Mariner-Mars 1964 Final Project Report

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MARINER-MARS 1964

FINAL PROJECT REPORT



Mariner-Mars 1964 spacecraft



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NASA SP-139

MARINER-MARS 1964FINAL PROJECT REPORT

Prepared under contract for NASA by

Jet Propulsion Laboratory

California Institute of Technology

ScientiSt and Technical In[ormation Division

OFFICE OF TECHNOLOGY UTILIZATION 1967

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

Washington, D.C.



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Editor's Note

This report presents the story of the Mariner-Mars 1964 project

from the time of its inception until October 1, 1965, when two-way

communications with the spacecraft were interrupted. By this date, the

objectives of the Mariner-Mars 1964 project had been completed. How-

ever, since that time, the spacecraft continued to function properly

and a follow-on project, called Mariner IV, was established to

continue operations with the spacecraft through 1966 and 1967. Plans

for tracking Mariner IV in its path around the Sun and for obtaining

additional telemetry data, as discussed in this report, were successfully

culminated during this follow-on project. Mariner IV continued to

function properly during 1966 and most of 1967, and the results of this

portion of its flight will be included in a subsequent report.

For Sale by the Superintendent of Documents,

U.S. Government Printing Office, Washington, D.C. 20402

Price $2.50 (paper cover)

Library of Congress Catalog Card Number 67-60049



Foreword

In November 1962, the National Aeronautics and Space Administration

established the Mariner-Mars 1964 project as the first phase of U.S. unmanned

exploration of Mars. Mariner IV, developed under that project, joined the grow-

ing list of U.S. space accomplishments when, on July 15, 1965, it flew within

approximately 9800 km (6100 statute miles) of the planet. A television instrument

photographed Mars, obtaining significant data about its surface; particles and

electromagnetic fields experiments yielded a vast storehouse of scientific infor-mation about the near-Earth and near-Mars environments and interplanetary

space; and, by measuring changes in the characteristics of radio signals as Mariner

IV passed behind Mar_ (as viewed from the Earth), information about the atmos-

phere of the planet was obtained.

Mariner IV was launched on November 28, 1964, on a trajectory which

would have taken it within approximately 242 960 km (151 000 statute miles) of

Mars. By a single, successful midcourse maneuver on December 5, 1964, its flight

path was altered to enable the close flyby. The total flight time was approximately

7 a_ months, but the story of Mariner IV actually began 2 years prior to its launch

and it extends beyond the July 15 close approach to Mars with chapters yet un-

told. This document presents that story from the time the idea for the mission was

conceived until October 1, 1965, when two-way communications with the space-

craft were terminated. Plans for tracking Mariner IV as it continues in its path

around the Sun and for attempting to obtain telemetry data during a future close

approach of Mariner IV to the Earth are also discussed.

Actually, the Mariner-Mars 1964 story involves more than the recorded facts

and figures included herein. The portion of the story which cannot be told on

paper is the human aspect: a story of a team whose members combined their skills

and talents toward one common goal. The Mariner-Mars 1964 project was an ex-

tremely complex undertaking, and, as such, required an enormous effort by many

people. The success of Mariner IV is a tribute to their efforts.

It is also a tribute to this country's economic strength and resources that a

mission such as Mariner IV could evolve from an idea to a reality in a little over

2 years. Such rapid progress is made possible by a fast-moving intricate mechanism

of many functioning parts. For Mariner-Mars 1964, the functioning parts included



MARINER-MARS1964

the engineers, scientists, administrators, and many other individuals who played a

direct role in the project and also, of equal importance, the support of the entire

country sharing our commitment to meet the challenges of space exploration.

A combination of the new concepts, methods, and techniques developed

under the Mariner-Mars 1964 project and those already proved by Mariner IV's

predecessors in space (such as the Ranger flights to the Moon and the Mariner II

flight past Venus) was indeed a winning combination for our first attempt at Mars

exploration. And, as was true with Mariner IV, the invaluable information

gathered both in the development stages of the project and during the Mariner IV

mission will be used in future space projects involving unmanned spacecraft which

will orbit the planets, soft-land on them, and explore their surfaces.

W. H. PICKERING,

Director, Jet Propulsion Laboratory,

California Institute of Technology.

vi



Contents

INTRODUCTION .................................................... 1

CHAPTER 1 Summary of Scientific and Engineering Results .......... 5

CHAPTER 2 Trajectory ........................................ 9

Earth-to-Mars Trajectory Characteristics ................ 9

Near-Earth Ascent ........................... 9

Heliocentric Transfer ......................... 12

CHAPTER 3

Mars Encounter .............................. 12

Heliocentric Orbit ............................ 1 3

Mariner-Mars 1964 Trajectory Design ................ 1 3Determination of Launch Interval and Arrival Dates.. 1 3

Aiming Point Determination .................... 1"/

Impact Probability Analysis .................... 20

Launch Constraints Analysis .................. 21

Final Prelaunch Trajectory Determination .......... 22

Orbit-Determination Operations ................ 23

Mariner IV Trajectory Results ....................... 27

Orbit-Determination Procedures ................. 27

Resulting Trajectory Values .................... 28

Orbital Elements of Mariner IV Trajectory ........ 35

Space Vehicle System Design and Testing Operations •. • 41

Launch Vehicle Design ........................... 41

Atlas D First Stage ............................ 41

Agena D Second Stage ........................ 43

Shroud System and Spacecraft Adapter ........... 45

Spacecraft System Design ........................... 46

Design Considerations ......................... 46

Preliminary Design ............................ 4"/

Spacecraft Mechanical Configuration ............ 50

Subsystem Functional Description ................ 58

Testing Operations ................................ 117

Environmental Testing Program .................. 11 7

Miscellaneous Qualification and Developmental

Testing .................................. 120

vii



MARINER-MARS 1964

CHAPTER 4

CHAPTER 5

CHAPTER 6

System Testing Program ........................ 122

Test Facilities ................................ 122

OperationaJ Support Equipment ................. 124

Flight History and Space Vehicle Performance .......... 131

Chronology of Mariner-Mars 1964 Flights ............. 131

Mariner III Flight .............................. 131

Mariner IV Flight ............................ 132

Mission Planning and Decision Philosophies ........... 155

Launch ..................................... 155

Midcourse Maneuver .......................... 156

Postmidcourse-Maneuver (Interplanetary) Cruise .... 1 56

Early Science Cover Deployment Exercise .......... 161

Encounter .................................... 162

Postencounter ................................ 164

Future Operations for Mariner IV ................ 166

Mariner IV Subsystem Performance ................... 16-/

Structure and Mechanisms ....................... 167

Radio ........................................ 168

Data Encoder ................................. 172

Video Storage ................................ 172

Command .................................... 173

Attitude Control .............................. 173

Central Computer and Sequencer ................. 175

Power ....................................... 1 75

Pyrotechnics ................................. 177

Propulsion ................................... 17"/

Temperature Control ........................... 178

Science ...................................... 1 79

Tracking and Data Acquisition ...................... 185

Scheduled Support J:or Mariner-Mars 1964 Missions ..... 185

Air Force Eastern Test Range .................... : 185

Goddard Space Flight Center .................... 18"/

Deep Space Network .......................... 188

Mariner IV Tracking and Data Acquisition Summary ..... 207

Space Flight Operations ............................ 213

System Functions and Responsibilities ................. 213

viii



* .b

CONTENTS

CHAPTER 7

APPENDIX A

APPENDIX B

BIBLIOGRAPHY

Data Processing Subsystem ......................... 216

Data Flow ........................................ 219

Real-Time Data ............................... 219

Non-Real-Time Data ........................... 219

General Paths ................................. 220

Operational Modes and Computer Programs ....... 220

Telemetry Data Forms After Processing ................ 229

Mariner-Mars 1964 Master Data Library ............... 230

Mariner IV Data Processing and Recovery Summary ..... 231

Data Recovery Summary ........................ 231

Data Processing Summary ....................... 232

Scientific Results and Conclusions ..................... 235Fields and Particles Experiments ...................... 235

Introduction .................................. 235

Helium Magnetometer ....................... 238

Cosmic Dust Detector ........................... 244

Ionization Chamber ............................ 24"/

Cosmic Ray Telescope .......................... 250

Trapped Radiation Detector ..................... 251

Solar Plasma Probe ............................ 254

Television and Occultation Experiments .............. 257

Introduction .................................. 257Television ................................... 261

Occultation .................................. 316

Project History and Organization .................... 323

Mariner-Mars 1964 Project Formalization ............ 323

Jet Propulsion Laboratory Organization and Management. 326

Lewis Research Center Responsibility ................. 327

Advisory Panel Participation ........................ 334

Abbreviations .................................... 335

................................................ 339

ix



Introduction

The planet Mars was named by the Romans for their ancient sanguinary god

of war because of its vermilion color. With the traditional belief that it may sup-

port life, Mars has overshadowed its far more brilliant Olympian sister, Venus, in

public interest. Since it is the planet in our solar system thought to be most like

Earth, scientists, philosophers, and writers have for many years speculated on what

this life, if it exists, might be like. As our knowledge of Mars has improved as more

sophisticated observational techniques have become available over the years, thefabled surface environment of Mars has appeared less and less promising in regard

to its ability to support terrestrial life forms. However, because of the amazing

........._,4_pt_h;1;tyof life on this anet, .... v ......... y ..... c on _,,m_ ,_nuu_ be ueu-J n

nitely excluded.

In order to answer the question of life on Mars, it will be necessary to land

instruments on its surface. Before this can be accomplished, the most desirable

location to attempt such a soft landing will have to be selected. However, before

either of these feats can become a reality, information concerning the atmospheric

and surface conditions of Mars and the environmental characteristics of the space

separating Mars from the Earth must be available to designers of the spacecraft.

Therefore, the first step in Mars exploration was a closeup (flyby) mission from

which the necessary planetary and interplanetary information could be derived.

Such a mission was that of Mariner IV.

The Mariner IV spacecraft was launched on November 28, 1964, and en-

countered Mars on July 15, 1965. The mission proved to be of immense scientific

and engineering importance. Scientific information is now available on regions of

the solar system never before penetrated with instruments. Observations from the

vicinity of Mars suggest entirely new concepts about the nature of the planet.

Spacecraft performance has proved our ability to design and construct a remotely

operated device of extreme complexity, and its continued operation established an

extremely high standard of reliability. Maintaining two-way communications over

distances up to 304 million km (190 million miles) demonstrates remarkable

advances in communications technology not thought possible a decade ago.

Design concepts used in the design of Mariner IV date back to 1959 when the

Vega project was begun at the Jet Propulsion Laboratory (JPL). From that project



MARINER-MARS1964

evolved the Ranger project, the first phase of U.S. lunar exploration. The Mariner

concept itself was formulated when a mission to Venus was planned for the 1962

flight opportunity. This mission was to be accomplished by using a 567-kg-class

(1250-1b) spacecraft launched by a vehicle consisting of a modified Atlas D first

stage and a Centaur liquid-hydrogen/liquid-oxygen, high-energy second stage.

However, when it became evident that the development of the Centaur stage had

not progressed sufficiently to make it available for the 1962 Venus launch period,

this mission series was canceled and another was formulated. The latter Venus

mission series, based on the use of an Atlas D/Agena B launch vehicle and a hybrid

spacecraft combining features of the Ranger and Mariner designs, formally

became the Mariner-Venus 1962 project. The Mariner I I spacecraft, developed

under that project, made history on December 14, 1962 (after 109 days of flight),

when on a predetermined trajectory it encountered Venus at a distance of 34 826

km (21 645 statute miles) from the planet. Valuable scientific data on Venus and

on interplanetary space were obtained.

The Mariner IV spacecraft was developed under the Mariner-Mars 1964

project, which was established as part of the National Aeronautics and Space

Administration (NASA) Planetary-Interplanetary Space Exploration program in

November 1962. Primary objectives were to make flyby scientific observations of

the planet Mars during the 1964-1965 flight opportunity and to transmit the

results back to Earth. Secondary objectives were to develop and study equipment

and techniques required for such a mission and to perform certain scientific meas-

urements during the trip.

NASA, through its Office of Space Science and Applications (OSSA),

assigned the Jet Propulsion Laboratory, California Institute of Technology, at

Pasadena, California, the management responsibility for the project under con-

tract NAST-100; the spacecraft system; and tracking, data acquisition, and space

flight operations activities. Responsibility for the overall direction and perform-

ance evaluation of the project was assigned to the OSSA Lunar and Planetary

Programs Office. Management responsibility for the launch vehicle, an Atlas D/

Agena D combination, was assigned to the NASA Lewis Research Center (LeRC)

of Cleveland, Ohio. NASA Goddard Space Flight Center at Greenbelt, Maryland,was assigned launch operations responsibility for the project.

Discussed in this document are the Mariner-Mars 1964 trajectory; space-

vehicle system design and testing operations; flight history and space-vehicle

performance; tracking, data acquisition, and space flight operations activities;

Mariner IV scientific (planetary and interplanetary) results and conclusions; and

2



INTRODUCTION

project management and organization. Thus, a comprehensive report is given

herein of activities from the inception of the project in November 1962 until the

end of the Mariner IV mission on October 1, 1965. Future plans involving the

Mariner IV spacecraft are also discussed.

A brief history of the Mariner-Mars 1964 project is given in appendix A, and

abbreviations used in this book are defined in appendix B.

3



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CHAPTER 1

Summary of Scientific and Engineering Results

The Mariner IV spacecraft, the first man-made probe to travel to Mars, was

far more than a technological experiment: it was an extremely complex instrument

designed to perform multiple scientific experiments to extend man's knowledge of

his own planet, interplanetary space, and the planet Mars. In addition, the devel-

opment of this spacecraft contributed a vast amount of engineering knowledge to

space technology which is essential to the design of future planetary and inter-

planetary space probes. Regarding the engineering evolution and technical

development of the spacecraft, the following list highlights only a small fraction of

the many technological demands placed on, and operational "firsts" required of,

Mariner IV:

1. The Mariner IV mission was the first to require 9 months of successful

spacecraft operation to achieve mission success.

2. The spacecraft was required to be fully automatic; i.e., it had to be

capable of completing its entire mission without ground-based intervention,

except for trajectory-correction maneuvers and, of course, tracking and data

acquisition.

3. The design of the spacecraft required at least two independent means of

initiating every specific function or event critical to the success of the mission.

4. The complexity of its assigned tasks required that the spacecraft contain

138 000 parts, as compared with 54 000 parts in its predecessor, Mariner II, with

only a 61-kg (135-1b) increase in spacecraft weight.

5. The spacecraft was required to communicate with Earth over extreme

distances: at least 2 _ times greater than those of previous missions. The Mariner

IV mission involved the first use of the S-band communications system.6. Since it was traveling away from the Sun during its journey to Mars,

Mariner IV had to withstand a widely varying range of thermal conditions and

required twice the solar panel area of Mariner II.

7. The solar pressure vanes at the ends of Mariner IV's solar panels were

unique in utilizing solar pressure effects (about a millionth of a pound per vane)

to assist in maintaining stable orientation of the spacecraft toward the Sun.

$



MARINER-MARS1964

B

8. Whereas previous spacecraft used the Earth for roll-axis stabilization,

Mariner IV was the first successful spacecraft to use a star (Canopus) for this

purpose. (The Earth could not be used since, during much of the flight, the Earth

appeared as a relatively dim crescent as it moved across the face of the Sun.)

9. In case more than one trajectory correction was necessary, the Mariner IV

trajectory-correction propulsion system was designed to be restartable. This was

the first time such a dual capability was available.

10. The spacecraft was required to store the data it acquired at Mars on

magnetic tape for transmission to Earth at a later time.

11. The Mariner IV mission was the first in which coherent radio transmission

was used to probe conditions on another planet (i.e., the first in which an occulta-

tion experiment was conducted).

In order to meet the unique and more stringent requirements of the mission to

Mars, every effort to insure proper operation of the spacecraft had to be made.

Rigorous testing programs were carried out, as were very demanding parts-screen-

ing and quality-assurance activities. Backup designs were accomplished in several

critical areas. Every effort was thoroughly documented so that all areas connected

with the mission would be as well informed and coordinated as possible. As it

turned out, these activities were well worth the effort involved, for the Mariner IV

spacecraft was truly a milestone in engineering technological accomplishment.

In addition, the scientific results of the mission greatly enhanced the level of

our knowledge concerning planetary and interplanetary conditions. During theMariner IV journey to Mars, approximately 23 million scientific measurements

were made. During the early part of the trip as it passed through the region of

space influenced by, the Earth, Mariner IV measured with great precision the Van

Allen radiation belts, the terrestrial magnetic field which holds them, and the

interface between the solar plasma (ionized gas) in space and the Earth's magnetic

field. The spacecraft measured the rise and fall of solar activity throughout its

journey, and, although the mission took place during a period of decreased solar

activity, over 20 solar flare events were detected. Solar wind velocities varied

widely during the flight, and magnetic fields fluctuated concurrently. About 235

micrometeorite (cosmic dust) impacts on the cosmic dust detector were recorded.

As it flew by Mars, Mariner IV proved conclusively that Mars has a very

small magnetic dipole moment compared with that of the Earth (less than 0.1

percent the value for the Earth), if it has one at all. This measurement was sup-

ported by the fact that no radiation belts were detected in the vicinity of the

planet. Without a magnetic field to deflect energetic particles, Mars then is



SUMMARY OF SCIENTIFICAND ENGINEERINGRESULTS

directly cxposed to bombardment by cosmic rays and solar plasma. The absence of

a magnetic field also implies that some feature of the Earth's internal structure is

missing in Mars: presumably a liquid core. It can thus be concluded that, without

such a core, Mars probably lacks much of the internal activity that results in

changes in the topography of the Earth (such as mountain building).

Measurements by the instrument which detected the presence of cosmic dust

throughout the flight indicated no concentration of solid matter in the vicinity of

Mars. In fact, the measurements seem to indicate that Mars has swept a dust-free

path in its orbit around the Sun and has thus reduced the quantity of matter in

that region.

The atmosphere on Mars was found to be extremely thin compared with that

on the Earth. Mars' daytime ionosphere appears to be approximately equivalent

to that of the Earth at night. Since Mars has a surface pressure measuring between

0.5 and 1 percent that of the Earth, thus providing little aerodynamic braking

assistance to facilitate a soft landing, it will be much more difficult than was ex-

p,.c_eu to design capsules capable of landing on the Martian surthce. However,

the discovery that density decreases quite rapidly in Mars' upper atmosphere

indicates that it may be possible to orbit at lower altitudes than were previously

thought feasible.

The most surprising discovery of Mariner IV was that the surface of Mars

closely resembles that of the Moon. The existence of craters seems to indicate

that the surface may be 2 to 5 billion years old and very well preserved, since none

of the erosive effects encountered on Earth would be encountered on Mars.

The close flyby of the Mariner IV spacecraft past Mars and the accurate

tracking of the spacecraft on its trajectory allowed improvements in the calculation

of the planet's mass. A new value, with significantly improved accuracy, for the

ratio of Mars' mass to that of the Sun was obtained.

If Mariner IV is still operating during its close approach to the Earth in

September 1967, even more data from the spacecraft will be received. These data

will be valuable since: (1) the measurements will come from a region of space,

some 16 million km (10 million statute miles) above the orbital plane of the Earth,

at a time of increased solar activity; and (2) the measurements will be madesimultaneously with those of the Pioneer and other spacecraft from different

regions of space about the Sun. All evidence obtained to date indicates that the

spacecraft is continuing to operate quite well and that it will be possible to obtain

these data in 1967.



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CHAPTER 2

Trajectory

EARTH-TO-MARS TRAJECTORY CHARACTERISTICS

For the best utilization of the available rocket energy, the relative motion and

positions of the planets about the Sun must be considered, since the spacecraft

itself, once freed from the Earth's gravitational pull, becomes a member (planet-

oid) of the solar system and, therefore, subject to the same inertial forces. As a

result of the changing planetary relationships, the available time of departure

(launch date), the speed of travel, the time of flight, and the flight path change

continually.

Of prime significance in scheduling an interplanetary trip is the knowledge

that a free-falling (orbiting) body travels in an imaginary plane which passes

through the center of a controlling body. For an Earth-to-Mars interplanetary

spacecraft, this controlling body is first the Earth, then the Sun, then Mars, and

again the Sun. Within each of these planes, the spacecraft follows certain geo-

metric paths that are mathematically definable and predictable. The trajectory

path describes various conic figures: Earth orbit: ellipse; Earth escape: hyperbola;Sun-centered transfer orbit: ellipse; Mars encounter and escape: hyperbola; and

Sun-centered permanent orbit: ellipse.

Near-Earth Ascent

The ascent phase for the mission (fig. 2-1) could be divided into three

portions: The powered-flight ascent, the parking-orbit coast, and the postinjection

ascent. The powered-flight ascent consists of the Atlas D and Agena D thrust

periods. At the end of the first Agena D thrust period, the Agena D/spacecraft

combination is placed in a nearly circular parking orbit at an altitude of approxi-

mately 188 km (117.5 statute miles). The Agena D/spacecraft combination

"coasts" in this orbit until the optimum point is reached for a final thrust phase

(near perigee or closest point of the required escape hyperbola), at which time the

Agena D engine is restarted. Injection takes place upon termination of this final

Agena D thrust period and, consequently, when the spacecraft is "injected" into

its hyperbolic orbit away from the Earth.



MARINER-MARS 1964

/--ATLAS DATLAS D I BOOSTER

SUSTAINER /GENA D

FIRST BURN q

ORBIT

COAST II POSITION / NI I

AGENA O

SECOND BURN %

POST,NJECTIO /--ESCAPEHYPERBOLAASCENT __i "'-._

"RI"_'_"-A SY M PT OT E OF

ESCAPE HYPERBOLA

OUTGOING RADIAL,

GENERAL DIRECTIONOF EARTH'S MOTION

FIGURE 2-1.--Typical near-Earth ascent trajectory profile (in near-Earth ascent

trajectory plane) for Mariner-Mars 1964 spacecraft.

The postinjection ascent describes an escape hyperbola with the Earth's

center as its principal focus. A characteristic of the escape trajectory is that after

a few hours the spacecraft travels essentially radially away from the Earth along

the outgoing asymptote of the escape hyperbola. This asymptote is a straight line

parallel to the outgoing radial (a straight line connecting Earth's center and the

geocentric point in space at which the spacecraft finally escapes the pull of the

Earth's gravitational field). The spacing between these two lines is determined by

the eccentricity of the hyperbola, which is, in turn, determined by the velocity of

the spacecraft at the time of injection. The direction of the outgoing radial is

defined by its celestially referenced right ascension and declination. The value of

the right ascension and declination is determined from the relative positions of

Earth at launch and Mars at encounter, and remains essentially fixed for a given

launch date.

Since the launch site remains at a fixed geographic latitude, the requirement

of the near-Earth ascent phase is to match the powered-flight portion (which

10



TRAJECTOR Y

begins at the launch site) to the required escape velocity vector. The direction of

this vector is determined by the asymptote of the escape hyperbola. The hyper-

bolic excess velocity is the geocentric velocity (i.e., the velocity in relation to the

Earth's center) which the spacecraft attains a few days after launch as it becomes

free of the gravitational pull of the Earth. (The magnitude of this velocity is pro-

portional to the square root of the injection energy.) The outgoing radial and the

geocentric position of the launch site define the plane of the near-Earth ascent

trajectory.

The injection energy (i.e., that energy required to effect the ballistic transfer

from Earth to Mars) is at a minimum every 25 months. This time period is deter-

mined by the harmonic relationship between the duration of time required by

Mars and the Earth to complete their orbital revolutions about the Sun (approxi-

mately 687 and 365 days, respectively). Any launch other than one at the optimum

time requires an increase in injection energy and a resultant decrease in allowable

spacecraft weight. (For a given spacecraft weight, there is a corresponding value

of injection energy which is achievable by the launch vehicle.) For any available

energy above the absolute minimum, there is a corresponding launch interval

(number of days) in which the spacecraft can be launched.

Each day in this launch interval has its own launch period or "window" of

only several hours or minutes. This launch window results from several inter-

related restrictions and conditions:

1. The geographically fixed launch site on the surface of the Earth.2. The Earth's center point.

3. The geocentrically referenced location and direction of injection into an

Earth-escape hyperbolic orbit. (The range of locations and directions varies with

celestial latitude of Mars at encounter and with the Earth's own orbit position

and, therefore, does not change significantly for any given day.)

4. The rotation of the Earth about its axis.

5. The 93 ° to 111 ° east of true north (i.e., north referenced to the Earth's axis

rather than to the magnetic poles) geographic launch range or corridor. (This is

an Air Force Eastern Test Range (AFETR) safety restriction to minimize the

hazard to populated areas below the launch vehicle ascent path.)

The parking-orbit coast time decreases as the time of launch is delayed during

a daily launch window, since the angle between the launch-site position vector and

the outgoing radial (projected backward) gets smaller as the Earth rotates. The

change in both launch azimuth and parking-orbit coast time results in a wide

geographic range of injection locations. The effect of the Earth's rotation is to

11



MARINER-MARS1964

move the launch site eastward 15 deg/hr and to change continuously the required

launch azimuth to coincide with the continuously inclining spacecraft orbit plane.

(The launch site, Earth center point, and asymptote of the escape hyperbola willall be in the orbit plane of the spacecraft when the location and direction condi-

tions of injection are satisfied.) Since the launch azimuth increases as the time of

launch increases, the period of time during a day is limited during which the

required range of azimuth headings is available. This period of time defines the

launch window.

Heliocentric Transfer

The heliocentric (Sun-centered) transfer orbit is an ellipse that essentially

intersects the Earth at launch and the planet Mars at encounter, with the Sun at

one focus. The hyperbolic excess velocity vector and the Earth's velocity vector

about the Sun add vectorially to determine the velocity at which the spacecraft

enters the heliocentric orbit. Since the spacecraft is launched "forward" from the

Earth's orbital velocity, the magnitude of the spacecraft's resultant velocity vector

is larger (relative to the Sun) than that of the Earth's velocity vector. As the space-

craft then travels outward from the Sun, it decreases in speed and the Earth finally

passes it (as viewed from the Sun).

The minimum velocity required to escape the pull of the Earth's gravitational

field is approximately 11.18 km/sec (6.95 statute miles/sec); the actual velocity

required to reach Mars is approximately 11.44 km/sec (7.11 statute miles/sec).

The additional velocity is necessary to move the spacecraft out farther away fromthe Sun and to displace it from the ecliptical plane (the orbital plane of the Earth

as projected on the celestial sphere 1) on a trajectory suitable for intersection of the

orbital plane of Mars (as projected on the celestial sphere) at the optimum time

for planetary encounter.

Mars Encounter

During the encounter with Mars, the primary source of gravitational attrac-

tion is the planet itself. The trajectory of the spacecraft is similar to that during the

near-Earth ascent phase (both described by hyperbolas), except that during Mars

encounter the spacecraft travels along an incoming hyperbolic path. Also, the

altitude of closest approach to Mars is several times greater for the Mars encounter

phase than for the near-Earth ascent phase.

1An imaginary sphere of infinite radius, with the observer at its center, on which all celestial bodies except

the Earth appear to be projected.

12



TRAJECTOR

Heliocentric Orbit

After Mars encounter and a hyperbolic escape from the gravitational pull ofthat planet, the spacecraft begins a new heliocentric orbit. The parameters of the

new elliptic orbit differ greatly from those of the preencounter orbit because of

large inertial perturbations introduced during the encounter.

MARINER-MARS 1964 TRAJECTORY DESIGN

Determination of Launch Interval and Arrival Dates

In order to determine the acceptable launch intervals and trajectory charac-

teristics for the Mariner-Mars 1964 mission, studies of the relationship between

flight time, launch date, and injection energy were required. Six major factors hadto be considered in the preliminary trajectory design:

1. Desire for scientific data gathering during a close flyby and transmission of

the data back to Earth.

2. Use of the Atlas D/Agena D launch vehicle and an approximately 189-km

(I17.5-statute-mile) nearly circular parking orbit.

3. Launchings of two spacecraft with at least a 2-day separation in arrivaldates.

4. Launchings from two separate launch pads with a maximum launch

azimuth spread of from 90 ° to 1 |4 ° east of true north. (In order to maximize the

probability of launching two spacecraft, it was necessary to maximize the size of

the launch window and thus the launch corridor. Therefore, permission was

granted for a launch corridor of 90 ° to 114 °, instead of the usual 93°to Ill°.)

5. Use of an attitude-stabilized spacecraft with the Sun and the star Canopus

as reference bodies.

6. A maximum allowable Earth-spacecraft communications distance of ap-

proximately 250 million km (156 million statute miles).

Two possible types of trajectories may be used: The type I trajectories are

characterized by a heliocentric transfer angle of less than 180 ° from launch to

encounter; the total transfer angle for type II trajectories lies between 180 ° and

360 °. Longer flight times and communications distances at encounter are allowed

by the type II trajectories.

In order to increase the probability of launching two spacecraft, a launch

period composed of both type I and type II trajectories was selected to provide a

maximum number of launch days. The two types of possible trajectories are

represented in figures 2-2 and 2-3 as two sets of closed contours. As may be seen,

13



MARINER-MARS 1964

the energy contours for both types provided overlapping flight times and com-

munications distances at encounter.

After careful study of all possible trajectories, a decision was made to use

those with small variations in arrival date. Since the relative positions of Earth,

Mars, and the Sun are fixed for a given arrival date, the selection of constant-

arrival-date trajectories allowed the use of a fixed high-gain antenna on the space-

craft. Also, because of the relatively constant spacecraft approach direction and

velocity, the use of a fixed aiming point was permitted. A single arrival date was

not allowable, however, because of the requirement for a separation of at least 2

days between the Mars encounter of each Mariner-Mars 1964 spacecraft. The 1965

480 I i v i

NOTE : INJECTION ENERGY C 3 IS TWICE THE TOTAL

ENERGY/UNIT MASS WITH UNITS km2/sec 2

440 ---- CLASS I (LOWER ENERGY CONTOURS)

-- -- -- CLASS 13" (HIGHER ENERGY CONTOURS) _ _ ....

360 -- / i / /

/ / / / /

// // // / / /

32o ; z // / I/ / L/"-_:_----- _ 18/ / / TYPE_ /L_;_'__ 15

28O 9.843 --

!/ ! ,,,,

160 -- 24 / _ _,_.///

120

15 25 4 14 24 4 14 24 3

OCT NOV DEC JAN

LAUNCH DATE

FIGURE 2-2.--Time of flight as a function of launch date.

14



TRAJECTOR Y

350

33O

T 1 1 T 1

NOTE: INJECT/ON ENERGY C 3 IS TWICE THE TOTAL /ENERGY/UNIT MASS WITH UNITS km2/sec 2

-- CLASS I (LOWER ENERGY CONTOURS)

-- -- -- CLASS 11"(HIGHER ENERGY CONTOURS)

%X

E

u)_

0

l.-

z

0c.p

+

310

290

27O

250

230

210

TYPE I

:'3= 21

18

15

12

t0

98.843

190

I 7O

150

I0

12

15

18

21

24

130

15 25 4 14 24 4 14 24

OCT NOV DEC JAN

LAUNCH DATE

FIGURE 2-3.--Earth-Mars communications distance as a function of launch date.

13

15



MARINER-MARS 1964

arrival dates selected --July 15 and 17 for type I trajectories, July 17 and 19 for type

II trajectories resulted in trajectories with near-minimum energy for each launch

date (fig. 2-4) and insured that a maximum possible launch interval would be

available once the spacecraft were built and the maximum injection energies

attainable were calculated. The communications distances at encounter for the

selected dates varied from 217.5 to 221.7 million km (135.2 to 137.8 million statute

miles). It was felt at that time that the maximum communications distance would

be 251 million km (156 million statute miles), reached approximately 25 days after

encounter; thus, sufficient time would remain after encounter for two complete

transmissions of the television picture data. (During the actual Mariner IV flight,

however, telemetry contact was maintained to a distance of approximately 309.2

million km (192.2 million statute miles) on the 78th day after encounter.)

Because of the direct relationship between the injection energy achievable

from the launch vehicle and the weight of the spacecraft, a trade-off between the

140

130

E

IIO

F-

_0 o

I-

@

/

,,,, OATE f"ULY 17, 1965

JULY 19,196,.5 --JULY 15,1965

JULY 17, 1965

[ , ,4 6 8 I0 12 14 16 18 20 22 24 26 28 30 2 4

NOV DEC

LAUNCH DATE

FIGURE 2-4.--Geocentric injection energy as a function of launch date.

/

/

6 8 I0

16



TRAJECTORY

length of the firing period and the spacecraft weight existed. Once the spacecraft

weight of approximately 261 kg (575 lb) was established, a launch interval ofapproximately 27 days was calculated.

Aiming Point Determination

Studies were also conducted to determine the region near Mars most desirable

for flyby. This aiming point could be specified by a vector B directed from the

center of Mars to the point at which the incoming asymptote pierced the T-R

plane. This plane, shown in figure 2-5, is defined to be normal to the incoming

asymptote of the spacecraft approach hyperbola, T is a unit vector parallel to the

ecliptic plane, and R is normal to T. The aiming point is usually defined by its

two components B oT and B'R or by its magnitude B (from the center of Mars)and the polar angle 0 measured from T to R. Spacecraft design and planetary-

science-experiment considerations imposed numerous constraints on the selection

of the aiming point; among the more important constraints were:

1. The Sun-probe (spacecraft)-near limb of Mars angle (SPM) had to be

greater than 0 ° during the entire encounter phase, since the Sun served as both a

power source and a reference body for the attitude-control subsystem. Thus, the

aiming point must be outside the SPM --0 ° contour in figure 2-6 (calculated for a

November 15, 1964, launch).

2. No part of Mars or its moons could fall within a region 4-26 ° in clock angle

and 920+36 ° in cone angle during the encounter phase so that Canopus would be

the only detectable body within the Canopus-sensor field of view; i.e., the Cano-

pus-probe-near limb of Mars angle (CPM) had to be greater than 36 ° during the

entire encounter phase. Therefore, the aiming point must be outside the region

defined by CPM = 36 ° in figure 2-6.

3. It was desirable for the spacecraft to pass within 40 200 km (25 000 statute

miles) of the surface of the planet to insure a maximum scientific return from the

fields and particles experiments on the spacecraft and the desired picture resolu-

tion from the television experiment.

4. The aiming point had to lie between 0 ° and 90 ° from the T-axis in the

T-R plane (fig. 2-5). This constraint was necessary to insure that the planet would

fall within the field of view of the planetary-scan system and to provide a picture

trace across a desirable region of the planet.

5. The Earth-probe-near limb of Mars angle (EPM) had to be less than 0 °

sometime during the encounter phase to enable the spacecraft to pass behind Mars

(as seen from the Earth). Then the occultation experiment could be performed to

17



MARINER-MARS 1964

determine the characteristics of the Martian atmosphere. Thus, the aiming point

had to lie within the region defined by EPM--0 ° in figure 2-6.

6. The probability that the spacecraft would impact Mars had to be less than

10-4. Therefore, the region defined by impact probability (IMP) --10 -4 in figure

2-6 had to be avoided.

After careful consideration of the constraints, aiming points were selected to lie on

the centerline of the Earth's occultation contour with B -- 12 068 km (7500 statute

miles). Since the centerline varied slightly with trajectory type and arrival date,

the aiming point would change slightly also.

The precise arrival time had to be selected also and was chosen so that the

spacecraft could be viewed by the Goldstone Deep Space Communication Com-

plex during the encounter sequence. Since the television sequence was designed to

occur before closest approach, the arrival time (defined as the time of closest ap-

proach to Mars) was chosen to be 1 hour past the middle of the Goldstone-view

period.

TARGET 5'0 (OUTGOING ASYMPTOTE)

PLANE OF THE

APPROACH

5"I

CLOSEST APPROACH

REFERENCE PLANE

T

il

_'I (INCOMING ASYMPTOTE)

TARGET-CENTERED

HYPERBOLA

R

FIGURE 2-5.--Definition of B • T, B • R system.

18



TRA JECTOR Y

ii

19



MARINER-MARS1964

Impact Probability Analysis

To protect Mars from possible contamination by viable organisms from

Earth, a NASA policy stated that any unsterilized spacecraft launched to Mars

must have less than a 10 -_ probability of accidental impact with the planet. This

requirement played a key role in the determination of the aiming point for the

Mars flyby. A retrorocket was provided on the Agena D for the first Mariner-Mars

1964 spacecraft (Mariner III) for firing after injection to insure that the require-

ment was met for the Agena D. However, since the second launch (that of

Mariner IV) was delayed (for design of a new shroud), it became desirable to

remove this retrorocket since the resulting weight reduction would allow an

increase in available injection energy and thus provide a few extra days for the

launch period. Therefore, it became necessary to bias the nominal aiming point at

Mars in order to meet the quarantine requirement.

Since the overall probability that the spacecraft or the Agena D or both

would impact the planet was dependent on 11 probabilities that various other

events would occur, each of these probabilities was computed and the values were

added. It was found that the new aiming point resulting from the combined effect

of both the injection velocity and yaw biases would allow the spacecraft to pass

sufficiently away from the planet that the impact probability at injection would be

approximately 0.9 × 10-q

In addition to the possibility of planetary contamination by accidental

impact, a possibility existed for contamination by viable particles expelled fromthe spacecraft as it passed the planet. These particles could come from the gases

expelled by the attitude-control subsystem jets, the gases expelled by the mid-

course motor, and/or the outgassing from the spacecraft.

The first possibility was analyzed as follows: The attitude-control gases

would follow a considerably different flight path than that of the spacecraft

because of their different relative velocities; it is reasonable to assume that the

interplanetary environment would destroy any viable organisms except those

emitted very near the planet. With the solar-pressure vanes functioning, the

number of attitude-control subsystem jet actuations would be expected to be

almost zero except during maneuvers. Since the effect of solar pressure on the

emitted particles would be 10 _ to 10 _ times greater than that on the spacecraft,

particles emitted anywhere but very near the planet would be blown far away

from it. A further study was made to estimate the probability that particles ejected

near the planet would impact it. During this period, the control jets would most

probably not be operating because of the solar-pressure vanes, and some viable

2o



¢,,

TRA JECTOR Y

particles would already have been expelled during the flight. Since tile spacecraft

nitrogen-gas tanks were assembled and filled in uhraclean rooms and since all

attitude-control nitrogen from the tanks had to pass through two extremely effi-

cient and reliable filters in series, the probability of impact on Mars by a viable

organism from this source was computed to be less than 10 -12.

It was estimated that the probability that viable organisms would be expelled

from the midcourse motor was essentially negligible because of the extremely high

temperature environment the particles would have to survive. Even if they sur-

vived, they would then have to travel at least 200 days through space to encounter

Mars, since the midcourse maneuver would occur within the first few days of the

mission. During that time, the particles would be exposed to continuous ultra-

violet radiation with a high probability of destruction. In addition, the particle

trajectory would be radically different from the spacecraft trajectory.

Particles outgassed from the spacecraft would also have a negligible proba-

bility of contaminating the planet for two reasons: (1) Since the temperature of

the spacecraft would be highest during the early phase of the flight, it was expected

that what little outgassing did occur would take place at that time, subjecting the

particles to months of ultraviolet radiation; and (2) even if the particles survived

the environment, they would undoubtedly be perturbed considerably off an

impact course by solar pressure.

Thus, it could be concluded with considerable certainty that a negligible

probability existed for contamination of Mars by particle ejection from the space-craft.

Launch Constraints Analysis

An analysis was conducted to define all constraints affecting the launch of the

Mariner-Mars 1964 spacecraft caused by the spacecraft itself, the launch vehicle,

the space flight operations, and the tracking and data acquisition activities re-

quired for a successful mission. The only constraints imposed by the spacecraft

were those due to the Canopus sensor and low-gain-antenna operational charac-

teristics. The Atlas D first stage and Agena D second stage imposed several impor-tant constraints on the launch which had to be considered in the establishment of

the available launch interval. Since space flight operations (discussed in ch. 6) for

both launches could be continuous, no constraints were imposed by this pos-

sible source. Also, adequate deep space-station coverage (discussed in ch. 5) was

planned for both launches. A constraint which resulted from the inability to deliver

21



MARINER-MARS 1964

and analyze realistically certain tracking and telemetry data much before 36

hours dictated that at least one launch window should elapse between the

launches.Various computer programs were used in this analysis. The final launch

window designs were shown on a launch constraints board maintained in the JPL

Space Flight Operations Center at the Air Force Eastern Test Range. This board

was kept current by a continuing analysis of all possible constraints until both

Mariner-Mars 1964 spacecraft had been launched.

Final Prelaunch Trajectory Determination

With a launch interval of 27 days, the first spacecraft launch was scheduled

for November 4, 1964. It was found that, by making the arrival date for thisspacecraft July 17, 1965, and by accepting some penalty in injection energy, it was

possible to utilize an aiming point which would provide good occultation and good

television coverage. As shown in figure 2-7, launch days November 4 to 10 re-

quired the use of type II trajectories; since it had been decided that only one

Mariner-Mars 1964 spacecraft would be launched during the type II trajectory

JULY 19

O)

l-

JULY L7

>

JULY 15

ITYPE Tr-

4 6

NOV

I I

ITYPE I

i

I0 IZ 14 16 IB 20 22 24 26 28 30

LAUNCH DATE_1964

[ [

2 4 6 8

DEC

FmURE 2-7.--Arrival date as a function of launch date.

22



TRAJECTORY

period, the second spacecraft would be launched sometime after November 10 and

would thus have a type I trajectory.

After considering all possible constraints, final precision trajectories were

computed, and a detailed simulation of the Atlas D/Agena D ascent trajectory

was computed to determine proper guidance and control settings. Also, the launch

and parking-orbit coast times were computed to yield arrival conditions at Mars as

specified. These were determined for a 90 ° to 114 ° launch azimuth interval. The

daily firing period was about 3 hours, with injection locations confined to a region

of about 10.5 ° in latitude and 69.6 ° in longitude over an area slightly west of

South Africa and extending into the Indian Ocean. Parking-orbit coast times

ranged from 19.1 to 34.6 min, depending on both launch date and launch

azimuth.

Orbit-Determination Operations

Control of the aiming point parameters B.T and B-R and of the time of

flight was essential for control of the flyby distance, look angles, and illumination-

for-science instruments; correct timing of the automatic encounter sequence; and

adequate deep space-station view periods at encounter. The launch vehicle theo-

retically is supposed to inject the spacecraft onto a trajectory having the required

arrival parameters; however, because of various uncertainties in guidance param-

eters which are unavoidable before the mission, the initial parameter values are

not necessarily acceptable once the spacecraft has been launched. Therefore, radio

tracking data received after injection were used to determine the initial trajectory

for the Mariner-Mars 1964 spacecraft; to ascertain whether or not a midcourse

maneuver was necessary; and, if so, to compute a velocity increment that, when

added to the spacecraft velocity vector, would correct the trajectory parameters.

Orbit-determination operations centered around the orbit-determination com-

puter program using an IBM 7090 digital computer. A functional block diagram

of the midcourse-maneuver operations program is given in figure 2-8.

Because of tolerances in the guidance system components, some errors in the

maneuver were unavoidable, and, therefore, the miss at the target would not be

totally nullified by execution of the maneuver. The total allowable miss param-

eter dispersions at the target caused by uncertainty in orbit determination and

maneuver-execution (based on 15 m/see (50 ft/sec)) errors were specified to be

less than 2262 and 4093 km (1406 and 2544 statute miles), respectively, or a 30-

rain error in flight time.

23



MARINER-MARS 1964

AUTOMATIC FLOW

MANUAL FLOWD,,,.

(i)

INTRODUCTORY

PRINTOUT

(

MIDCOURSE

DECISION

(2)

J 0

RESIDUAL

MISS

MIDCOURSE

COMMAND

GENERATIONPROGRAM

FLYBY (5) I

TRAdECTORY

FINE PR NT

PROPULSION

uJ_

DISPERSION (4)]_ELLIPSE

0

(6)CAPABILITY

ELLIPSE

GENERATOR

1CONTOUR (7)_PLOTTING

FIOURE 2 8.--Functional block diagram of naidcourse-maneuver operations program.

1

IIIIIIII

)lIIIIIII

II

4IIIII

After the midcourse maneuver, an estimate of the new trajectory would be

made to determine whether a second maneuver was required and to allow control

of the remainder of the mission in a manner that would enable the maximum

24



TRAJECTORY

probable data return from the science instruments on the spacecraft. Throughout

the cruise and encounter portions of the mission, the orbital estimate would be

updated repeatedly as new tracking data became available.

One of the first requirements in precise orbit determination is a source of

accurate tracking data. These data can include almost any type of measurement

that in some way describes the position or velocity of the spacecraft at some point

along its trajectory. The most important measurements made by deep space sta-

tions are of angles and two-way Doppler effect. The angle data are valuable

primarily during the very early portion of a mission when the trajectory geometry

is changing rapidly. The two-way Doppler effect, however, is most valuable and

accurate after the spacecraft has left the vicinity of the Earth. (The two-way

Doppler effect is described in ch. 3 and is discussed in more detail in ch. 5.) The

two-way Doppler system developed by the Deep Space Instrumentation Facility

(DSIF) is probably the most accurate source of spacecraft tracking data in exist-

ence. Using extremely precise rubidium frequency standards (stable to about 1

part in 10 n over both long and short periods of time), it can provide low-noise,

unbiased tracking data from a spacecraft transmitting at 10-watt power from a

range of over 100 million km (62.5 million statute miles).

An estimate of the orbit of the spacecraft is computed from the tracking data

by a weighted least-squares fitting technique. By this technique, a theoretical

spacecraft trajectory is computed that best fits the observations concerning the true

trajectory. Since a free-space trajectory must obey a known deterministic set ofequations of motion, a set of trajectory initial conditions is all that is required. The

independent variables that specify a spacecraft trajectory and whose effects can be

seen in the tracking data consist not only of the initial spacecraft position and

velocity, but also of the masses of the various gravitating bodies, the lunar and

planetary ephemeris scaling factors, and the reflectivity of the spacecraft (since its

trajectory is perturbed by the force of the impinging sunlight). It is because the

effects of these parameters can be accurately measured by Doppler tracking that

the determination of the orbit of a spacecraft can frequently produce values of

physical constants that are far more accurate than those available by other

methods.Although the direct result of the least-squares fit to the tracking data is the

set of trajectory initial conditions, the parameters of greatest interest are the target

conditions resulting from integrating (i.e., mathematically projecting) the tra-

jectory forward to its point of closest approach to Mars. These parameters are

expressed in the B-plane system previously defined in figure 2-5.

25



MARINER-MARS1964

w

The least-squares orbit computation also produces an estimate of the orbit

accuracy. This accuracy is a function of the way in which the data are weighted

and of the a priori uncertainties attached to the physical constants and the stationlocations (with respect to the exact center of the Earth). In the Mariner-Mars 1964

computer program, 66 of these uncertainties were treated as error sources. The

accuracy estimate appears in the form of standard deviations on the estimated

parameters. (Standard deviations are also called 1or values.) These also can be

mathematically projected forward to the target to show uncertainties in the B-

plane, and are expressed in terms of the semimajor axis, semiminor axis, and

orientation angle of the l_r dispersion ellipse in the plane.

In classical least-squares fitting, each data point is given a rating indicating

the assumed variance on the point. Thus, measurements thought to be more

accurate exert a stronger influence on the estimate. All possible sources of error

are considered.

A study to provide a detailed analysis of the relative effects of various error

sources which degrade orbit accuracy showed that: (1) orbit uncertainties would

be only moderately affected by changes in data weighting; (2) uncertainties in the

Earth's gravitational constant and in the astronomical unit would contribute a

negligible amount to orbit uncertainty; and (3) for an early launch date, the major

source of error would be uncertainties in station locations (with respect to the

exact center of the Earth), and, for a later launch date, the major source of error

would be uncertainties in the solar pressure forces.

Another study demonstrated a steady increase in orbit-determination ac-

curacy as later launch dates were used, with an apparent discontinuity near

November 11, 1964, when the transition from type II to type I trajectories would

occur (which caused this date to be eliminated from consideration as a possible

launch date). It was found that, following a midcourse maneuver performed from

3 to 10 days after launch, 3 to 5 days of tracking would reduce the semimajor axis

of the dispersion ellipse to less than 2010 km (1250 statute miles). After 60 to 90

days of tracking, solar pressure forces would be known well enough to be a negli-

gible source of error. After that time, the semimajor axis would remain relatively

constant at from 302 to 503 km (188 to 312.5 statute miles) until shortly beforeencounter.

It was also felt that tracking the spacecraft for a few hours once or twice each

week from a single station would result in the same orbit accuracy as continuous

tracking during the heliocentric portion of the flight. Tracking during this phase

could provide an estimate of the astronomical unit with a standard deviation of

26



TRA JECTOR Y

less than 503 km (312.5 statute miles). Tracking data for the encounter phase

could reduce current estimates of the uncertainty of the mass of Mars from 0.1 to

2.0 percent to a value no greater than 0.001 percent.

MARINER IV TRAJECTORY RESULTS

Orbit-Determination Procedures

For the Mariner IV flight, the reduction of AFETR data and the generation

of tracking predictions proceeded as scheduled from launch through signal acqui-

sition by the Woomera Deep Space Station (DSS) 41. All results were near

nominal. During the first few hours of the mission, the primary orbit-determina-

tion function was the computation of the tracking predictions. These were

necessary so that the deep space station could establish and maintain contact with

Mariner IV. The first data obtained after launch resulted from tracking the Agena

D while the Agena D/spacecraft combination coasted in a parking orbit. Collected

at the AFETR and processed, these data provided an estimate of the initial con-

ditions of the parking orbit.

Transfer-orbit predictions were generated by assuming a nominal second

Agena D burn and were transmitted to the Space Flight Operations Facility

(SFOF) at JPL for comparison with prelaunch nominal predictions. In addition,

the initial conditions from AFETR orbit computation and the raw data were also

transmitted to SFOF. From SFOF, the transfer-orbit predictions were passed onto various deep space stations. Once the second Agena D burn occurred, the

Agena D was again tracked by AFETR, and the data were processed to provide

an estimate of the transfer orbit. (Although the Agena D orbit differed from that

of the spacecraft after separation, the difference was small for a sufficiently long

period that predictions based on the Agena D orbit could be used to acquire the

spacecraft signal.) The procedure used during the parking orbit was then re-

peated, with predictions, initial conditions, and raw data being sent to SFOF.

Once Woomera DSS 41 acquired the spacecraft signal, the tracking data

were transmitted to SFOF and processed to provide a more accurate orbit and a

better set of predictions. The first fit to DSIF data indicated a near-standard tra-

jectory with a correctable miss. All subsequent orbit computations were consistent

with the original fit, and, 16 hours after launch, the computed miss parameters

had converged to nearly constant values. The primary task then became the

determination of the most accurate orbit possible for use in the computation of the

midcourse maneuver.

27



MARINER-MARS1964

Continuous orbit updating was performed from launch until 20 hours after

launch. From that time until the midcourse maneuver, the orbit was updated

approximately every 6 hours. Tracking data monitoring was nearly continuous

during that time. The orbit used for the computation of the December 5 midcourse

maneuver included data through the first 6 days after launch. Throughout the

premaneuver portion of the flight, the quality of the tracking data was excellent.

Two-way coverage was nearly continuous, and, with the exception of certain

isolated points, no "bad" data were detected.

During the cruise period following the midcourse maneuver, tracking data

monitor runs were made to validate quality every 2 or 3 days. The new data were

used to update the tracking data file. Orbits were computed (in real time) at an

average of one every 2 weeks during this cruise period until 2 weeks before en-

counter. As the Mars encounter approached, the orbit-determination effortincreased. Orbits were then computed daily.

Resulting Trajectory Values

The values of the aiming point parameters (at the time of closest approach)

which were desired by the execution of the midcourse maneuver were as follows:

B = 12 079 km (7507 statute miles)

B-T = 6042 km (3755 statute miles)

B-R = 10 460 km (6501 statute miles)

Time = 01:47:00 GMT, July 15, 1965

(All distances expressed here are distances from the center of Mars, and the time

refers to Greenwich Mean Time (GMT).) The values of the aiming point param-

eters actually obtained by the execution of the maneuver (as computed from the

best available values from 5 days before until 5 days after encounter) were as

follows:

B = 15 3364-10 km (9532±6 statute miles)

B.T = 8188=k20 km (50894-12 statute miles)

B.R = 12 9704-20 km (80614-12 statute miles)

Time = 01:00:58 (4-10 sec) GMT, July 15, 1965

By subtracting these actual values from the desired values, the orbit errors were

determined to be the following:

Error in B = 3257 km (2025 statute miles)

Error in BoT = 2146 km (1334 statute miles)

Error in BoR = 2510 km (1560 statute miles)

Error in time = -46 min 2 sec

28



TRAJECTORY

0 o __

330 ° 300 °

/ X \

270" 240" 210"

400

_MINAL/

,COUNTER, _OMIN/A/L _ <

/

180 °

30° 60" 90 ° i 20 ° 150 o

F[OURE 2-9.--Near-Mars trajectory, as computed by using premidcourse, preencounter, and

postencounter (actual) values of aiming point parameters.

A comparison of the encounter trajectories resulting from using premidcourse,

preencounter, and postencounter (actual) aiming point parameter values is given

in figure 2-9. The errors in predicting the aiming point parameters (predicted

minus actual values) for 10 hours before encounter to 6 hours after encounter are

plotted in figure 2-10. The error in the aiming point was approximately three-

fourths of the specified 1 _ (standard-deviation) requirement established for the

mission. The allowable miss at the target caused by both uncertainties in orbit

determination and maneuver-execution errors was specified to be less than 4675

km (2906 statute miles), with the requirement that errors caused by uncertainties

in premidcourse orbit determination be less than 2262 km (1406 statute miles).

The premidcourse orbit errors are difficult to determine since the estimates are

very dependent upon the values of the astronomical unit (AU) and solar pressure

which are used. Preliminary results showed an error of 1006 to 1508 km (625 to

937 statute miles) in BoT and an error of 302 km (188 statute miles) in B.R to be

the most likely premidcourse orbit errors.

During the mission there was no evidence of errors of this magnitude. (The

results of four postmidcourse orbit computations are shown in fig. 2-11.) Since

the maneuver command had been transmitted 10 min 35 sec earlier than the time

for which the maneuver was originally computed, it was expected that the aiming

29



MARINER-MARS 1964

0

400

2OO

rr0

-200bJ

Z

0

0

-600

/

/J

/.1o--

J

/

_7 ERROR IN TIME OF CLOSEST

APPROACH, sec

rl ERROR IN B" R, krn

Z_ ERROR IN B, krn

0 ERROR IN B" 1", km

-I000

-iO -8 -6 -4 -2 E 2 4 6

TIME FROM ENCOUNTER, hr

FIGURE 2-10.--Errors in predicting aiming point parameters as a function of time

from encounter.

point would be approximately 116 km (72 statute miles) away from the nominal

location in the +T direction. Also, normal performance variations in the space-

craft guidance equipment and rocket motor would be expected to result in errors

in the velocity increment, causing a residual error in the arrival parameters.

Although the orbit parameter values were stable during the first few months

of the mission, certain fluctuations of 503 to 1006 km (312.5 to 625 statute miles)

were observed late in the mission due to fluctuations in the solutions of such items

as the astronomical unit, upon which the spacecraft trajectory is dependent. (The

fact that the data were not strong enough during the first few months of the mission

to obtain solutions of these various items accounts for the stability of the orbit

parameter values during that time.) In fact, the fluctuations in the aiming point

parameters were intensified when, approximately 1 month before encounter, the

3O



TRAJECTORY

210

180

150'

240

IMPACT CIRCLE,

/MARS SURFACE

.f

\DISPERSION

LIPSE (95%)

SUN OCCULTATION CONTOUR

(_ Isf POSTMIDCOURSE ORBITt" 2d POSTMiDCOURSE ORBIT

_@_) 3d P OSTMIDCOURSE O0/_BRB,4th POSTMIDCOURSE

120

270 500

CONTOUR

330

;AIMING REGI ON

-- AIMING POINT

I

IO 000

90 60

20000 mi

f"OCCULTATION

CONTOUR

T 0

3O

FIGURE 2-11.--Point of closest approach as determined by postmidcourse orbit computations.

mass of Mars was added to the solution set. Each new value for a solution became

the a priori value the next time the orbit was computed; thus, no set of orbit com-

putations obtained during that period used the same a priori assumptions.

During the last 5 hours before encounter, two computers were operating: one

using a 1_ value of 503 km (312.5 statute miles) for spacecraft position and the

other using a value of 201 km (125 statute miles). Both used a 1_ value of 0.1 m/sec

(approximately 0.3 ft/sec) for spacecraft velocity. At approximately 1 hour before

encounter, both computers used the 503-km (312.5-statute-mile) value for space-

craft position.

The values given for the aiming point parameters actually achieved, plus

solutions for the geocentric position and velocity, the solar pressure, the station

locations, the astronomical unit, and the mass of Mars for 5 days before until 5

days after encounter, were obtained from computations performed 3 _ months

31



MARINER-MARS 1964

,w

k

after the Mars encounter (i.e., on October 30, 1965). Table 2-I lists the a priori

1_ statistics which served as input for the orbit-determination computer program.

Doppler tracking data were used in the fit. The date to which the initial

position and velocity were referenced was July 10, 1965. When the solutions were

obtained, the geocentric position had changed by slightly less than 1006 km (625

statute miles), the geocentric velocity by approximately 1 m/sec (3 ft/sec), the

solar pressure by 0.005 percent, and the station locations by less than 10 m (0.006

statute mile). The solution for the astronomical unit and its associated 1 o- value as

determined by the orbit-determination program was 1504389564-241 km

(93 498 419+ 150 statute miles), as compared with its a priori value of 150 439 992

+2011 kin (93 499 063+1250 statute miles). The solution for the mass ratio of

Mars to the Sun was (0.322728+0.000015)×10 .6 , as compared with its a

priori value of (0.322804224-0.006456) × 10 -_.

In other computed encounter orbits, the solutions for the mass ratio of Mars

to the Sun varied less than 0.01 percent. However, for reasons probably connected

with the stability of the single-precision (8-digit) computer program, a solution

which included the astronomical unit always resulted in larger residuals between

computed and observed data. Still, the numerical value of the solution fell within

a range of 201 km (125 statute miles) of the value quoted previously. Even for

orbits fitting data from the midcourse maneuver through 5 days after encounter,

the range of values of the mass ratio of Mars to the Sun was less than 0.1 percent,

Table 2-[.--Input for orbit-determination computer program

Parameter A priori value A priori 1¢ value

Geocentric position (spherical):

Kilometers .......................................

Statute miles .....................................

Geocentric velocity (spherical):

Meters/second ....................................

Feet/second ......................................

Mass ratio of Mars to Sun . .. ... ... .. ... ... .. ... ... .. ..

Solar pressure coefficient ...............................

Astronomical unit (as determined by JPL Venus radar-

bounce experiment) :Kilometers .......................................

Statute miles .....................................

Station locations (spherical): a

Meters ..........................................

Statute miles .....................................

0.32280422 X 10 6

1.2067

10 056

6250

149 598 500

93 499 063

10

30

2% or 0.006456X10 6

5 % or 0.06

2011

1250

50

0.03

aValues obtained from surveys, Ranger project results, and preencounter Mariner IV solutions.

32



Ii

TRAJECTORY

and the range of values of the astronomical unit was less than 1006 km (625 statute

miles). These latter results are surprising because no attempt was made to solve

for Earth-Mars ephemeris errors and also because the single-precision computer

program is not entirely suitable for computing such orbits.

The uncertainty in the radius of Mars as obtained from astronomical meas-

urements was approximately 50 km (3! statute miles). The time that Mariner IV

entered the occultation region (as indicated by the loss of the spacecraft radio-

frequency (RF) signal) compared very favorably with the time predicted by the

encounter orbit, which used an average value for the Mars radius of 3397 km

(2111 statute miles). This favorable comparison indicated that the encounter orbit

was valid to within at least 50 km (31 statute miles).

A remarkable reduction in orbit parameter error is realized when the valuefor the astronomical unit obtained from the orbit-determination computer pro-

gram is used in determining the orbit. This decrease is evident in plots of the

various errors for various times after the midcourse maneuver. These plots, given

2OO

-200

-40C

-6oo

CL

C: -8001

or_LLJ

-1000

-1200

-I_00

CLOSEST APPROACH

I

TO MARS

r

f--._____

/.A

I JAN I FEBAU,499770m

U = 149 598 500 km --

REAL-TIME RUNS

I II MAR I APR I MAY I JUNE I JULY I

MONTH, /965

FIOURE 2--12.--Error in predicting B. T as a function of date for various values of the

astronomical unit (AU).

33



MARINER-MARS 1964

200

0

E -2001

-4oJz

_.)

a. -600

z_

-8OO

w

- i O00

-1200

JAN

J

L

=

(_ AU = 149597470 km

• AU = J49.598500 km

A REAL-TIME RUNS

I I IFEB I MAR I APR I MAY

I

IJ_.--=

II

,IJUNE I JULY

MONTH, 1965

FIGURE 2--1 3.--Error in predicting B. R as a function of date for various values of the

astronomical unit (AU).

in figures 2-12 to 2-14, were made after the end of the mission with all a priori

conditions held constant and data added in 1-month blocks. The errors represent

predicted minus actual values. It should be stressed that this reduction in error isnot offered as proof of the validity of the smaller value of the astronomical unit,

but is intended to suggest that the currently accepted larger value used in deter-

mining the encounter orbit might account for a large part of the error of that orbit.

The new value is not yet an accepted value, and the appropriate 1 _ value remains

2011 km (1250 statute miles).

The errors in the predicted encounter orbit were definitely the result of more

than one cause, however. The fact that the computations were done primarily in

single-precision arithmetic constitutes a deficiency of the computer program.

When the new double-precision orbit-determination program becomes available,

much more accurate orbit determinations will be possible, and it is anticipatedthat a definitive analysis of the Mariner IV orbit errors will be enabled. Another

deficiency was that the capability for properly solving for the Earth-Sun and Sun-

Mars ephemeris elements was not available in a checked-out orbit-determination

program. Since the motion of the solar pressure vanes was somewhat larger than

expected, their position was not properly represented in the computer program as

34



T

TRAJECTOR

-r

0

o_

l-

ID

600

48O

I I I ' [AU = 149 597470 km

• AU : 149 598 500 km I

IEAL-TIME RUNS

I iI

CLOSEST

APPROACH

TO MARS --

4

360

L9

Z

I--

n,-

(3.

Z

0(rre"

_J

120

I JAN ] FEB I MAR ] APR I MAY I JUNE I JULY I AUG

MONTH, 1965

FIGURE 2-14.--Error in predicting time of closest approach as a fimction of date

for various values of the astronomical unit (AU).

an average value. The telemetry data indicated random forces as large or larger

than 0.5 percent of the solar pressure force due to valve leakage in the attitude-

control subsystem. The forces of the magnitude encountered could result in B-

plane errors of 201 to 402 km (125 to 250 statute miles). There is considerable

doubt that these random forces can ever be completely "fit out" of the solution

because of their nondeterministic character.

Orbital Elements of Mariner IV Trajectory

The geocentric (Earth-centered) characteristics of the Mariner IV trajectory

are listed in table 2-II, and certain orbital data for Earth and Mars are given for

reference in table 2-III. Tables 2-IV and 2-V list the heliocentric (Sun-centered)

and areocentric (Mars-centered) orbital elements, respectively, of the trajectory of

Mariner IV. Various trajectory elements are also included as part of the chro-

nology in chapter 4.

35



MARINER-MARS 1964

J

From launch to closest approach to Mars, Mariner IV traveled approxi-

mately 523 million km (325 million statute miles) on a heliocentric trajectory

which was slightly above the ecliptic plane (fig. 2-15). The inclination of the orbitbefore encounter was slightly more than 0.1 °. As the spacecraft passed close to and

below the planet, the gravitational pull of Mars raised tile trajectory so that the

Table 2-11.--Geocentric characteristicsof Mariner IV trajectory

Characteristic

Parameter:

Radius, R:

Km ........................................

Statute miles ................................

Inertial speed, V:

Km/sec ....................................

Statute miles/sec ... .. .. .. .. .. .. .. .. .. .. .. .. .

Earth-fixed speed, v:

Km/sec ....................................

Statute miles/sec ... .. .. .. .. .. .. .. .. .. .. .. .. .

Geocentric latitude, q5, deg .......................

Longitude, 0, deg ................................

Right ascension,O, deg ..........................

Path angle of inertial velocity, r, deg ..............

Azimuth of inertial velocity, 2, deg ................

Path angle of Earth-fixed velocity, % deg ...........

Azimuth of Earth-fixed velocity, a, deg .............

Time of event, T, GMT .......................... I

Date of event . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Hyperbolic orbital element:

Semimajor axis, a:

gm ........................................

Statute miles ................................

Eccentricity, e ...................................

Injection

6872.9574

4295.5984

Preencounter

Inclination to Earth's equator, i, deg ...............

Longitude of ascending node, % deg ...............

Argument of perigee, w, deg ......................

Perigee distance, p:

Km ........................................

Statute miles ................................

Time of perigee passage, T, GMT ... .. .. .. .. .. .. ..

Date of perigee passage ...........................

11.206 585

7.004 116

10.775 777

6.734 861

--28.130 141

86.212 637

20.742 333

12.650 441

90.421 969

13.165 104

90.439 744

15:07:57

Nov. 28,1964

--41 535.874

--25 959.921

1.158 074 0

28.133 045

111.637 27

245.659 52

Postmidcourse

6565.7425

4103.5891

15:03:53.852

Nov. 28,1964

2 022 402

1 264 001

3.151 569 1

1.969 730 7

142.005 96

88.753 73

15.678 090

186.016 06

142.854 06

89.557 318

47.958 569

1.271 643 8

270.006 58

16:09:25

Dec. 5, 1964

Postencounter

228 218 340142 636 463

30.785 551

19.240 969

16 550.953

10 344.346

--5.108 255 7

289.338 86

192.497 94

30.007 818

110.912 33

0.053 300 23

269.967 05

21:27:02

July 23,1965

36



TRAJECTOR Y

inclination of the spacecraft orbit was over 20 times greater than that before en-

counter. The new orbit carries the spacecraft alternately 2.7 ° above the ecliptic

and 2.8 ° below the ecliptic.

Table 2-111.mOrbital data for Earthand Mars

Characteristic Earth Mars

Length of ellipse:

Km .........................................................

Statute miles .................................................

Perihelion distance:

Km .........................................................

Statute miles .................................................

Aphelion distance :

Km .........................................................

313 270 400

195 794 000

146 147 200

91 342 000

151 123 200

Statute miles .................................................

Eccentricity of orbit ................................................

Inclination of orbit to ecliptic, deg, min of arc .........................

Period, Earth days .................................................

Inclination of equatorial plane to orbital plane, deg ....................

Period of rota tion, hr, min .. .. .. .. .. ... .. .. .. .. .. .. .. .. .. .. .. .. .. ..

94 452 000

0.0i7

0 t

365

23.5

23, 56

452 944 000

283 090 000

205 328 000

128 330 000

247 616 000

154 760 000

0.093

1 t, 51 t

687

25.2

24, 37

Table 2-1V.mHeliocentric orbital elementsof Mariner IV trajectory

Elliptical orbital element Preencounter orbit Postencounterorbit

Semimajor axis, a:

Km .................................................

Statute miles .........................................

Eccentricity, e ............................................

Inclination to the ecliptic, i, deg .............................

Longitude of ascending node, _, deg .........................

Argument of perihelion, w, deg .............................

Perihelion distance, p:Km ..................................................

Statute miles .........................................

Time of perihelion passage, T, GMT .........................

Date of per ihel ion passage a . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .

Period, P, days ...........................................

190 929 830

119 331 144

0.227 502 96

0.125 699 63

68.665 534

352.565 27

147 492 730

92 182 956

23:11:28

Nov. 23,1964

526.645 30

200 588 100

125 367 563

0.173 220 07

2.543 740 1

226.755 45

200.649 08

165 842 220

103 651 388

07:25:19

Nov. 16, 1964

567.113 21

aThe Mariner IV launch occurred past the perihelion point of its heliocentric orbit.

37



MARINER-MARS 1964

Table 2-V.mAreocentr ic (Mars-centered) orbital elements of Mariner IV trajectory

Hyperbolic orbital element

Semimajor axis, a:

Kin .................................................

Statute miles .........................................

Eccentricity, e ............................................

Inclination to the ecliptic, i, deg .............................

Longitude of ascending node, _, deg .........................

Argument of periapsis, 6o, deg ...............................

Per iapsis dis tance, p:

Km .................................................

Statute miles .........................................

Time of periapsis passage, 7", GMT ..........................

Date of per iapsis passage . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Preencounter prediction

-22 046

--13 779

0.658 94

60.458

188.009

289.546

12 322

7 701

01:04:49.5

July 15, 1965

Actual Mars-encounter orbit

--22 092

--13 808

0.697 53

58.186

187.499

289.321

13 201

8 251

01:00:58.1

]uly 15, 1965

/-EARTH AT LAUNCH

/ NOVEMBER 28, 1964

I .f" ..si---_ 6o

_o_,-,-'t \ i'_00,. /\.. _ ____,_o 2,o/,- I

90 G)" / - "-'_. ,50 18o _l/ /

\ ,,'\. __'--i_li/..u. , /

,_o_#,,__,.....--_....--,,

,o ,oLo,;::;..JULY 15, 1965

FIGURE 2-15.--Mariner IV trajectory.

38



TRAJECTORY

The pull of the gravitational field of Mars also added energy to the spacecraft.

Because this energy was added nearer aphelion (point in orbit farthest from the

Sun) than perihelion (point in orbit nearest to the Sun), the effect was to make the

orbit more nearly circular. Thus, the postencounter orbit of Mariner IV will

never pass as close to the Sun as the orbit beginning at launch. The preencounter

and postencounter orbits of Mariner IV are shown in figure 2-16.

The postencounter orbit has been computed for the 7 years following launch.

2[0 ° 180 ° 150 °

POSTENCOUNTER PREENCOUNTER

TRAJEC]

120 °

270 ° 90 °

300 ° 60 °

330 o 0 ° 30 o

FIouRE 2-16.--Heliocentric plan view of Mariner IV preencounter and postencounterorbits.

39



MARINER-MARS 1964

s

360 --

320

280

x

E

.24C

m 20C

o

160

±

Ld 120

8O

,o:/...o .¢

1964 i965

NCOUNTER

l/]_FIRST CLOSE APPROACH

| TO EARTH

.................... I,,,,,I ..... I,,,,,L ....

1966 1967 1968 1969

DATE

_COND CL/OSE APPROACH

TO EARTH

1970 1971

FmURE 2-17.--Mariner IV distance from Earth, 1965-1971.

In these 7 years, the spacecraft will twice pass relatively close to Earth, once in

September 1967 and again in July 1970 (fig. 2-17.) At its first close approach to

Earth, Mariner IV will be about 50 million km (31 million statute miles) away.

This range is within the communications capability of the spacecraft low-gain

omnidirectional antenna. Provided the radio subsystem is still operable at that

time, an attempt will be made to reaequire Mariner IV in September 1967.

40



CHAPTER 3

Space Vehicle System Design and Testing Operations

LAUNCH VEHICLE DESIGN

The Mariner-Mars 1964 launch vehicle consisted of an Atlas D first stage to

provide the thrust (propelling force) required to place the Agena D/spacecraft

combination on the required ascent trajectory, and an Agena D second stage toprovide the dual-burn capability and attitude control required to place the space-

craft in a predetermined geocentric (Earth-centered) transfer orbit. A shroud and

adapter structure protected the spacecraft from the external environment during

_t_ ascent mruugn that portion of the Earth's atmosphere offering resistance to its

passage. A cutaway view of the launch vehicle is shown in figure 3-1.

Atlas D First Stage

The outer structure of the Atlas D, called the airframe, was a stainless-steel

cylinder. Into this airframe were integrated the main propellant tank section, the

aft (bottom of Atlas D when mounted on launch pad) booster section which could

be jettisoned, and the forward interstage adapter section. Equipment pods on the

ATLAS D/AGENA DADAPTER -7 SHROUD x,

BOOSTER SUSTAINER /

MARINER

MARS 1964

ATLAS D SPACECRAFT

AGENA D SPACECRAFT

ADAPTER RING

FIOURE 3-1.--Exploded view of Atlas D/Agena D launch vehicle.

41



MARINER-MARS1964

outside of the main section housed the necessary electrical and electronic units and

other components.

The main section housed in a single 253 547-N (57 000-1b) thrust Rocket-

dyne sustainer engine (one thrust chamber, designed for high-altitude perform-

ance) and two 2224-N (500-1b) vernier (fine steering) engines (one thrust chamber

each, for attitude control and velocity trim). The vernier engines were derated

from a thrust of 4448 N (1000 lb) to 2980 N (670 lb) to reduce the amount of

propellant consumed by those engines, thus making more propellant available

for the main engines to deliver higher effective specific impulse. 1 When assembled,

this section was approximately 20.1 meters (67 ft) long and approximately 3

meters (10 ft) in diameter at its base. Its weight when fully fueled was approxi-

mately 117 930 kg (260 000 lb).

The booster consisted of two 667 230-N (150 000-1b) thrust engines (one

thrust chamber each, designed for low-altitude performance). The booster

propellants were delivered, under pressure, from the main propellant tank section

to the combustion chamber. The width of the booster section was approximately

4.8 meters (16 ft), and its weight was approximately 3175 kg (7000 lb).

All three types of engines in the Atlas D stage started and developed their

full rated thrust while the vehicle was held on the launch pad. After liftoff and

after the booster engines burned out (approximately 2 min after launch; called

booster-engine cutoff or BECO), the booster-engine section jettisoned or sepa-

rated from the main section. The sustainer engine continued to burn until no morethrust was produced (approximately 5 min after launch; called sustainer-engine

cutoff or SECO). The swiveled vernier engines then provided the final correction

in velocity and attitude before they shut down (called vernier-engine cutoff

or VECO).

The fuel used by the Atlas D rocket-engine system was RP-1 (kerosene), and

the oxidizer was liquid oxygen. Primary power was supplied by a remotely

activated silver-zinc battery, and secondary power was supplied by a three-phase

inverter (a dc-to-ac conversion device). A telemetry system monitored and trans-

mitted functional conditions from prior to launch until Atlas D/Agena D separa-

tion, after which it ceased operation. No aerodynamic control surfaces such as

fins and rudders were necessary since the Atlas D was stabilized and controlled

by "gimbaling" or swiveling the engine thrust chambers by means of two inde-

pendent hydraulic pressure systems. A standard pneumatic system provided: (1)

l Specific impulse is a ratio of the thrust developed to the amount of propellant required to produce that thrust.

42



SPACEVEHICLESYSTEMDESIGNAND_TESTINGOPERATIONS

continuous adequate pressurization of the main propellant tanks, hydraulic res-

ervoirs, and engine lubrication oil tanks; (2) pressure for engine control functions;

(3) inflight repressurization of various propellant tanks; and (4) activation pres-

sure for the booster separation systems. The helium gas used to support the Atlas D

structure was expanded to the proper pressure by the heat of the exhaust from the

engine gas generator. This use of pressurized helium gas for support instead of

internal bracing resulted in a considerable weight savings.

The Atlas D used a flight programer, an autopilot, and 10 gimbaled thrust-

chamber actuator assemblies for stabilizing and steering the vehicle along the

desired trajectory by controlling the orientation of the engine thrust vectors. The

autopilot was set before launch to control the attitude of the vehicle automatically.

Guidance commands were furnished by a ground radio guidance system andcomputer. The radio inertial guidance system used two radio beacons which

responded to the ground radar. A guidance system decoder processed the guidance

commands, and the necessary steering corrections were then monitored to the

flight-control system. Pitch steering commands in such a vehicle are normally

generated by the flight programer from launch until the booster is jettisoned, and

by a combination of the flight programer and the guidance system from booster

jettison until the end of the Atlas D sustainer-powered flight. However, in order

to permit early correction of deviations from the nominal ascent trajectory, the

capability to steer the vehicle by use of the guidance system prior to booster

jettison was implemented in the Atlas D stage used for the Mariner-Mars 1964

missions. (Actually, this capability was not required for the Mariner IV flight

since the small deviation from the nominal trajectory made its usage unnecessary.)

A significant gain in payload was made possible by new guidance equations

which determined more efficient Atlas D sustainer steering commands. Formerly,

steering commands were generated which attempted to return the vehicle as

rapidly as possible to the nominal flight path. These new commands attempted to

correct the vehicle trajectory only enough to arrive at the proper aiming point.

Agena D Second Stage

The overall nominal thrust rating of the single Agena D rocket engine was

71 171 N (16 000 lb) in vacuum, with a nominal thrust duration of 240 see. The

outer structure of the Agena D (the airframe) was a cylinder approximately 1.5

meters (5 ft) in diameter and 6.29 meters (20.6 ft) long (as measured from the

rearward end of the engine nozzle to the forward section of the forward equip-

ment rack). When fueled, its weight was approximately 6940 kg (15 300 lb).

43



MARINER-MARS1964t--

Four major sections made up the Agena D second stage: the forward section,

the tank section, the rearward section, and the booster adapter section. The for-

ward section carried guidance, flight-control electronics, telemetry, command,tracking, electrical power, and propellant-pressurization equipment. The tank

section stored the fuel (unsymmetrical dimethylhydrazine) and oxidizer (inhibited

red fuming nitric acid) necessary for operation of the main rocket engine and also

provided the support structure between the forward and rearward sections. The

rearward section provided for structural support and attachment of the rocket-

engine assembly, pneumatic attitude-control thrust valves, and nitrogen storage

spheres. Ready access to all parts of the engine, plumbing, and wiring was made

possible by the rearward section open-frame design. The booster adapter section

was the interconnecting structure between the Agena D second stage and the

Atlas D first stage; housed within this section were two retrorockets used for sepa-

ration of the Atlas D from the Agena D. This section remained with the Atlas D

upon separation of the two stages.

The propulsion system was composed of an engine with a dual-restart capa-

bility and a propellant and pressurization subsystem. The capability to restart the

engine after it had already been fired once in order to reach an Earth orbital

speed made possible a significant increase in payload and a change of orbital

altitude. Propellant tanks were pressurized with helium to insure proper propel-

lant pump operation. The propellant tank and sump designs were improvements

over previous designs in that the following were provided: (1) improved scaveng-

ing (removal of burned gases from the cylinder), and (2) containment of sufficient

propellant within the tank sumps (reservoirs) to eliminate the need for ullage

rockets. (However, since flight qualification of this feature for the Mariner-Mars

1964 missions was not possible prior to launch, a positive continuous ullage

control system was included to insure the second Agena D start.)

Power was supplied by two primary batteries, a single- and three-phase

inverter, and a dc-to-dc converter. The communications system monitored and

measured, by electrical signals, functional and environmental conditions of the

Agena D and the spacecraft during the ascent phase to spacecraft injection and

during the Agena D retromaneuver.The Agena D guidance and control system served to: (1) maintain the Agena

D at the proper attitude at all times, (2) provide switch closures at the proper

times to accomplish the desired sequence of events during ascent, and (3) provide

the propulsion shutdown signal after the desired velocity had been achieved.

Major components of the system were an inertial reference package, a horizon

44



,q

SPACEVEHICLESYSTEMDESIGNAND TESTINGOPERATIONS

sensor, a flight-control electronics unit, a hydraulic attitude-control system, a

pneumatic attitude-control system, a velocity meter, and two electromechanical

timers. Error signals from the inertial reference package were applied to the flight-

control electronics unit for processing and were then applied to the hydraulic and

pneumatic attitude-control systems. Two hydraulic actuators "gimbaled" the

engine, providing thrust-vector control; six nitrogen-gas attitude-control jets were

provided to torque the vehicle.

Shroud System and Spacecraft Adapter

A shroud is the covering which protects a spacecraft during the prelaunch

checkout and the ascent phase of flight. Since previous shrouds were too small, anew design for the Mariner-Mars 1964 spacecraft was formulated which used a

one-piece fiber-glass-and-magnesium section connected to a beryllium nose dome.

This new shroud was to be separated after launch by springs which would eject

it forward "over the nose" from the spacecraft. However, after the first Mariner-

Mars 1964 spacecraft (Mariner III) launch on November 5, 1964, this shroud

failed to eject clear of the spacecraft. The failure mode was a structural failure

caused by skin separation from the fiberglass honeycomb core.

Efforts were then directed toward the immediate development of an all-metal

shroud for use on the second Mariner-Mars 1964 spacecraft (Mariner IV). In

this design, the fiber-glass structure was replaced by a magnesium section with an

inner thermal liner. The first replacement shroud was completed within 17 days

after the Mariner III launch. To compensate for the additional weight of the new

shroud, the Agena D retrorocket (for use after injection to insure that the Agena D

would not impact the planet) was removed, and an Agena D command-destruct

unit was replaced by a self-destruct system (which separated with the Atlas D in-

stead of at injection). The removal of the retrorocket was possible because of a

change in the mission profile (discussed in oh. 2). During the Mariner IV

mission, this all-metal shroud successfully ejected over the spacecraft without

striking it and did not subsequently collide with the spacecraft.

The shroud-to-Agena D attachment structure was part of the spacecraft-to-

Agena D attachment structure called the spacecraft adapter. As illustrated in

figure 3-2, the basic structure was comprised of two conical sections, one con-

verging from the common base ring, through which the adapter was attached to

the Agena D forward equipment rack, to the mounting plane of the spacecraft,

and the other diverging from the common base ring to the shroud mounting ring.

45



MARINER-MARS 1964

4GENA D S i

FORWARD

EQUIPMENT

RACK

SHROUD

ASPACECRAFT MOUNTING PLANE

// ,/

COMMON BASE RING

u

-L

FIGURE 3-2.--Spacecraft adapter.

The attachment of the adapter to the shroud and to the spacecraft was released by

firing explosive squibs. Shroud ejection was accomplished by four ejection spring

assemblies mounted in lower end of the shroud. The spacecraft/Agena D separa-

tion system used four equally spaced spring mechanisms mounted on the adapter

that acted against four mating pads on the spacecraft. Most measuring and

monitoring instrumentation was mounted on the adapter.

SPACECRAFT SYSTEM DESIGN

Design Considerations

Although vital knowledge and experience were gained during the Mariner-

Venus 1962 project, the problems involved in sending a spacecraft on a mission to

Mars in 1964 were more numerous and more complex. The energy required to

ship a pound of payload to Mars in 1964 was actually slightly less than that needed

for the trip to Venus in 1962. However, this slightly lower energy requirement was

practically the only aspect of the Mars mission that was not considerably more

difficult than the comparable aspect in the Venus mission.

One example of the difficulties was service life. Mariner II had to operate for

about 2500 hours on its flight to Venus, while the Mariner-Mars 1964 spacecrafthad to be designed for 6000 to 7000 hours of flight life for the trip to Mars and

beyond.

Electrical power was another consideration. While only a small amount (less

than 200 W) was required, the electrical power had to come from sunlight. The

amount of available power was, of course, dependent upon the distance from the

spacecraft to the Sun. Mariner II had one of its two solar panels partially disabled

46



h_


enroute to Venus, but, since it was going closer to the Sun in its flight, the space-

craft drew nearly as much power from the undamaged panel at Venus as it had

received from both panels near Earth. However, in going from Earth to Mars

rather than to Venus, a spacecraft would be receding from the Sun and thus the

solar power would decrease. Therefore, the Mariner-Mars 1964 spacecraft had to

have more than twice the solar panel area of Mariner II, or approximately 6.5!

sq m (70 sq ft) instead of just 2.51 sq m (27 sq ft).

The decrease in solar radiation would cause the spacecraft to become colder

during the trip, rather than hotter as in the Mariner II mission. This fact had to

be considered in the design of the temperature-control devices for the spacecraft.

Beyond Mars is the asteroid belt, consisting of thousands of planetoids in

independent solar orbits. Therefore, astronomers believed that the meteoric

intensity might increase in the direction of this belt. In addition, the Mars path

lay across several "cometary" meteor streams. The spacecraft might encounter

more space dust than had been experienced during the Mariner II mission, where

only two impacts had been recorded by the spacecraft's detector. Even this

number is somewhat misleading if it is not realized that the total area of the

Mariner II spacecraft was about 200 times that of its small dust detector; there-

fore, the detector recorded only a fraction of the particles actually hitting the

spacecraft.

Simple distance to be traveled was another consideration in spacecraft design.

When Mariner II was at a maximum distance of 86.9 million km (54 millionstatute miles), radio waves from it took nearly 5 min to reach the Earth. Since

communications during the Mariner-Mars 1964 missions would involve distances

of at least 241 million km (150 million statute miles), the delay would be about

three times as long. (During the actual Mariner IV mission, the communications

time delay reached a maximum of 17 min 6 sec at the end of the mission on

October 1, 1965, when Mariner IV was approximately 306 million km (190

million statute miles) from the Earth.) Therefore, the communications system

had to be more powerful--actually nine times more powerful since radio strength

decreases as the square of increasing distance. Both the ground and flight units

had to be improved.

Preliminary Design

The flight plan for the Mariner-Mars 1964 missions was the same as that for

the Mariner II mission; i.e. : the launch by an Atlas/Agena, attainment of a cruise

47



MARINER-MARS1964

0

mode, the performance of a trajectory correction 2 as necessary early in the flight,

the gathering of interplanetary data, and the turn-on of planetary experiments for

data gathering about the planet. The basic design of the Mariner-Mars 1964spacecraft for which this flight plan was applicable is illustrated in figure 3-3.

The spacecraft, fully stabilized in attitude, used the Sun and the star Canopus

as references. Cold-gas jets pointed the spacecraft in all three axes (pitch, yaw, and

roll), and external torques were counteracted in two of the axes by changing the

positions of movable solar pressure vanes. Gyroscopes were available for initial

acquisition and for inertial control during the midcourse maneuver.

For power, photovoltaic cells were arranged on panels with a body-fixed

orientation for cruise operations, and a rechargeable battery was provided for

launch, midcourse maneuvers, and backup support. Power-conversion equipment

delivered regulated 2400-hertz square-wave, 400-hertz, and unregulated dc elec-

tricity for distribution to the spacecraft subsystems. A central computer and

sequencer, an extremely accurate electronic clock, provided synchronization

signals for frequency regulation and performed the sequencing of switching

operations on the spacecraft. A guidance system permitting midcourse maneuvers

and a propulsion system capable of executing two such maneuvers were included.

A two-way S-band (2110 to 2120 MHz for Earth-to-spacecraft transmission

and 2290 to 2300 MHz for spacecraft-to-Earth transmission) communications

system carried a steady stream of telemetry information to Earth; commands to

the spacecraft; and angle-tracking, Doppler, and ranging information for orbit

determination. A low-gain antenna and a fixed high-gain antenna were included

in the radio system of the spacecraft. Either of these antennas could be used to

transmit or receive. Switching between the antennas was accomplished by logic

on the spacecraft or by ground command. The command system detected and

decoded incoming command messages and passed them to the various equipment

on the spacecraft. Two types of commands were possible: a direct command (DC),

which resulted in direct action by the receiving system; and a quantitative com-

mand (QCI-1, -2, and -3), which was transferred to the central computer and

sequencer to be stored for tgiter use. A data encoder formatted, sequenced, and,

as necessary, provided analog-to-digital conversion of the telemetry data.

2Commonly called the "midcourse-correction maneuver" or simply "midcourse maneuver." Actually theselermsare misnomers since the maneuver occursearlier than the midpoint ofthe flight, and, rather than correctinga mistake, the maneuver increasespossible accuracy. However, since these are the accepted terms describing the

maneuver, they are used throughout this document.

48



SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

(o) TOP VIEW

LOW-GAIN ANTENNA--

CSS_rC DETE CTOR-- _'_.4

-Y__SOLAR PANEL

TRR _;PETDION DET_J

PLASMARO._____sE,,,SORJ_-_..__ 1____

-- "?'--_/----d._'OATAJ.,AN0AW _O%%,N0LBAy.,

GAS JETS

/-HIGH-GAIN ANTENNA

// * PITCH

_RD

Z._N SENSO_ \_-BAY I POWER \ /

PLL_---BAY-- 11" POSTINJECTION

PROPULSION SYSTEM

SCIENTIFIC EQUIPMENT AND

DATA AUTOMATION SYSTEM

(b) BOTTOM VIEW

LOW _ GAIN

ANTENNA WAVEGUID E _-_ _,_.| F _-MAGNETOMETER

"1 ) / SENSOR SO'AR

I_/ P_ESSURE

/1 _ I_ rSOLAR PANEL DAMPER VANE--_

_._,,,r'_. _\. _ II I AND PI NPULLER __

URE ',_./_ I _ _._ _ ,_SOLAR PRESSJRE _,.._._11 _ I "_/--BAy'O'm POWER REGULATOR

_VANE ACTUATOR ,,,,,,,,,,,_".__._ ll_J_J--__ l _ / AND BATTERY

_/_ _ _'-__NOPUS SENSOR

SENSOR (SECONDARY BAY ]Z__ ATTITUDE CONTROLsIENsOR (SECONDARY) ANU L;.L._:_ NTR

WIDE-ANGLE ACQUISITION LENS

BAY _ RF COMMUNICATIONS

FIGURE 3-3.--Mariner-Mars 1964 spacecraft design.

49



MARINER-MARS1964P

The spacecraft carried a number of scientific instruments to measure the

particles and electromagnetic fields between the Earth and Mars and in the vi-

cinity of both planets. At planetary encounter, a television instrument took

pictures of the surface of the planet. A data automation system furnished control

and synchronization, performed necessary data conversions and encoding func-

tions, and buffered the science data, transmitting them to the data encoder at the

appropriate times and rates. Data to be prepared for direct transmission to Earth

were handled in the real-time portion of the data automation system. Data to be

stored on magnetic tape in the spacecraft's tape recorder for later transmission

to Earth at a much slower rate than that during recording were handled in the

non-real-time portion. Stored data consisted of the pictures taken by the television

camera, the science data obtained during the encounter phase, and informationon spacecraft and television system performance.

Spacecraft Mechanical Configuration

The designers of the Mariner-Mars 1964 spacecraft knew that more space-

craft weight would be needed than the 204 kg (450 lb) that the Atlas D/Agena B

launch vehicle allowed for the trip to Venus. A new version of the Agena second

stage, the Agena D, being basically a collection of improvements from better

propellant utilization to lightweight materials, added 36 kg (80 lb) to the possible

spacecraft weight, and the increase in available energy between the Mariner-Venus 1962 and Mariner-Mars 1964 missions added still more (approximately

18 kg (40 lb)) to the possible spacecraft weight. Thus, the new version of the space-

craft could weigh approximately 258 kg (570 lb).

In order to minimize nonelectronic weight within a total spacecraft weight

of about 258 kg (570 lb), it was necessary to: (1) integrate structurally the elec-

tronic packaging design into the basic spacecraft structure (i.e., make the elec-

tronics the structural foundation, thereby almost eliminating the need for a

skeletal structure), (2) minimize the number of articulated or deployed elements,

and (3) attach the spacecraft to the Agena D second stage in a direct manner. A

weight breakdown for the final spacecraft weight of 260.68 kg (574.74 lb) is given

in table 3-I. With an overall height of 2.90 meters (9.4 ft), the spacecraft was

composed of approximately 138 000 parts. The contractors and subcontractors

for the various spacecraft subsystems and components are listed in table 3-II. The

chosen mechanical configuration, illustrated in figure 3-4, is described in the

following paragraph.

5O




i

The primary spacecraft structure was an octagonal electronics compartment

approximately 45.7 cm (18 in.) high and 138.4 cm (54.5 in.) across the points.

This shape represented a minimum departure from existing design technology and

yet assured a minimum structural weight. The two octagonal rings of the structure

were separated by longerons mounted at the octagon corners. The eight rectan-

gular shallow openings formed by the rings and the longerons served to form

"bays." Seven of these bays contained spacecraft electronics assemblies, and the

eighth contained the midcourse (postinjection) propulsion system so located be-

tween two solar panels that exhaust heating of the panels would be minimized

during motor firing. The cabling in each bay was mounted on the center member

of each electronics assembly. Electrical cables for the bays were contained in two

concentric rings.

Table 3-1.--Weight summaryfor Mariner-Mars 1964 spacecraft

Item

Engineering subsystems:

Structure ..................................................................

Radio (including 3.37 kg (7.43 lb) for antennas) ................................

Data encoder ...............................................................

Video storage ..............................................................

Command ................................................................

Attitude control ...........................................................

Central computer and sequencer .. ... ... ... ... ... ... ... ... ... .... ... ... ... ... .

Power (including 35.84 kg (79.02 lb) for solar panels) ...........................

Pyrotechnics ...............................................................

Propulsion ................................................................

Temperature control .......................................................

Science subsystem and ancillary equipment:

Television ..................................................................

Helium magnetometer ......................................................

Cosmic dust detector .......................................................

Ionization chamber ........................................................

Cosmic ray telescope .......................................................

Trapped radiat ion detector . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .

Solar plasma probe .........................................................

Instrument simulator (replacing ultraviolet photometer, to maintain inertial mass

and temperature balance within spacecraft) ..................................

Data automation system ....................................................

Planetary scan system .......................................................

Spacecraft wiring ..............................................................

Total ..................................................................

Weight

kg lb

35.58 78.44

18.97 41.83

10.17 22.43

7.66 16.89

4.59 10.12

28.71 63.29

5.16 11.38

68.02 149.97

5.54 12.21

21.57 47.55

7.04 15.53

5.12 11.28

3.07 6.77

.95 2.10

1.32 2.90

1.17 2.58

.98 2.17

2.91 6.41

2.98 6.57

5.34 11.78

3.11 6.85

20.72 45.69

260.68 574.74

51



MARINER-MARS 1964

Q

,I,--

Table 3-11.--Contractors and major subcontractors for Mariner-Mars 1964 project

Name ] Responsibility

Contractors

General Dynamics/Convair (now General Dynamics/Astronautics), San

Diego, Calif.

Rocketdyne, division of North American Aviation, Inc., Canoga Park,

Calif.

General Electric Co., Defense Electronics Division, Syracuse_ N.Y .........

Burroughs Corp., Defense, Space, and Special Systems Group, Paoli, Pa...

Lockheed Missiles & Space Co., Sunnyvale, Calif .....................

Bell Aerosystems Co. , Buffalo, N.Y .. ... .. .. ... .. ... .. ... .. .. ... .. ... .

Jet Propulsion Laboratory, California Institute of Technology, Pasadena,

Calif.

Atlas D (purchased through the U.S.

Air Force Systems Command,

Space Systems Division)

Atlas D propulsion systems

Atlas D radio command guidance

Ground guidance computer

Agena D

Agena D propulsion system

Spacecraft

Major subcontractors '_

Advanced Structures Division, Whittaker Corp., La Mesa, Calif. ..........

Airite Products, division of the Electrada Corp., Los Angeles, Calif ....

Alpha-Tronics Corp., Monrovia, Calif ...............................

Anadite Co., Los Angeles , Calif . .. ... .. .. ... .. ... .. ... .. .. ... .. ... ..

Anchor Plating Co., El Monte, Calif ................................

Applied Development Corp., Monterey Park, Calif ....................

As trodata, Inc., Anahe im, Calif . .. ... .. ... .. .. ... .. ... .. ... .. .. ... .

Barnes Engineering Co., Stamford, Conn ... .. .. .. .. .. .. .. .. .. .. .. .. .

Bendix Corp., Scintilla Division, Sidney, N.Y .......................

Bergman Manufacturing Co., San Rafael, Calif .. ... .. .. ... .. ... .. ...

Cannon Electric Co., Los Angeles, Calif ............................

CBS Laboratories, division of Columbia Broadcasting System, Inc.,

Stamford, Conn.

Computer Control Co., Inc., Framingham, Mass .....................

Correlated Data Systems Corp., Glendale, Calif ......................

Data-Tronix Corp., King of Prussia, Pa .............................

Spacecraft high-gain antennas

Midcourse-propulsion fuel tanks,

nitrogen tanks

Data automation system analog-to-

pulsewidth converters

Surface treatment of structural ele-

ments and chassis

Gold plating

Ground telemetry decommutators,

printer programers

Time code generator/translators;

ground command read, write,verify equipment; encoder simu-

lator; and spacecraft system test

data system

Canopus star sensor elect ronics

Connectors

Chass is forgings

Connectors

Image dissector tubes for Canopus

star sensors

Real-time data automation system

logic cards for scientific instru-

ments; operational support equip-

ment; and data automation system

voltage-to-pulsewidth converters

Spacecraft external power source and

solar panel simulators and voltage-

controlled oscillators

Voltage-controlled oscillators

aIn addition to these subcontractors, over 1000 other individual firms contributed to the Mariner-Mars 1964

project.

52




Table 3-11.--Contractors and major subcontractors for Mariner-Mars 1964 project_Continued

I

Name ] Responsibility

Major subcontractors

Delco Radio Division, General Motors Corp., Kokomo, Ind ........... i Telemetry format simulatorsDigital Equipment Corp., Los Angeles, Calif ........................ Data automation system operational

support data system

Dunlap & Whitehead Manufacturing Co., Van Nuys, Calif. ............ Midcourse propulsion and structural

Dynamics Instrumentation Co., Monterey Park, Calif. .................

The Electric Storage Battery Co., Raleigh, N.C .......................

Electro-Optical Systems, Inc., Pasadena, Calif. .......................

Electronic Memories, Inc., Los Angeles, Calif .........................

Engineered Electronics Co., Santa Ana, Calif. ........................

Fargo Rubber Corp., Los Angeles, Calif. ............................

Franklin Electronics, Inc., Bridgeport, Pa ...........................

General Dynamics Corp., General Dynamics/Electronics, San Diego, Calif.

General Electrodynamics Corp., Garland, Tex .......................

Grindley Manufacturing Co., Los Angeles, Calif .....................

Hi-Shear Corp., Torrance, Calif. ...................................

Hughes Aircraft Co., Microwave Tube Division, Los Angeles, Calif....

IMC Magnetics Corp., Westbury, N.Y .............................

International Data Systems, Inc., Dallas, Tex ........................

Kearfott Division, General Precision, Inc., Los Angeles, Calif. .........

Lawrence Industries, Inc., Burbank, Calif. ..........................

Lockheed Aircraft Service, Inc., division of Lockheed Aircraft Corp.,

Ontario, Calif.

Lockheed Electronics Co., division of Lockheed Aircraft Corp., Los Angeles,

Calif.

Magnamill, Los Angeles, Calif ......................................

Massachusetts Institute of Technology, division of Sponsored Research,

Cambridge, Mass.

Metal Bellows Corp., Chatsworth, Calif. .............................

Milbore Co., Glendale , Calif . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Mincom Division, Minnesota Mining & Manufacturing Co. ,Los Angeles,

Calif.

elements

Ground telemetry consoles and as-

sembly of planetary scan system

electronics

Spacecraft batteries

Ionization chamber assemblies; as-

sembly and test of spacecraft solar

panels; modification and test of

spacecraft power subsystem; and

spacecraft assembly cables

Magnetic counter assemblies for

spacecraft central computer and

sequellccr

Non-real-time data automation system

Midcourse-propulsion fuel-tank

bladders

Ground telemetry high-speed digital

computers

Assembly of television subsystems

Vidicons and television tube test set

Midcourse-propulsion jet vanes; fuelmanifolds; oxidizer cartridge shell;

and supports

Squibs

Traveling wave tubes

Solar vane actuators

Ground command modulation

checker and telemetry power sup-

plies

Gyroscopes and jet vane actuators

Printed circuits

Spacecraft low-level positioners

Solar cell modules and magnetic shift

register for central computer and

sequencer

Structural elements and chassis

Solar plasma probes

Midcourse-propulsion oxidizer bellows

assembly

Midcourse-propulsion engine compo-

nents

Ground telemetry tape recorders

53



MARINER-MARS 1964

m.

Table 3-11.--Contractors and major subcontractors For /Vlariner-Mars 1964 project--Concluded

Name Responsibility

Major subcontractors

Motorola, Inc., Military Electronics Division, Scottsdale, Ariz ........

Nortronics, division of Northrop Corp., Palos Verdes Estates, Calif .....

Philco Corp., Palo Alto, Calif ...................................

Proto Spec, Pasadena, Calif . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Pyronetics, Inc., Santa Fe Springs, Calif ..........................

Rantec Corp., Calabasas, Calif ..................................

Raymond Engineering Laboratory, Inc., Middletown, Corm .........

Ryan Aeronautical Co., Aerospace Division, San Diego, Calif. .......

Siemens & Halske, A.G., Munich, West Germany ..................

Space Technology Laboratories, division of Thompson Ramo Wool-

dridge, Inc., Redondo Beach, Calif.

Sperry Utah Co., division of Sperry Rand Corp., Salt Lake City, Utah.

State University of Iowa, Iowa City, Iowa .........................

Sterer Engineering & Manufacturing Company, North Hollywood, Calif..

Texas Instruments, Inc., Apparatus Division, Dallas, Tex ............

Textron Electronics, Inc., Heliotek Division, Sylmar, Calif ...........

Thompson Ramo Wooldridge, Inc., Redondo Beach, Calif ...........

Univac, division of Sperry Rand Corp., St. Paul, Minn .............

University of Chicago, Chicago, Il l .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .

WEMS, Inc., Hawthorne, Calif ..................................

Wyman-Gordon Corp., Los Angeles, Calif .........................

Spacecraft transponders; command sys-

tems and associated operational sup-

port equipment; and DSIF equivalent

operational support equipment

Development and support of attitude-

control electronics

Integrated circuit sequence generator

system; spacecraft antenna feeds; and

spacecraft antenna subsystem tests

Chassis and subchassis

Midcourse propulsion system explosive

actuated valves

S-band circulator switches and pre-

selection and band rejection filters

Spacecraft video-storage-subsystem tape

recorder

Spacecraft solar panel structure

RF amplifier tubes

Spacecraft central computer and se-

quencer and associated operational

support equipment

Magnetometer mapping fixture

Trapped radiation detectors

Valves and regulators for attitude-con-

trol gas system

Spacecraft video-storage-subsystem elec-

tronics; spacecraft data encoders and

associated operational support equip-

ment; helium magnetometers; atti-

tude-control gyro electronics assem-

blies; and data demodulators

Silicon photovoltaic solar cells

Thermal control louvers and power

converters

Spacecraft data automation system

buffer memory

Spacecraft cosmic ray telescopes

Spacecraft television electronics modules

and spacecraft attitude-control elec-

tronics modules

Spacecraft structural forgings

54




EA RT H DE TE CT OR

SQUARE-ROOT SUM

:'NTL'

SOLAR PANEL RIN/G_ (_(

CABLE TROUGH

TRAPPED jRADIATION DETECTOR

-Y SOLAR PANEL -_FT ..'.-_PITCR GAB JETi_,'-

__1__ _ _ SUI'I i_ii_iR_L_ ROLL AND YAW GAS JEt 7

:' - ¼ 1

PLASMA PROBE _--CASE HARNESS SUPPORT STRUCTURE (TYPICAL}

(O) TOP VIEW OF BASIC OCTAGON, ONE SOLAR PANEL SHOWN IN DEPLOYED POSITION

FIGURE 3--4.--Mariner-Mars 1964 spacecraft mechanical configuration.

LO_/AVEGAIN ANTENNA

LOW-GAIN ANTENNA

G RO UN D P LAN E _

+ ROLL

SOLAR PRESSURE LOW-GAIN ANTENNA DAMPER_ ABSORPTIVITY STANDARD CANOPUS SENSOR

\SHUTTER ACTUATOR

VANE LATCH HELIUM MAGNETOMETER IONIZATION uPPER THERMAL _IELD

_, , /-_ H EL IU M M AG NE TO ME " /'_ CHAMiBER 'Q___'_ C_SENS{_,\,_ _r_ PLANESHUTTER

SOIE_E

' IS NOT SHOWN IN

TELEVISLON

CAMERA

S CI EN CE C OV ER

I

/ W IOE -AN GLE ACQUISITION

SENSOR

NARROW-ANGLE

®@@@@®®@@®@@@@ NO @

SOLAR PANEL O_UISE _SOLAR PANEL OPEN SWITCHSOLAR PRESSURE

SOLAR PANEL D_PA_ R LA_ cHCH_UI_ _ _ S OL AR P ANE L OPEN SWITCHV AN E ( ST OW ED ) DAMPER _PLASMA PROBE

(b) SIDE VIEW OF SPACECRAFT, SOLAR PANELS FOLDED

FIGURE 3-4.--Continued.

55



MARINER-MARS 1964

#

,t m

SQUARE ROOT SUN

SENSOR (PRIMARY)

RF COMMUNICATIONS

BAY

LOW-GAIN

IONIZATION

CHAMBER -

bJ

__o<lF-

20

oat_ Z

E AR TH DE TE CTOR

ATTITUDE CONTROL AND CCB, S

BAY _]

.t.

SOLAR PANEL OPEN 1"_,_ c)

DATA E NCODER D 1 _J_vJ

AND COMMAND BAY IT[

SCIENTIFIC EQUIPMENT AND

OAT" AUTOMATION SYSTEM

BAY _]]]

POWER REGULATOR

AND BATTERY

+ PITCH

8

PROPULSION

SUBSYSTEM

THERMAL

SHIELO

BAY

-ROOT SUN

SENSOR (PRIMARY)

SUN GATE DETECTOR

SEPARATION- BAY ]_ISQUARE - R OOT

INITIATED TIMER _ H G / SUN SENSOR

CANOPUS SENSOR ._ iSECONDARY)

BAY _rrff BAY _Zl SEPARAT,ONL'NEARPYROTECHNICS /'_/-/_( _" _ , _ POTENTIOMETER

ARMING _ _'///'/'\ \ / ' _ _ PLUNGER

SW,TCH \ \ \

>'- -X

•x _ g

B E

INELIGHT

DISCONNECT

BAY TT _ _ _ //_/ . /_.' Ioi , /J _ LOWER THERMAL

_19 deq (S;OW_ED _ SHIELD

POSITION) BAY IllI

SQUARE-ROOT W_ _ COSMIC RAYSUN SENSOR + YA J TELESCOPE

(SECONDARY)

(C) TOP AND BOTTOM VIEWS OF SPACECRAFT

FIGurE 3-4.--Concluded.




Four rectangular solar panels were attached onto the primary structure at the

top of the octagon. These panels, providing an area of approximately 6.51 sq m

(70 sq It), were stowed in a nearly vertical position above the spacecraft during

the boost phase and were deployed after spacecraft separation from the Agena D

second stage.

To maintain attitude control, a dual nitrogen-gas system was devised which

consisted of 2 gas supplies and regulators supported in the internal cavity of the

octagon, 12 cold-gas jets mounted on the tips of the solar panels, and necessary

plumbing. A solar-pressure-vane auxiliary attitude-control system was mounted

on the tips of the solar panels. The installation on each panel consisted of an elec-

tromechanical actuator, a thermomechanical actuator, and a solar pressure vane

of 0.65 sq m (7 sq ft) of reflective surface area. The vanes were stowed along the

backs of the solar panels and unfolded to a nominal position upon deployment of

the solar panels after spacecraft separation. The total span of the spacecraft with

solar panels deployed and solar pressure vanes extended was 6.79 m (22.24 ft).

Sun sensors were mounted on both the top and bottom surfaces of the octagon

so that the Sun would be in the field of view of at least one set of Sun sensors

regardless of the spacecraft's angular orientation. The Canopus star sensor,

mounted on the bottom of the octagon, was provided with a clear field of view

between two solar panels. An Earth detector was mounted on the side of a Sun

sensor pedestal. The attitude-control system was linked by logic circuitry to the

Canopus and Sun sensors and to the three gyroscopes on the spacecraft.The high-gain antenna, a 116.8- by 53.3-cm (46- by 21-in.) ellipse in plan-

form (as seen from above) and a parabola in cross section, was attached to the

spacecraft by a conical support structure (superstructure) mounted on top of the

octagon. The antenna was supported in a fixed position such that it would be

pointed in the direction of Earth at planetary encounter. The low-gain omnidi-

rectional antenna was mounted on the end of a cylindrical tube approximately

10.2 cm (4 in.) in diameter and 223.5 cm (88 in.) long. This tube, which acted as

a waveguide and support structure for the antenna and three cruise-science experi-

ments, was attached at its base to the top of the octagon.

Flexible thermal shield blankets were attached to the superstructure on top

and the central cavity on the bottom. Louvers were provided on six of the bays, and

rigid aluminum shields covered the sides of the octagon structure not housing louvers.

The cruise-science experiments were mounted about the spacecraft, and the

television instrument--a planetary-science experiment--was mounted on a scan

platform at the bottom center of the spacecraft. These experiments were located to

$7



MARINER-MARS1964

m

¢

satisfy their field-of-view requirements without significantly compromising other

spacecraft system operations. The magnetometer was positioned near the top of

the low-gain omnidirectional antenna waveguide to separate it from the main

mass of the spacecraft while keeping it far enough from the antenna top to prevent

interference with the antenna pattern. The ionization chamber experiment was

located on the waveguide to provide a maximum field of view (since it was spheri-

cally sensitive) without significantly affecting the high-gain antenna pattern or the

magnetometer measurements. The cosmic dust detector was mounted on the

superstructure. The remaining cruise-science experiments were positioned to

satisfy their individual requirements. The television instrument was mounted on

the scan platform at such an angle as to optimize visual lighting conditions at the

planet. The platform was driven by a scan actuator mounted on the top of the

octagon through the tube running along the centerof the spacecraft to the platform.

Subsystem Functional Description

Structureand mechanisms

A spacecraft structure integrates the spacecraft subsystems into a rigidly sup-

ported functional spacecraft. The Mariner-Mars 1964 spacecraft structure was

composed of the basic octagon, the superstructure, and the scan platform. The

basic octagon housed the electronic equipment, cabling, midcourse propulsion

system, and attitude-control gas supplies and regulators. Several of the science

experiments, the Sun and Canopus star sensors, the superstructure, the side and

lower thermal shields, and all other exterior protrusions were mounted to the

octagon. The superstructure supported the high-gain antenna, the cosmic dust

detector experiment, the solar panels through their boost dampers, one low-gain

antenna damper, and the upper thermal shield. The superstructure is shown

attached to the basic octagon in figure 3-5. The scan platform provided the

mounting surface for the television instrument and sensors.

Spacecraft mechanisms, or those items required to produce or retard move-

ment of components during the missions, included the following:

1. Boost dampers, to reduce the vibration inputs to the solar panels and the

low-gain omnidirectional antenna during the boost phase and statically position

these items.

2. Pyrotechnics arming switch, to initiate commands on the spacecraft and

apply power to (arm) the pyrotechnics subsystem at spacecraft separation from the

Agena D second stage; and separation-initiated timer, to serve as a backup at

about 40 sec after separation.

58



8

.'SPACE VEHIClE SYSTEM DESIGN AND TESTING OPERATIONS*.

FIGURE-5.--Superstructure attached to basic octagon.

3. Solar panel deployment springs (one for each panel), to deploy the solar

panels to the cruise position.

4. Cruise dampers, to place the solar panels in the deployed position and

dampen solar panel excursions during propulsion maneuvers.

5. Scan actuator, to rotate the scan platform and television instrument to

search for and point to the surface of Mars.

6. Scan inhibit switch, to inhibit power to the scan actuator until the pin-

puller, which torsionally restrained the scan platform, was fired and to give a

telemetry indication that the pinpuller had released the platform.7. Science cover, to protect the television instrument and sensors from sun-

light and cosmic dust.

Radio

The radio, data encoder, video storage, and command subsystems made up

the Mariner-Mars 1964 spacecraft telecommunications system. In conjunction

59



MARINER-MARS 1964

with the Deep Space Instrumentation Facility (DSIF), these subsystems served to:

(1) track the angular position, radial velocity, and range of the spacecraft; (2)

provide engineering and scientific telemetry data from the spacecraft; and (3) use

ground commands to control spacecraft operation. The radio subsystem consisted

of a phase-coherent S-band transponder using a transmitter configuration de-

signed for maximum reliability of spacecraft-transmitted signals by the inde-

pendent use of redundant radiofrequency (RF) exciters and power amplifiers.

Selection of the exciters and amplifiers was by spacecraft logic, with ground com-

mand backup.

The transponder receiver and its power supply were mounted in bay V of the

basic octagon. The transponder transmitting portion, including a Mariner II-type

triode cavity amplifier and a longer life traveling-wave-tube power amplifier,

their associated power supplies, two band rejection filters, one four-port circulator

switch, one five-port circulator switch, two directional couplers, and control unit

switching logic, was mounted in bay VI. The modules were interconnected by

shielded two-conductor cable and RF coaxial cable. All circuitry, except the triode

cavity amplifier and the traveling-wave-tube power amplifier, was solid state. At

about 2300 MHz, the radio put out about 10 watts of power.

A functional block diagram of the radio subsystem is given in figure 3-6.

Operation was as follows:

1. A modulated or unmodulated RF signal transmitted to the spacecraft from

DSIF was received.2. The frequency and phase of the received signal was coherently translated

by a fixed ratio.

3. The received signal was demodulated by the automatic phase control

receiver, which tracked the carrier modulated signal in a phase-locked loop, and a

composite command signal was sent to the spacecraft command subsystem.

4. The range code, if transmitted to the spacecraft from DSIF, was de-

modulated.

5. The transmitter signal was modulated continuously with a composite

telemetry signal and the demodulated ranging signal if ranging were turned on.

6. The modulated RF signal was transmitted to the deep space stations of

DSIF.

Three transmitting and receiving antenna modes were available to provide

the required coverage during all phases of the missions; they are: (1) transmit low

gain, receive low gain; (2) transmit high gain, receive high gain; and (3) transmit

high gain, receive low gain. The low-gain omnidirectional broadcast antenna was

6{1




CS CIRCULATOR SWITCH

P/S POWER SUPPLY

AGC AUTOMATIC G AI N C ON TR OL

IF INTERMEDIATE FREQUENCY

VCO VOLTAGE-CONTROLLED OSCILLATOR

CCaS CENTRAL COMPUTER AND SEQUENCER

TELEMETRY

TELEMETRY

HI_H

GAIN

BALd

MI

PREAN

_ICED

ER I

LIFIER

AGC

IN

AMPLIFIER

DETECTOR LIMITER

ON/OFF

A GC O UT PU T

AGC INHIBIT

I DET

FREQ

SE

;TOR

M©

MODULATION

2 OUTPUT

!NCY

TRANSFER

COMMAND

I

ISO-

INHIBIT INHIBIT

CS3 CS3

C_ __ CS4 CS4 CS5_ I L _SS

CC_S C

AGC ANTENNA POWER AMPLIFIER INPUTINHIBIT CONTROL CONTROL EXCITER CONTROL 2.4 k¢

IHATT,TU.+ + + _C',! + +NHIBIT DIRECT DIRECT DIRECT CCGS DI E CCGS DIRECT

COMMAND COMMANID COMMAND COMMAND INPUT COMMANO

FIGURE 3-6.--Radio subsystem functional block diagram.

AUXflOSCI_

RANGING

MODULATION

TELEMETRY

MODULATION

INPUT

+

l R/SECEIVER

+INPUT

Z.4 kC 50vrms

61



MARINER-MARS1964,m

the primary antenna for the Earth-to-spacecraft link. This antenna did not have to

be pointed directly at the Earth in order to receive a signal. Communications

during the first half of the mission were maintained through this antenna by using

the 10-kW DSIF transmitters. A 100-kW DSIF transmitter would have been

required to support the second half of the mission if the spacecraft receiver had not

been switched to the fixed high-gain antenna (a narrow-beam elliptical reflector

antenna). This antenna could be used only when pointing directly at the Earth or

during the latter half of the mission. The high- and low-gain antennas are shown

on the spacecraft in figure 3-7. The switchover between the antennas could be

effected by a central computer and sequencer command or by ground command.

Failure-mode switching was also available to switch the receiver to the low-gain

antenna to take advantage of its broader coverage should roll reference be lost.The direction or angular position of the spacecraft in its flight could be cal-

culated from the pointing angles of the narrow-beam ground antennas. The two-

way Doppler shift could be measured to provide a value for the spacecraft radial

velocity (range rate) with respect to the station. Since the spacecraft was receding

from the ground antenna, the frequency of the narrow bandwidth radio signal

received by the spacecraft was less than the ground-transmitted frequency (called

the Doppler effect). Similarly, the ground-received signal was lower in frequency

than the spacecraft-transmitted signal. The operating frequency could be con-

trolled by either an oscillator on the spacecraft or, for more precise determination,

a stable oscillator on the ground, which generated a signal to be transmitted to the

spacecraft, where it was multiplied by a known factor (240/221), amplified, and

then sent back to Earth.

A "turnaround" ranging system capable of measuring the Earth-to-space-

craft range to a distance of 1 million km (625 000 statute miles) was included. The

ground-transmitted signals could be modulated by a long binary wave train

known as the range code. This code could then be retransmitted as modulation on

the spacecraft signal. When received by the ground station, the range code could

be shifted in time relative to the original signal by the round-trip radio propaga-

tion time, and thus provide a measurement of spacecraft range. (This capabilitywas not, however, used during the missions.)

Spacecraft performance and scientific data, both digital and analog, were

transmitted in a continuous stream of telemetry information in the form of binary

digits (bits). Data transmission was in digital form at rates of either 8½ or 33½

bits/sec. Bit-rate selection was by command from the central computer and se-

quencer, with ground-command backup capability.

62



Y

8

h


FIGURE-7.-High- an d low-gain antennas mounted on spacecraft.

Data encoder

The data encoder, located in bay IV , accepted engineering data from 90analog channels, time-multiplexed (commutated) them into a predetermined

sequence, and converted these data into 7-bit binary words. It also accepted

digital data from the science, video storage, and command subsystems and time-

multiplexed these with the engineering data. The data encoder subsystem gener-

ated a cumulative count of specific spacecraft events, as well as a cyclic, binary,

pseudorandom code from which bit and word synchronization (sync) could be

63



MARINER-MARS1964

obtained. A composite telemetry signal, consisting of telemetry data biphase-

modulated on a square-wave subcarrier and linearly added to a synchronization

subcarrier, was generated and sent to the RF phase modulator. Two data rates--8½ and 33½ bits/sec--were provided to take advantage of the higher signal

strengths available during the early part of the mission to support a higher bit rate.

The lower rate was used after the Earth-to-spacecraft range became so large that

the ground-received signal level decreased to the point where excessive bit errors

were introduced at the higher bit rate.

A functional diagram of the data encoder subsystem is shown in figure 3-8,

and the telemetry commutation is illustrated in figure 3-9. Table 3-III gives the

telemetry channel assignments. The input system, which included a commutator,

programer, and signal-conditioning circuits, was that portion that received,

routed, and conditioned the input data. The commutator consisted of solid-state

switches driven by complementary flip-flop registers called sequencers. These,

together with the logic circuits, were designated the commutator programer. The

signal-conditioning circuits converted nonstandard inputs to standard levels or

formats required by the basic system.

The conditioned time-multiplexed analog data from the input system were

received by the basic system of the data encoder. This system also generated the

serial data train, the synchronization code, the two subcarriers, and all clock

frequencies for tile encoder. The primary elements of the basic system were the

rate selector, pseudonoise generator, analog-to-digital converter, biphase modu-lator-mixer, and power supply.

Four data modes were available for use during the missions:

1. Data mode 1, used during maneuvers. Only engineering data (420 bits/

data frame) were sampled. This mode could also be used to increase the engineer-

ing data sampling during cruise to aid in failure analysis if required.

2. Data mode 2, used during launch, initial acquisition, and cruise. Blocks of

engineering data 140 bits long were alternated with blocks of science data 280

bits long.

3. Data mode 3, used during Mars encounter while the television and the

planetary scan system were viewing the planet. Only science data (420 bits/data

frame) were sampled.

4. Data mode 4, used to play back the stored science data from the magnetic

tape of the tape recorder in the video storage subsystem. (After the transmission of

each complete picture, approximately 2 hours of real-time data mode 1 engineer-

ing data were transmitted.)

64




SPACECRAFT

INPUTS TO

COMMUTATOR

S WI TC H A NA LO G-TO -D iG IT AL

CONVERTER/PSEUDONOISE

SIGNAL TO

TRANSMITTER

SYNC TO

ANALOG-TO-

DIGITAL

CONVERTER

-- _,4- kc

REFERENCE __RATE

COMMAND

iNPUTS

VIDEO

OAS STORAGE

2f$ SYNCHRONIZATION SUBCARRIER FREQUENCY TIMER INPUT

DAS DATA AUTOMATION SYSTEM COMMAND COMMAND

LLA LOW-LEVEL AMPLIFIER BIT SYNC LOC K t NPU T

FIGURE 3-8.--Functional block diagram of data encoder.

1TO V IDE O

RECORDER

These modes were differentiated by the specific data format within the mode with-

out consideration of the bit rate. Selection of data modes was by ground command

or logic on the spacecraft, depending on the sequence of inflight events. The sub-

system accommodated 100 commutated measurements.

Video storage

The video storage subsystem was required to accept and record digital tele-

vision picture data from the data automation system at a fixed rate of 10 700

bits/sec and to store a minimum of 20 pictures of video data on magnetic tape.

These data were then reproduced, for coding by the data encoder, at a rate of 8 ½

bits/sec during the picture playback period. Because of the low data rate, it took

more than 8 hours to transmit one picture (containing approximately 250 000

65



MARINER-MARS 1964,t

DECK I0 RATE I HIGH IOENGWORDSAT ? BITS/WORD [DECK II R ATE I (HIGH) SCIENCE FORMATTED DATA I

. 02 03_104_ 05 06 07 08 091 liD [ I 2 113 114 IIS 116 t7 118 119 N DATA MODE 2 THERE ARE 280 BITS OF

lOOt DE 3KIDECKI | | _ I SCIENCE DATA (40 ENGINEERING

2,_XI 2!X[ [ | I I 122xl I ] I l [ I 1 I WORDSEQUIVALENT) II

I [DECK21RATE2 (MEDIUM) ] I IDECK22 RATE2 (MEDIUM) I

210 211 212 213 214 215 216 217 218 219 220 221 222 223 224 225 226 227 228 229

40xDECK41x DECKI3x t_

I I DECK 4C RATE 3 (LOW-LOW) I DECK 41 RATE 3 ( LOW-LOW)

i i

I I DECK 20. RATE 2 MEDIUM_ I IDECK 30. RATE 3 (LOW)I 2OOlll 201 I 202 ] 203 204 205 206 207 208 209 .._ 300 301 302 303 304 305 306 07 308 309 DN DATA NUMBER

RATE TiME BETWEEN SAMPLESiODE PHASE TELEMETRY DECK DATA MODE I DA TA M OD E 2 J

I MIDCOURSE ENGINEERING ONLY 53 I/3bps 8 I/3bp$ _3 I/3bps 8 I/3 bps

* DECK SYNC WORD ( 7 LOGIC ONES) LAUNCH ENGINEERING AND I HIGt 4.20 sec IEB =ec 12 6 sec 504 $ec** COMMAND DETECTOR MONITOR 2 ANO CRUISE REAL-TIME SCIENCE I21ME[ 420sec }_c (28rnin)l126;ec (21m,n)lSO4sec (84rain) /

DN <- 31 : IN LOCK REAL-TIME

DN>32=OUTOFLOC K 3 ENCOUNTER SCIENCE ONLY I ILO'l_ 420sec (7rain) }Osec(2Bf/_lri)/126)sec(21min)SO4Osec (14hr) I

POST- ENGINEERING AND NON- 131LOW 840_c,4minl PoOsec(56rnin)125_3sec42min)llO_:)80sec(2.Bhr) IOTE FOR CHANNEL ASSIGNMENTS, 4 ENCOUNTER REAL-TIME SCIENCE [ I LL_SEE TABLE 3-111

FIGURE 3-9.--Data encoder telemetry commutation.

Table 3-111.--Telemetry channel assignments

Channel

100

101

102

103

104

105

106

107

108

109

110

111

112

113

114

115

115

116

Time between samples, see

Data mode 1 Data mode 2Measurement

Synchronization ..............................

Deck 200 ....................................

Deck 210 ....................................

Receiver automatic gain control (coarse) .........

Command detector monitor ....................

Pitch gyro, fine Sun sensor .....................

Yaw gyro, fine Sun sensor ......................

Roll gyro, Earth position detector ...............

Canopus intensi ty . .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Power switching and logic output voltage ........

Deck 220 .....................................

Receiver static phase error .....................

Pitch position ................................

Yaw position .................................

Roll position, roll search .......................

Event counter 1: pyro amplifiers, gyros on, solar

panel 4A1 open ... .. .. .. .. .. .. .. .. .. .. .. .. ..

Event counter 2: CC&S events, solar panel 4A3

open ......................................

Event counter 3: pyro arm, pyro amplifiers, solar

panel 4A5 open, end-of-tape loop .............

33a/_ 81_

bits/sec bits/sec

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

33_

bits/sec

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

12.6

8¼bits/sec

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4

50.4




Table 3-111.--Telemetry channel assignments--Continued

Channel

116

117

118

119

200

201

202

203

204

2O5

206

207

208

210

211

212

213

214

215

216

217

218

219

220

221

222

223

224

225

226

227

228

229

30O

301

302

303

304

305

306

307

308

309

Measurement

Event counter 4: command events, Sun acquire,

solar panel 4A7 open, scan platform unlatched..

Thrust chamber pressure

Propulsion nitrogen tank pressure ................

Propellant tank pressure .......................

Synchronization ... .. .. .. .. .. .. .. .. .. .. .. .. .. .

Low deck posit ion .. .. ... .. ... .. ... .. ... .. ... .

Deck 500 ....................................

Dual booster regulator input current ............

Power switching and logic current to commutator

converter ..................................

Main booster regulator output current ...........

Battery voltage ... .. .. .. .. .. .. .. .. .. .. .. .. .. ..

2400-Hz inverter output voltage ................

Attitude-control X, -- Y gas pressure ............

Attitude-control --X, Y gas pressure ..............

Receiver local oscillator drive ...................

Decks 400 and 410 ............................

Decks 420 and 430 ............................

High-gain antenna drive .......................

Low-galn antenna drive ......................

Receiver automatic gain control (fine) ...........

Battery charge curren t .........................

Propellant temperature ........................Attitude-control X, -- Y gas temperature .........

Attitude-control -- X, Y gas temperature .' ........

CC&S timing of events ........................

Maneuver booster regulator output current .......

Solar panel 4A1 current .......................




Battery current drain ..........................

2400-Hz inverter output current ................

Oxidizer pressure ... .. .. .. .. .. .. .. .. .. .. .. .. ..

Exciter power output ..........................

Cathode, helix current ..........................

Excite r voltage 1 . ... .. ... .. ... .. ... .. ... .. ... ..

Excite r voltage 2 . ... .. ... .. ... .. ... .. ... .. ... .

Canopus sensor cone angle .....................

Recorder pressure . .. .. ... .. ... .. ... .. ... .. ... .

Solar vane position, -bX .......................

Solar vane position, -X .......................

Solar vane position, +Y .......................

Solar vane position, --Y .......................

CC&S 28-V monitor ..........................

Time between samples, sec

Data mode 1 Data mode 2

33a/_ 8_/_

bits/see bits/see

4.20 16.8

4.20 16.8

4.20 16.8

4.20 16.8

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

420 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.O

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.0 168.0

42.O 168.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

420.0 1680.0

331/'_ 81/_ "

bits/see bits/see

12.6 50.4

12.6 50.4

12.6 50.4

12.6 50.4

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 5O4.O

126.0 504.0

126.0 504.0

126.0 504.O

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 5O4.O

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

126.0 504.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

1260.0 5040.0

67



MARINER-MARS 1964

a

o

w--

Table 3-111.--Telemetry channel assignments--Concluded

Channel

400

401

,I02

403

404

405

4O6

407408

409

410

411

412

413

414

415

416

417

418

419

420

421

422

423

424

425

426

427

428

429

430

431

432

433

434435

436

437

438

439

Measurenlent

Time between samples, sec

Data mode 1 Data mode 2

Synchronization ... .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Bay I temperature .. ... .. ... .. ... .. ... .. ... .. .

Bay III temperature ..........................

Spare .......................................

Bay V temperature ............................

Bay VI tempecature ...........................

Bay VII louver position indication ..............

Power regulator assembly temperature ...........Propulsion nitrogen tank temperature ............

Solar panel 4A1 front temperature ..............

Canopus sensor temperature ....................

Scan actuator temperature .....................

Absorptivity standard temperature I .............

Absorptivity standard temperature 2 .............

Science cover and spacecraft identification ........

Standard cell current ..........................

Radiation-resistant cell current .................

Standard cell voltage ..........................

Television temperature ........................

Ionization chamber temperature ................

Temperature reference .........................

Bay II temperature ............................

Bay III louver position indication ...............

Bay IV temperature ...........................

Crystal oscillator temperature . .. ... .. ... .. ... .. .

Bay I louver position indication .................

Bay VII temperature ..........................

Spare .......................................

Batte ry temperature . .. ... .. ... .. ... .. ... .. ... .

Solar panel 4A5 front temperature ...............

Lower ring temperature above Canopus sensor ....

Upper ring temperature under Sun sensor ........



Upper thermal shield temperature ...............

Lower thermal shield temperature ........... ....

Recorder temperature .........................

Instrument simulator (replacement for ultraviolet

photometer) temperature ....................

Trapped radiation detector temperature .........

Magnetometer temperature . ... .. ... .. ... .. ... .

33_

bits/see

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0840.0

840.0

840.0

840.0

840.0

840,0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0

840.0840.0

840.0

840.0

840.0

840.0

8¼bits/sec

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.03360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.0

3360.03360.0

3360.0

3360.0

3360.0

3360.0

33_

bits/see

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.02520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.0

2520.02520.0

2520.0

2520.0

2520.0

2520.0

8IAbits/see

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 O80.0

10 080.010 080.0

10 08O.0

10 080.0

10 080.0

10 O80.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 O80.0

I0 080.0

10 080.0

10 080.O

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 080.0

10 O80.010 080.0

10 080.0

10 080.0

10 080.0

10 080.0

68




bits). The total bit storagc capacity of thc tape recorder had to be at least 5.24x

10 G bits, contained on two tracks of 0.64-cm (0.25-in.) tape 100 meters (330 ft)

long. Design bit error rate was less than 1 bit error in 105 bits, not including errors

accumulated while establishing lock or those due to track changes.

The tape recorder, located in bay V, had three operating modes:

1. Launch mode, in which the tape transport was operated at the recording

speed, although no actual recording took place. The record motor was started

shortly before launch and it continued running until after spacecraft separation.

The tape was then positioned automatically at the proper location in preparation

for planetary encounter.

2. Record mode (fig. 3-10), in which application of 2400-hertz power and

single-phase 400-hertz power prior to encounter energized all sections of therecorder system except those associated with the launch mode. The recorder was

then ready to record at least 20 picture sequences while making two complete

passes of the tape, changing tracks after the first pass and returning to the initial

track after the second pass. Response to record .......... J- then '-' "" '.2Ol].lllldl IU_ 3¢ig" a s 1i1111 kii L(2(.I.

During recording, a pulse was generated for each complete pass of the tape by the

end-of-tape circuit as the end-of-tape foil passed over sensor contacts. The data

6 _ OSE FREQUENCY- OSE RECORD

MAIN DOUBLED RECORD MOTOR MONITOR

24-kc POWER_ _OWER -6 _ MONITOR OSE OPERATIONAL S UP POR T EQ UI PM EN T

SUPPLY _+_0 ____ t H I DAS DATA AUTOMATION SYSTEM

RETuRN-TO- c- .............. _'_ @ _ _ F--------_

R ATA ' I /

_ RETORN-TO-ZEROTO _--"I'>--4 -j _ ' _ 1"FREQUENCY-DOUBLED CONVERTER [ STOP _ / RF((3_r)\ I I _ |

L IDATA AND SYNC COMBINER) I l NORMAL I I 2,-_;_b- Y-J .... _.) /

S:0LEP :SE o TART I 1 Joo-Hz OWER_ o LAUNC. , 'r S "---C___

UMRl/Ir A/ I _ I I / _ 330- it TAPE

I NT E G_ RA'TOR _ _ [ ! I v _ ENDLESS LOOP

MONITOR -- I L IE%OL ?R JDAS I

START RECORD _L_--_ I I"--.. 1 _ _ I DATA ENCODER

COMMAND _ - J+ )--f-_ _'-- J-- t) I TRACK _l---J_ i_'_ _ I ( PLAYBACK

N',B,TC "_ I _ _ _ _ E I COMMANDDAS

STOP RECORD >----- i _r'_ I r.._ "L-c_ I I _ OSE PLAYBACK

COMMAND / _- +)-'--_ _ P I_ I - MOTOR ON

T _'L_ I V L AT_-HING I I I

OSE I RELAY DATA ENCODER

TARTECORD ; END-OF-TAPECOMMAND _ q CHANGE I

I CIRCUIT ' OSE

O ;OPRECOROI I? iiii!2 - i _ _ END-OF-TAPE

M A "_ _ i '_ L _ ( UMBIUCAL

CO M NO i I_ I _ I TRACK STEP

COUNT TWO ',

sTeP C,RC TSI ENOOPTAREOSE TRACK FLIGHT COMMAND

INDICATOR TRACK STEP

FIGURE 3 10.--Functional block diagram of video-storage-subsystem record mode.

69



MARINER-MARS 1964

automation system was mechanized to terminate the encounter record sequences

upon receipt of the second end-of-tape signal or at the end of the 22d picture if an

end-of-tape pulse had not yet been received. For telemetry purposes, the end-of-

tape pulses registered on an event counter in the data encoder. Pictures were

recorded in two-sequence frames.

3. Reproduce mode (fig. 3-11), in which after encounter, the recorder system

was switched into the reproduce mode by a command from the data encoder.

When the correct playback tape speed and phase for accurate readout of the data

was achieved, the recorder system was "phase locked." This reproduce mode

could continue indefinitely after encounter. Complete reproduction of 20 picture

sequences required from 8 to 10 days. Provision was included for switching the

data encoder from data mode 4 (picture playback) to data mode 1 (engineering

data) during the time data were absent because of erasures on the tape made

between record sequences. As in the record mode, the end-of-tape pulse served to

change tracks and it registered on an event counter in the data encoder.

OSE PLAYBACK

+B AMPLIFIER MONITOR

POWER ÷ 20 _'

I FREQUENCY-l _ | _DATA ENCODER _DOUBLEO TO NRZ,_I._ I PEAK_ _ _l._ 1' _T-RA_

DATA OUTPUT - _ l I CONVERTER AND ! I DET ECT OR I _ I t SWaT

iSYNC SEPARATOR AMPLIFIER !AOUTPUT A,2

_1/-- SL AVED BiTSYNC

@ATA ENCODER@ODE 4/t

DI\

DATA ENCODER_MASTER BIT-

SYNC INPUT

L CE_

i

RECORD-

START-STOPMBILICAL RELAY_NTEGRATOR 4

MONITOR I

START

PHASE COMPARATOR

OSE OPERATIONAL SUPPORT EQUIPMENT

NRZ NONRETURN-TO-ZERO

VCO VOLTAGE-CONTROLLED OSCILLATOR

PREAMPLIFIERS f ....

--J 330-h TAPE

POWER END-OF-TAPE

- L- L_F""_ _ DATA ENCODER

T -_ - END-OF-TAPEI INHIBIT

L_ _ DATA ENCODER

: PLAYBACK

COMMAND

OSEPLAYBACKMOTOR ON

OSE

END-OF-TAPE

@_ UMBILICAL-- ( TRACK STEP

OSE TRACK FLIGHT COMMAND

INDICATOR TRACK STEP

FIOURE 3-11.--Functional block diagram of video-storage-subsystem reproduce mode.

7O




Comman_

The command subsystem, located in bay IV, was required to process andexecute any of 29 direct command (DC) words and 3 quantitative command (QC)

words sent by the DSIF ground stations. A description of the ground commands

is given in table 3-IV, and a functional block diagram of the subsystem is given in

figure 3-12. The groutld-transmitted signal was modulated by command data to

control various spacecraft functions. The spacecraft receiver demodulated the

composite received signal and routed the command subcarriers from the radio

subsystem to the command subsystem. The command subsystem detected the

command word information (a sequence of binary digits), decoded the informa-

tion content, and issued the necessary command signals to the appropriate space-

craft subsystems. The command word format is shown in figure 3-13.

The subsystem was composed of a command detector, a decoder, and a

transformer-rectifier (to convert the spacecraft 2400-hertz voltage into the ac and

dc voltages required by the command subsystem). A command word was com-

posed of 20 bits, the first 3 of which began the command decoder function. The

next 6 bits, the command address bits, identified which command word was sent.

The remaining bits were quantitative information bits. For a DC, a momentary

switch closure took place in the recipient subsystem or subsystems circuitry. For

a Q C, a sequence of binary digits representing roll, pitch, and yaw midcourse-

maneuver information was directed to the central computer and sequencer.

SIGNAL PHASING WITHIN THE COMMAND SUBSYSTEM

BIT I I

SYNC

I i

i

ONEBIT II

I

ZEROBIT

COMMAND WORD BITS_

I .......... IBIT SYNC _ ._IiCOMMAND E -- .....

FROM SPACECRAFT_=I_IDETECTOR I DELAYED SYNC

2f s SYNCHRONIZATION SUBCARRIER

FREQUENCY

_-_IDC WORD OUTPUTS TO

[ SPA CECR AFT SUB SYSTEMS

_"_QC WORD OUTPUT)

IP ALERT PULSES _ SPACECRAFTsuBsYSTEMC_S

BIT SYNC J -

_--= EVENT PULSES _ SPACECRAFT

BIT SYNC _ DATA ENCODERUBSYSTEMDETECTOR LOCK

SIGNAL

5 0 V rm s,

SQUARE WAVE,_2,4 kc

i _--- + 6 V dc

TRANSFORMERCOMMAND._1 w - 6 V dc

RECTIFIER _ + 28 V dc

UNIT _ 25 Vrm$, SQUARE WAVE, 2.4 kc

FIGURE 3--12.--Functional block diagram of command subsystem.

71



MARINER-MARS 1964,p

Table 3-1V.--Ground commands

Command Function Effectdesignation

DC-1 Command data mode 1 ...............

DC-2

DC 3

DC-4

DC-5

DC-6

DC-7

DC-8

DC-9

DC-IO

I)C-11

DC-12

DC-13

Command data mode 2, turn on cruise

science

Command data mode 3 ..............

Command data mode 4 ..............

Switch data rates .....................

Switch analog-to-digital converter/pseu-

donoise generators

Switch power amplifiers ..............

Switch exciters . .. .. .. .. .. .. .. .. .. .. .

Switch ranging .. .. .. .. .. .. .. .. .. .. ..

Transmit high, receive low ............

Transmit high, receive high ...........

Transmit low, receive low .............

Inhibit maneuver command, inhibit pro-

pulsion command

Transfers data encoder to data mode 1 (engineering

words) as soon as transfer is acceptable to the

data encoder transfer logic

Transfers data encoder to data mode 2 (20 engineer-

ing words, 40 science words) as soon as transfer

is acceptable to the data encoder transfer logic.

Applies 2400-Hz power to cruise-science instru-

ments

Transfers data encoder to data mode 3 (science

words) as soon as transfer is acceptable to the

data encoder transfer logic

Transfers data encoder to data mode 4/data mode 1

(television picture data/engineering words) as

soon as transfer is acceptable to the data encode_

transfer logic. (If television picture data are avail-

able from video-storage subsystem, they are telem-

etered; if no such data are present, as between

pictures, then engineering data are telem-

etered.) Removes 2400-Hz power from cruise-

science ins truments

Transfers data encoder (operating at either 8x/_ or

331/_ bits/sec) from one bit rate to the other

Transfers data encoder (with 2 such items: A and B)

from one analog-to-digital converter/pseudonoise

generator to the other

Transfers radio (with 2 amplifiers: traveling-wave-

tube A, and cavity B) from one power amplifier

to the other

Transfers radio (with 2 exciters: A and B) from one

exciter to the other

Transfers spacecraft radio ranging receiver (with 2

positions: on and off) from one position to the

other

Causes radio circulator switches to be conditioned

so that spacecraft transmits on high-gain antenna

and receives on low-gain antenna


so that spacecraft transmits and receives on high-

gain antenna


so that spacecraft transmits and receives on low-

gain antenna

Removes attitude-control excitation power from

CC&S control lines so that attitude-control func-

tions controlled by CC&S are disabled. Prevents

pyrotechnics control circuitry from firing motor

start and stop squibs

72




Table 3-1V.--Ground commands--Continued


DC-14 Remove maneuver command inhibit...

I)C-15

DC-16

DC-17

LI(A-I _5

DC-19

DC-20

DC-21

DC-22

DC-23

DC-24

DC-25

Canopus gate inhibit override .........

Narrow-angle acquisition .............

Cycle Canopus cone angle ............

Gyros on: inertial control, positive roll..

Gyros off: normal control .............

Remove roll control ..................

Roll override: negative increment ......

Change tracks . .. ... .. ... .. ... .. ... ..

Arm second propulsion maneuver ......

Inhib it scan search ...................

Turn on planetary science, unlatch cover

Reverses state of all relays acted upon by DC-13.

Returns attitude-control and pyrotechnics sub-

systems to CC&S control

Causes Canopus sensor roll error signal to be applied

to roll gas jet electronics at all times whether or

not roll acquisition logic is satisfied. Prevents roll

search signal from being applied to roll channel

and prevents roll acquisition logic violations from

turning on the gyros

Initiates narrow-angle acquisition signal, thereby

conditioning data automation system logic to

begin television picture-taking sequence and to

transfer data encoder to data mode 3

Changes voltage on deflection plates of Canopus

sensor's image dissector, causing step change in

Canopus scnsor cone anglc

Turns on gyros (in inertial mode) and Canopus

sensor Sun shutter. Turns off Canopus sensor.

Turns on turn command generator. Conditions

attitude-control circuitry for commanded roll

turns. (Succeeding DC-18 commands cause

clockwise 2.25 ° roll turns)

Serves as reset for DC-15, DC-18, and DC-20

Turns off Canopus sensor. Turns on Canopus sensor

Sun shutter. Inhibits roll acquisition logic fromturning on gyros

Simulates Canopus acquisition logic violation.

Turns on gyros. Applies negative roll search

signal to roll gas jet electronics. Causes spacecraft

to begin counterclockwise roll search to acquire

a new target. (If preceded by a DC-18, causes a

2.25 ° counterclockwise roll turn by spacecraft)

Changes video-storage-subsystem tape tracks by

applying power to record head and gating output

of playback amplifiers

Sets relays in pyrotechnics subsystem such that

CC&S commands M-6 and M-7 are routed to

squibs allotted to second motor burn

Removes 400-Hz single-phase power from scan

platform drive motor

Causes 2400-Hz power to be applied to encounter

science loads, video storage subsystem, and cruise

science loads (if 2400-Hz power was off to cruise

science). At same time, applies 52 V dc from

booster regulator to 400-Hz single-phase inverter,

which in turn supplies power to scan system drive

motor and video-storage-subsystem record motor.

73



MARINER_MARS 1964a

Table 3-IV.--Ground commands--Concluded


DC-26

DC-27

DC-28

DC-29

QCI-1

QC 1-2

QC1-3

Turn off planetary science, cruise science,

and battery charger

Initiate midcourse maneuver ..........

Turn on battery charger, turn off video

storage subsystem

Arm first propulsion maneuver ........

Command required pitch turn duration

for maneuver

Command required roll turn duration formaneuver

Command required motor burn duration

for maneuver

Enables battery charger boost mode. Causes

pyrotechnics subsystem to energize solenoid that

releases scan platform science cover

Removes 2400-Hz power from all science loads

(allowing video-storage-subsystem 2400-Hz power

to remain on) and 52-V-de power from 400-Hz

single-phase inverter. Enables battery charger

boost mode

Starts maneuver sequence by issuing CC&S com-

mand M-1 (turn on gyros), applying power to

maneuver clock, and removing maneuver clamp

and flip-flop reset signal from CC&S maneuver

circuitry

Removes 2400-Hz power from video storage sub-

system. Enables charge mode of battery charger

Sets relays in pyrotechnics subsystem such that

CC&S commands M-6 and M-7 are routed to

squibs allotted to first motor burn

Sets pitch turn polarity and preloads CC&S pitch

shift register such that, at a 1-pulse/see counting

rate, the register will fill in required time interval

for attitude-control subsystem to pitch turn the

spacecraft the amount required for a given mid-

course maneuve

Sets roll turn polarity and preloads CC&S roll shiftregister such that, at a 1-pulse/sec counting rate,

the register will fill in required time interval for

attitude-control subsystem to roll turn the space-

craft the amount required for a given midcourse

maneuver

Preloads CC&S velocity shift register such that, at a

20-pulse/see counting rate, the register will fill in

required time interval for a midcourse motor burn

yielding required velocity change for a given mid-

course maneuver

Three telemetry signals were telemetered to Earth from the command sub-

system. These signals time shared the telemetry transmission channel and were

sampled at a rate determined by the telemetry mode. The signals were as follows:

1. Detector bit synchronization pulses, to be conditioned into information

concerning the detector voltage-controlled-oscillator frequency. From this infor-

74




DIRECT COMMAND FORMAT

COMMAND BIT NO. I 12 ]3

COMMAND BIT COM-

IDENTIFICATION MANDDECODER

START

COMMAND BIT VALUE 1 J1 Jo

, is1617B19oi"21'3i"1'S1'61'71'B1'9i2012'2212312412Si26- BITS2-_6AVENOS,ON.FICANCENDC'S

CO',_AND__-:,NQC'S,-,E,'ORMAR',FTHEC'.SADDRESS __ _ z_ COMMAND.REFER TO QUANTITATIVE COMMAND FORMAT

VARIABLE ZERO FOR DC'S; VARIABLE FOR QC'S

COMMAND BIT NO.

CC&S COMMAND BIT NO.

CC&S COMMAND BiT

iDENTIFICATION

QC PITCH TURN

COM-

MAND ROLL TURN

VALUES MOTOR BURN

NOIES:

QUANTITATIVE COMMAND FORMAT

26

18

1. COMMAND BITS 10 AND 11 ARE ADJUSTED TO ENSURE AGAINST SINGLE BIT ERRORS CAUSING AN INCORRECT COMMAND WORD OUTPUT.

2. COMMAND BITS 9 THROUGH 11 (CC&S COMMAND BITS 1 THROUGH 3) ARE NOT USED QUANTITATIVELY BY CC&S BUT ARE USED TO REMAIN

COMPATIBLE WITH PREVIOUSLY DESIGNED HARDWARE (MARINER VENUS 1962).

3 . COMMAND BIT 14 (CC&S COMMAND BIT 6) IS ADJUSTED IN QC'S TO GIVE AN ODD NUMBER OF ONE BITS IN COMMAND BITS 9 THROUGH

26 (CC&S BITS I THROUGH 18).

4. COMMAND BIT 26 (CC&5 COMMAND BIT 18) MUST BE A ONE TO P ROD UCE A CLOCKWISE (POSITIVE) SPACECRAFT ROTATION ABOUT THE

SPECIFIED SPACECRAFT AXIS, A ZERO IN THIS BIT POSITION WILL RESULT IN A COUNTERCLOCKWI SE (NEGATIVE) SPACECRAFT ROTATION

ABOUT THE SPECIFIED SPACECRAFT AXIS. _LARITY BIT FOR MO TOR BURN COMMAND IS ALWAYS ONE,

S. COMMAND BITS 15 THROUGH 25 (CC&S COMMAND BITS 7 THROUGH 17) ARE A PSEUDOBtNARY CODE REPRESENTATION OF THE TUR N OR

MOTOR BURN DURATI ON.

FIOURE 3-13.--Command word format.

mation, the proper ground-command signal frequency was ascertained so that a

minimum command lock acquisition time could be accomplished.

2. Detector lock signal, to be converted into one word of information regard-

ing the detector lock condition. The data indicated when the command subsystem

was capable of detecting and processing the command signal.

3. Command event pulses, approximately 10 sec after the command detector

recognized the start of a command word. The pulses indicated a normal decoding

timing function.

Attitude control

Attitude control is required to maximize solar panel electrical output by

orienting the panels perpendicular to the Sun's rays and to maintain communica-

tions through directional antennas. The Mariner-Mars 1964 attitude-control sub-

system (fig. 3-14), located in bay VII, established and maintained three-axis (pitch,

75



MARINER-MARS 1964,m

P

CANOPUS SENSOR CONE

ANGLE CONTROL

CCSS MT I

2_

s-_/"

4_

INHIBIT

OC-21 COMMAND ROLL ¢'---1

OVERRIDE ?

'+I

\ ,

_ S WI TC HI NG 1

---

;QUASI"

SUN

SENSC

i

LQUISI

SUN

SENSC

i POWE

E--IMAR

SUN

ZNSOF

YAW

PITCH --

ATTITUDE

CONTROL

dc POWER

RIMAF

SUN CONI

_ENSO _ dc P(POW,,FINTERMITTENT

LOAD I __?- j soy i_ I RO*ERI_

--J 240O Hz I

_.__-E

I

CANOPUS ACQUISITION

UNLESS OTHERWISE SPECIFIED,

ALL LATCHING RELAYS HAVE HOLD

CURRENT SUPPLIED THROUGH

SEPARATION CONNECTOR VIA

ATTITUDE-CONTROL TURN-ON

FIGURE 3-14.--Functional block diagram of

76




f %0,_LOV#-##R,,_,,_E##_ESET %;,iR#,_SREsE

CANOPUS GATE OVERRIDE RESET

DC-13 CC_S

_+, _c-,8ROLL×,SNERT,ALAC_UR VR,oNA_E#ER

' AE6 CCSS

" ' TURN ' I_'1_,

ANEUVER_ __]

i ; S C Cg S M-2

RO LL TU RNCOMMAND SCCP_S M-

IC

4 ONE-SHOT !

-- E-de 9 POLL

ROLL OVERRIDE COMMAND ; INCREMENT I

(-) ;NCREMENT L_

GYRO -

YAW U

PITCH TURN COMMAND

r

i3-PHASE, 400-Hz I

I I NV ERTER ]

GYRO SPIN

YAW i MOTOR

L AUTOPILOT]

GYRO SIGNAL I

GENERATOR

POWER

F 40 V _

GYRO _ GYRO INTERMITTENT,o,, ........PO W_ ___.._ MOTOR

_ _ ,,_ o:_ TELEMETRY

GYRO_CONT ROL DELAY - MI-- _' EVENT

26 V

a I CRUrSE I

, INVERTER ;

I I 2400 Hz

I _C_I . ,_E,v,%

SEPARATION SWITCH_GENA

ACITTITUDE CONTROL

POWER

1

I

I4i

I SC CBS M-3

I

SCCBS M-4

RQ...,>__

is;......

1

4F

S e.._.__

RCCe, S L-2

attitude-control subsystem.

77



MARINER-MARS 1964

+ YAW (CLOCKWISE)

+y '

Jt

+PITCH (CLOCKWISE)

÷xLOW-GAINANTENNA

SOLAR y /_'_ __

+ ROLL (CLOCKWISE)

+Z

FIOURE 3-15.--Spacecraft coordinate system.

yaw, and roll) stabilization of the spacecraft by using the Sun and the star Cano-

pus as references. (The spacecraft coordinate system is shown in figure 3-15. The

spacecraft may be compared with an airplane pointing in the direction the solar

panels faced--that is, toward the Sun--although the spacecraft usually traveled

at right angles to this direction. Yawing moved the nose to the right or left, pitch-

ing moved the nose up or down, and rolling spun the craft around.) Two-axis

(pitch and yaw) Sun stabilization alined the spacecraft Z-axis with the space-

craft-Sun line, keeping the sensitive surface of the solar panels facing the Sun. Roll

stabilization about the Z-axis with the star Canopus as the reference insured that

the high-gain antenna beam included Earth during the latter portions of the

mission. Whenever the spacecraft started to yaw or pitch, the cruise- and acquisi-

tion-phase Sun sensors produced error signals proportional to the angular position

displacement about the yaw and pitch axes. The acquisition Sun sensors facing

away from the cruise-phase Sun direction supplemented the cruise Sun sensors so

78




+26 V

CRUISE 1 ACQUISITION

SENSORS

NULLAxIS "16 V_ •

T SHADOW

-16V

-26VSENSOR

FIGURE 3-16.--Sun sensors.

OUTPUT TOSWITCHINGAMPLIFIERS

that the Sun would be in the field of view of at least one set of Sun sensors regard-

less of the spacecraft's angular orientation. Each Sun sensor consisted of a photo-

resistor mounted beneath a shadow mask (fig. 3-16). A Sun gate kept the power to

the acquisition Sun sensors on during acquisition and thus could be used in

determining when acquisition occurred.

Similarly, the Canopus sensor (fig. 3-17) produced, when a star was in its

field of view (11 ° in cone and 4 ° in clock), a error signals proportional to the

angular position displacement of the star from the spacecraft roll axis. An elec-tronic logic was set to respond to any object more than one-eighth as bright as

Canopus. Including Canopus, there were seven such objects visible to the Canopus

sensor as the spacecraft swung around in its search mode. A Canopus gate was

used to switch the roll-axis control system to the acquisition mode.

It was anticipated that star identification would be a major problem since the

only information from the Canopus sensor, other than the error signal, was a

brightness measurement and the cone angle of an object within ±5 °. Therefore, a

map-matching technique was developed to identify objects seen during the roll

search mode. Fundamental to the plan was an a priori telemetry-type map of the

expected sensor brightness output as a function of clock angle. A mathematical

model of the Canopus sensor and the sky, including the Milky Way, was developed

8Using the spacecraft center as the vertex (intersection) of the lines forming the angles, the position of an ob-

ject is described by its cone angle from the spacecraft-Sun llne and its clock angle from the spacecraft-Canopus line,

as projected on a plane at right angles to the Sun's direction (i.e., on the plane of the solar panels). These angles

are defined for the case of the Earth in fig. 3-18.

79



*c

MARINER-MARS 1964 .

FIGURE-1 7.-Canopus sensor.

so that, with available trajectory information, an IBM 7094 computer could be

programed to print, before launch, a map of the expected telemetry output of the

brightness channel during roll search. T h i s then could be matched with an actual

telemetry map to identify observed objects.

Initially, the actual telemetry map produced by a complex computer pro-

gram after receipt of the telemetry was statistically correlated with the a priori

map and other data available from the spacecraft. Whenever an object was ac-

quired after a roll search, the probability that each acquirable object had been

acquired was calculated. If the object were not Canopus, another roll search was

instituted and another computer run made until Canopus could be identified as

the object acquired. This procedure was necessary because of the lack of knowl-edge concerning the absolute calibration of the Canopus sensor for all stars and

the integrated background of stars in the field seen by the sensor. Until the sensor

could be calibrated in flight, the various uncertainties made it essential that the

best possible analysis techniques be prepared before launch.

A second technique for star identification was much simpler, but required

actual flight experience to prove reliability. With this technique, a continuous

80




Z [TO SUN_

E (TO EARTH)

SPACECRAFT

C (TO CANOPUS]

FmvaE 3 18.--Graphic definition of clock and cone angles. _ =cone

angle; /3= clock angle.

strip chart recorder was used to plot, in real time, the star sensor brightness telem-

etry. The a priori map was transcribed to a transparent overlay to the same scaleas the real-time telemetry plot for instantaneous comparison during a roll search.

This technique became the primary technique because of its accuracy and speed.

A typical plot--that for the first roll search of Mariner IV--is shown in figure

3-19.

The Mariner-Mars 1964 spacecraft were the first to use a star as a reference

object; earlier spacecraft such as Mariner II had sighted on the Earth. However,

81



MARINER-MARS 1964

..... PREDICTEO STAR MAP

ACTUAL TELEMETERED STAR MAP

EARTH

CANOPUS I

KOCHAB

PHECDAMERAK

REGULUS

CANOPUS

ALFARD /_

GAMMA \VELORUM \

AND S-302

I I I I I I I I40 80 120 160 200 240 260 320 360

CANOPUS SENSOR CLOCK ANGLE, deg

FIGURE 3-19.---Typical plot used in star identification procedure.

during the Mariner-Mars 1964 missions, Earth would transit across the face of the

Sun and through much of the flight it would appear as a relatively dim crescent.

Therefore, a bright reference source, such as Canopus, at a wide angle away from

the Sun was necessary. (The need for a second reference can be realized by

imagining a weight suspended from a single long cord. This weight would spin

unless a second cord, approximately at right angles, was attached, in which case

the weight would stabilize. Thus, Canopus simply served as the stabilizing cord of

the spacecraft.) An Earth detector (fig. 3-20), consisting of a photocell in series

with a fixed resistor, was used to verify roll stabilization by sensing reflected sun-

light from Earth after acquisition of Canopus.

The spacecraft was maintained at certain angular positions or rotated at

prescribed angular rates through the application of appropriate torques obtained

by the expulsion of gas through pairs of jet nozzles. These nozzles were located so

as to produce couples 4 about each of the principal axes. The gas jets were opened

4A couple is a system of two forces, equal in magnitude, parallel in direction, but opposite in sense.

82




FIGURE--20.-Earth detector assembly.

and closed by solenoid valves operated by switching amplifiers. These amplifiers

responded to position error signals from the Sun and Canopus senscrs. The gas

supply was contained in two completely independent nitrogen-gas systems. Oscil-

lations could be damped out by mixing a rate error signal from a rate gyroscope

(gyro) with the position error signal; however, the gyros were not operating during

most of the flight. Three single-degree-of-freedom gyros provided the rate error

signals when they were operating (during initial acquisition and for inertial con-

trol during the midcourse maneuver). When used during acquisition, all gyros had

an output to the switching amplifiers that was proportional to the sensed space-

craft angular rate.

For the long cruise period, a “derived rate” switching amplifier system was

devised which enhanced reliability by using, in place of the gyros, a passive stabi-

lization technique. Fundamentally, this technique depended upon a lag feedback

around the switching amplifier with a long time constant of about 100 sec. When-

ever there was a switching amplifier output to activate the gas jets, the long time

constant produced a feedback signal to the input. While the gas jets were open, the

spacecraft was accelerating, and the feedback appeared the same as a rate feed-back obtained by gyro integration of acceleration (thus, the term “derived rate”) ,

provided that the gas jets were open for a much shorter period than the derived-

rate time constant and the initial spacecraft rate was nearly zero. This system was

capable of controlling any cruise-mode disturbances.

During the midcourse-maneuver mode, the attitude-control subsystem was

required to orient the spacecraft so that the thrust vector of the rocket motor on

83



MARINER-MARS 7964 .

the spacecraft was alined to a predetermined direction in space. This orientation

was maintained during the motor-burn period, and the cruise orientation was then

reestablished. For inertial control used during the midcourse-maneuver sequence,

gyros provided error signals proportional to angular rate. This mode was con-

trolled by the central computer and sequencer. The rate error signal from a gyro

could be fed to a switching amplifier to maintain an angular rate within the rate

deadband orientation. This error signal could be integrated to give position error.

The midcourse autopilot was used to control the attitude of the spacecraft during

the motor-burn period. Control was accomplished by continuous adjustment of

the angular positions of four je t vanes mounted in the midcourse-motor nozzle.

Since the motor was not mounted along any of the three spacecraft axes, motion

about all three axes (pitch, yaw, and roll) was possible during midcourse motorburn. The motion of each je t vane was controlled by a mixture of the signals of all

three gyros. Each jet vane had its own control system consisting of an autopilot

amplifier, a jet vane actuator, and a feedback loop. Power was switched on only

during the midcourse-maneuver sequence. The three gyro signals and the feed-

back signal were combined in different proportions at the input to each amplifier.

The je t vanes were then adjusted so that the motor thrust vector passed through

the spacecraft center of gravity, nullifying the gyro error signals.

-SOLAR PRESSURE

VANE

-BACK SIDE OF

SOLAR PANEL

FIGURE- 21 .--Solar pressure vane and hardware mounted

back s ide of solar panel.

to

84





MARINER-MARS 1964

Table 3-V.--Central computer and sequencer (CC&S) commands

Command designation

L_l ...... . .. . .. .. . .. . .. . .. .. . .. . .. . .

L-2 ...... . .. . .. .. . .. . .. . .. .. . .. . .. . .

L-3 ...... . .. . .. .. . .. . .. . .. .. . .. . .. . .

M-1 (set) ............................

M-1 (reset) ..........................

M-2 (set) .......................... i

M-2 (reset) ..........................

M-3 (set) ...........................

M-3 (reset) ..........................

M-4 (set) ...........................

M-4 (reset) ..........................

M-5 (set) ...........................

M-5 (reset) ..........................

M-6 (pulsed) ..... .. .. .. .. .. .. .. .. .. .

M-7 (pulsed) ..... .. .. .. .. .. .. .. .. .. .

MT-1 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-2 (set) ... .. .. .. .. .. .. .. .. .. .. .. ..

MT-3 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-4 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-5 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-6 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-7 (set) ...........................

MT-8 (set) ... .. .. .. .. .. .. .. .. .. .. .. .

MT-9 (set) ...........................

CY-1 (pulsed) ..... .. .. .. .. .. .. .. .. .. .

Function

Deploy solar panels, unlatch scan platform

Turn on attitude-control subsystem

Energize Canopus sensor, turn on solar pressure vanes

Gyros on, data encoder to data mode 1

Gyros off, data encoder to data mode 2

Spacecraft on inertial control

Spacecraft off inertial control (reacquire Sun and Canopus)

Turn polarity (set if positive)

Turn polarity reset

Pitch-turn start

Pitch-turn stopRoll-turn start

Roll-turn stop

Ignite midcourse motorTurn off midcourse motor

First Canopus sensor cone angle update

Second Canopus sensor cone angle update

Third Canopus sensor cone angle update

Fourth Canopus sensor cone angle update

Switch transmitter to high-gain antenna

Switch data encoder bit rate to 81/_ bits/sec

Encounter science on (all science on)

Encounter science off (all science off)

Cruise science off, start data playback

Backup switching for radio subsystem (pulse occurring every 66_ hr)

outward from the center of the spacecraft; its field of view (11 °) covered an area

in the shape of a narrow cone. As the spacecraft moved about the Sun, the angle

between the Sun-spacecraft line and the Canopus-spacecraft line changed slowly.

The total change during the mission was approximately 28 ° . Therefore, the angle

of the Canopus sensor had to be altered four times during the mission to keep

Canopus in view. Each CC&S command to update the angle changed the cone

angle one increment of 4.6 ° .

PoweT

The power subsystem was designed to generate standard voltages for distri-bution to spacecraft power users and to turn on and off various spacecraft loads.

A functional diagram of the subsystem is given in figure 3-23. The components

of the power subsystem were as follows: (1) four photovoltaic solar panels with a

combined active area of 6.51 sq m (70 sq ft); (2) a 1200-W-hr, silver-zinc, 18-cell

battery; (3) dc voltage regulating devices; (4) 2400- and 400-hertz inverters; and

86



..L SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

0 3 6I I 1INCHES 91

\

FIGURE-22.-Assembled CC&S subsysteiri.

(5) battery charging, load switching, and frequency control devices. The solar

panels were mounted at the top of the spacecraft’s basic octagon, and the re-

mainder of the subsystem was contained in bays I and VII I.

Power during the launch-to-Sun-acquisition phase and during maneuvers

was provided by the battery. In order to insure maximum reliability, the sub-

system was designed to require no battery power after initial Sun acquisition,

except during maneuvers; however, the battery was maintained in a state of full

charge as a backup source of power if needed. Battery capacity was such that both

the launch and maneuver phases could be completed without battery recharging.

After Sun acquisition, the battery could be recharged by using the flight battery

a7



MARINER-MARS 1964R

(

>.

0

'-O

o

-a

.o

_4Oq

o

88




charger. After the battery was fuiiy charged, the charger could be turned off by

direct radio command to prolong battery life. If the battery should be needed

again as a result of losing Sun acquisition, a radio command could reapply the

charger.

When the spacecraft became Sun oriented, the 7056 photovoltaic cells on

each of the four solar panels converted sunlight into approximately 700 watts of

raw electrical power, which, in turn, was converted into various forms to operate

the spacecraft and recharge the battery. (At Mars distance from the Sun, the

battery still generated 300 watts, leaving an adequate safety margin in the event

of solar cell damage in the space environment.) Since launch trajectories allowed

the spacecraft to spend some time in the Earth's shadow, and since there is a

large solar panel power capability near Earth, it was necessary to be able to limit

the output voltage of each section of the solar panels immediately after the space-

craft left the Earth's shadow to 50 volts by incorporating six 50-watt zener diodes.

For increased reliability, the voltage regulating elements of the subsystem

consisted of two booster regulators: the main booster regulator, which would

normally be on throughout the entire flight, supplying power to all spacecraft

loads except the communications converter (which accepted unregulated power

directly from the battery or solar panels); and the maneuver booster regulator,

which would be used to power a large portion of the attitude-control subsystem

(with turn-on controlled by that subsystem) and which would be on during the

launch and midcourse-maneuver phases.

Under normal conditions, the main 2400-hertz inverter received dc power

from the main regulator. An identical inverter received dc power from the maneu-

ver regulator and supplied 100-volt peak-to-peak, 2400-hertz voltage to the

attitude-control subsystem. Also operating from the maneuver regulator was a

28-volt root-mean-square, 400-hertz, three-phase inverter that delivered step

square-wave power to the gyro spin motors. A 400-hertz, single-phase, square-

wave inverter supplying nominal outputs of 56 and 65 volts to the science scan

platform and video storage subsystem, respectively, operated from the main reg-

ulator. This inverter was turned off except during encounter.Some spacecraft subsystems received power whenever the power subsystem

was operating; others were turned on and off during various parts of the mission

by logic on the spacecraft or by direct radio command. The actual switching of

these loads was accomplished by the power subsystem in the power distribution

assembly. This unit accepted commands from other spacecraft subsystems and

translated the commands into relay closures.

89



MARINER-MARS 1964

Pyrotechnics

By definition, a pyrotechnic is a mixture of oxidant and reductant designed toperform some nonpropulsive function. The pyrotechnics subsystem on the

Mariner-Mars 1964 spacecraft supplied power and performed switching for the

following operations: pin retraction (by pinpullers) of eight solar panel latches and

one scan platform latch, first and second starts and stops of the midcourse motor,

and solenoid release of the science cover. The pyrotechnics subsystem accepted

commands from the appropriate sources and provided the energy necessary to fire

the proper squibs (small, electrically fired explosive charges) and to activate the

solenoid. The subsystem contained a total of 29 such squibs. A pyrotechnics firing

unit is shown in figure 3-24.

The subsystem was not energized (armed) prior to spacecraft separation so

that an inadvertent or spurious command prior to separation could not cause a

premature squib firing. At separation, the subsystem was armed by the pyro-

technics arming switch (fig. 3-25), which also supplied the arming event indica-

tion to the data encoder, an Agena D isolation-amplifier-off function, prelaunch

monitoring of the disarmed state of the switch, and turn-on of the attitude-control

FIGURE-24.-Pyrotechnics firing unit.

90



*.L SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

FIGURE-25.-Pyrotechnics arming

switch.

FIGURE-26.--Separation-initiated timer.

subsystem. This pyrotechnics subsystem arming event a t separation was backed up

by the separation-initiated timer (fig. 3-26), and the attitude-control subsystem

turn-on was backed up by the CC&S. The separation-initiated timer provided

two time-delayed switching functions sequenced from spacecraft separation : arm-

ing of the pyrotechnics subsystem and deployment of the solar panels.

Propulsion

The Mariner-Mars 1964 propulsion subsystem (fig. 3-27) was functionally a

monopropellant-hydrazine, regulated-gas-pressure-fed, constant-thrust rocket for

use in trajectory-correction (midcourse) maneuvers. Principal components were :

(1) a high-pressure gas reservoir (nitrogen tank); (2 ) a pneumatic pressure regu-

lator for reducing inlet pressure to a constant value; (3) a propellant tank and

propellant bladder containing the hydrazine; (4) n ignition reservoir; (5 ) explo-

sive valves for start and stop functions; and (6) a rocket motor (midcourse motor).The midcourse-motor installation is shown in figure 3-28. A catalyst was included

to accelerate hydrazine decomposition. The nominal vacuum thrust capability

was 222 newtons (50 lb). The tankage was sized for a maximum velocity increment

of 81 m/sec (approximately 267 ft/sec) to a 272-kg (600-lb) spacecraft, but

velocity increments as small as 0.2 m/sec (approximately 0.67 ft/sec) were

possible.

91



MARINER-MARS 1964

^

COMPONENTS

No DESCRIPTION

I ROCKET ENGINE

2 N 2 FILL VALVE, IGNITION CARTRIDGE

3 N 2 RESERVOIR, IGNITION CARTRIDGE

4 OXIDIZER VALVE, START I

5 OXIDIZER VALVE, START 2

6 OXIDIZER RESERVOIR, IGNITION CARTRIDGE

7 CHECK VALVE

8 NITROGEN TANK

9 OXIDIZER FILL VALVE I, IGNITION CARTRIDGE

I0 OXIDIZER FILL VALVE 2, IGNITION CARTRIDGE

II PROPELLANT VALVE, START 1

12 PROPELLANT VALVE, START 2

13 PROPELLANT VALVE, SHUTOFF I

14. PROPELLANT VALVE, SHUTOFF 2

15 FILL VALVE, PROPELLANT

16 PROPELLANT BLADDER

17 PROPELLANT TANK

IB PREPRESSURIZATION VALVE, NITROGEN

19 FILL VALVE, NITROGEN TANK

20 PRESSURE REGULATOR, NITROGEN21 NITROGEN FILTER

22 NITROGEN VALVE, START I

23 NITROGEN VALVE, START 2

24 NITROGEN VALVE, SHUTOFF I

25 NITROGEN VALVE, SHUTOFF 2

_] TWO-WAY VALVE, EXPLOSIVELYOPERATED, NORMALLY OPEN

TWO-WAY VALVE, EXPLOSIVELYPERATED, NORMALLY CLOSED

ANGLE VALVE, MANUALLYPERATED

CHECK VALVE

FILTER

PRESET REGULATOR

COMPONENT NUMBERS

INSTRUMENTATION NUMBERS

INSTRUMENTATION

PRESSURE TRANSDUCERS

(_NITROGEN TANK

(_PROPELLANT TANK

(_)THRUST CHAMBER

IGNI TI ON CA RT RID GEN2 RESERVOIR

TEMPERATURE TRANSDUCERS

PROPELLANT

NITROGEN TANK

NOTE :

EVENT REGISTER DATA (BLIP DEA)

ON EXPLOSIVE VALVE START AND

SHUTOFF CURRENT

FIOURE 3-27.--Schematic drawing of propulsion subsystem.

92




NOMINAL UNDEFLECTED THRUST

AX iS ( MOTOR CENTERLINE)/

/

MOTOR TANK THRUST PLATE _-'-'_

MIDCCOURSE MOUNTING BLOCKJ_''/_l_

SUBSTRUCTURE /

SPACECRAFT

MOUNTING BRACKET

-- SUBSTRUCTURE

l-_ _-P,VOTII POINT

FIGURE3-28.--Midcourse-motor installation.

The subsystem, located in bay II, was capable of two starts, and thus two

midcourse maneuvers could be performed if necessary. The ignition and thrust

termination signals were generated by the pyrotechnics subsystem upon command

from the CC&S. Ignition and thrust termination were controlled by explosive

valves, which fired simultaneously to initiate nitrogen, propellant, and oxidizer

flow and to terminate nitrogen and propellant flow. "Ganged" valves in parallel

were used to meet the two-start (dual-burn) requirement. Jet vanes capable of

deflecting the midcourse-motor jet stream provided the thrust vector control

through ±5°i The thrust produced by the system could be calibrated so accurately

that the change of velocity resulting from the maneuver could be metered by the

burning time of the motor alone.

Temperature control

The four major variables affecting the temperature of spacecraft components

are incident solar radiation, electrical power expenditure, thermal transfer be-

tween components, and thermal radiation from the spacecraft into space. The

Mariner-Mars 1964 spacecraft temperature-control subsystem was comprised of

all devices and treatments employed to maintain temperatures within specified

bounds.

93



MARINER-MARS 1964

Th e principal active temperature-controi devices were six variable-emittance

(varying amounts of heat radiated per unit area) louver assemblies mounted on

six of the electronic assemblies (fig. 3-29). Passive devices included removable

multilayer blankets (fig. 3-30) shielding the top and bottom of the spacecraft

basic octagon, polished aluminum shields on electronic assemblies, and numerous

other shields covering the spacecraft exterior surfaces. Surface treatments (e.g.

gold plating and polished aluminum) having particular thermal properties were

also used extensively on spacecraft components.

An engineering experiment called the absorptivity standard (fig. 3-31) was

included on the spacecraft to enable: (1) the measurement of the solar absorptance

of four typical spacecraft surfaces in direct sunlight; (2 ) the measurement of the

change or degradation of the surface properties with time in space; and (3) thedetermination of the accuracy of flight predictions based upon the temperatures

measured during space simulator testing. The instrument measured the tempera-

ture of an irisuiated f b t piate normal to t'he direction of solar irradiation by

signaling the time at which a known temperature was reached. A white paint, a

black paint, an aluminum silicone paint, and polished aluminum were the sample

surfaces.

Science

The Mariner-Mars 1964 science subsystem consisted of scientific instruments

and ancillary equipment selected to meet the mission objectives. Interplanetary

FIGURE3-29. -Louver assemblies installed on

spacecraft.

94



SPACE VEHlClE SYSTEM DESIGN AND TESTING OPERATIONS

cruise-science instruments performed experiments during the trip, and planetary-

science experiments were conducted at Mars encounter. The science instruments

may be described as follows:

Television.-The television experiment, to provide topographic reconnais-

sance of part of the surface of Mars, used a single camera employing a narrow-

angle telescope. The camera (fig. 3-32) was a shuttered system utilizing a slow-

scan vidicon capable of storing an image with negligible degradation for a 24-sec

frame time. The instrument took and encoded television pictures upon receipt of

frame and line start signals from the data automation system (DAS). It was

composed of seven functional parts (fig. 3-33) :

1. Optics and shutter, the mechanical assembly of the telescope and the

combination shutter-filter. The telescope was an f /8 Cassegrain with a 30.48-cm(12-in.) effective focal length and a 1.05" by 1.05" field of view. Filtering was

accomplished by a disk with two cutouts which contained red filters and two cut-

outs which contained green filters. These filters were introduced into the light path

by a rotating solenoid and remained until a second current pulse energized the

solenoid. Eight such pulses caused a complete rotation of the shutter filter wheel.

FIGURE -3O.-Blanket shielding for temperature control.

95



MARINER-MARS 1964 .

FIGURE-30.-Concluded.

2. Camera head, containing the vidicon tube, preamplifier circuitry, 110-kHz

oscillator, and filtering and distribution circuitry. The vidicon tube was a photo-

sensor that provided the electrical video signal. The 110-kHz oscillator provided

vidicon beam modulation which served as a carrier for amplitude-modulated

video information. The video signal was amplified by the tuned preamplifier.Filtering and distribution of the voltages required for operation of the vidicon tube

were accomplished through the appropriate circuitry.

3. Sweeps, consisting of the horizontal and vertical sweep circuitry and the

4-kHz oscillator. This circuitry provided the voltages necessary for horizontal and

vertical deflection of the vidicon beam both during picture readout and during the

time that the vidicon target was being erased and primed for the next picture.

96




(_ POLISHED ALUMINUM CYLINDER

(_) DOUBLE CONICAL SUPPORT

(_) ALUMINUM MOUNTING RING

(_ POLYCARBONATE PLASTIC BASE

(_) POLYCARBONATE PLASTIC COVER

(_ TEMPERATURE TRANSDUCER

(_ ALUMINIZED MYLAR WRAPPED SUPPORT CONE

(_) ALUMINIZED MYLAR DISCS

SAMP

FIGURE 3--31 .--Absorptivity standard.

FIGURE 3 32.--Sectional view of television camera.



MARINER-MARS 1964b

SWEEPS

CAMERA

HEAD

II

OPTICS

AND

SHUTTER

CAMERA

CONTROL

AND

COMPUTER

I POWER

SUBSYSTEM

, I

I TRANSFORMER _ 2.4 kc 1 50 V rms

RECTIFIER

I[ DAS

I D PLANET IN VIEW

c PLANET IN VIEW

"D'[-'-_'_r'VAIDE, O, l--lIJ TELEDVTSION _ :II:IIT:_ ;IICc;;:E E _AA;: Oz_E;s

I ENCODER _1 ANALOG-TO- PULSEWIDTH DATA

AM RA _ LINE S TA RT C OM MA ND

-- ; 22;c02;ND

DATA

ENCODER

TEMPERATURE SUBSYSTEM

TRANSDUCER

I

FIGURE 3-33.--Functional block diagram of television instrument.

Voltages were adjusted to scan a 0.56- by 0.56-cm (0.22- by 0.22-in.) area of the

vidicon target. The 4okHz oscillator was gated on during the interval between

pictures and caused the vertical sweep voltage to vary at that rate, providing

erasure of the image on the vidicon target.

4. Camera control and computer, containing gain-control computer, planet-

acquisition, and control circuitry. By means of this circuitry, simple computer

operations were performed, and control of certain functions affecting subsystem

operation and dynamic range was provided.

5. Camera control logic, consisting of the external DAS command logic, the

divide-by-25 logic control circuitry, frame control logic, and control logic gates

circuitry. These circuits provided logic control of all camera timing functions and

the necessary switching functions.

6. Video channel, consisting of an RF amplifier and detector and the video

output amplifier. Amplification, demodulation, and dc clamping for the video

signal were provided.

98



t


7. Television data encoder, containing a data output conditioner, accumu-

lator and shift register, analog-to-digital comparator and gates, 1-MHz clock and

gate, ramp generator, matrix, matrix buffers, programer, ground-support-equip-

ment buffer circuitry, analog-to-pulsewidth converter circuitry, and line syn-

chronization generator circuitry. The encoder provided analog-to-digital conver-

sion of the television picture data to 6 bits per picture element. Also, analog-to-

pulsewidth conversions were performed on certain critical television performance

data, and line synchronization information was reconstructed.

Helium magnetometer.--The magnetometer instrument (fig. 3-34), designed

to measure the magnetic field strength in three orthogonal axes, consisted of a

sensor, an oscillator and phase shifter, a commutator, an RF power supply, and

servoamplifiers. The sensor and electronics modules are shown in figure 3-35.This experiment employed the principle that the absorption of light by a cell

containing helium gas is a function of the angle between the ambient magnetic

SENSOR I

I

_'_ IGNITER

BACKUP

COMMAND I

I

Z-AXIS

BANDPASS I

AMPLIFIER :

AND

COMMUTAT_

BANDPASS

AMPLIFIER

ANDCOMMUTATOR

II

PHA SE SHIFT

AND

DEMODULATOr

DRIVE

DEMOOULATOF

DRIVE

PHA SE SHIFT

_, AND

CLOCK PULSE DEMODULATOI

DRIVE 1

IGNITER

'q_-- BACKUP

DATA

AUTOMATION I COMMAND

:i: Z-AXIS DATA

:i: X-AXIS DATA

I

:b Y-AXIS DATA

3 2

4

CALIBRATE

POWER SUBSYSTEM

FmuI_E 3-34.--Functional block diagram of helium magnetometer.

99



MARINER-MARS 1964

MODULE CONTA!N!NG MODU LE COkTAlNlNGBASIC SIGNALCHAIN CIRCUITRY

RF POWER SUPPLY,IGNIT ION LOGIC. HELIUMMAGNETOMETER POWER SUPPLY, AND ANCILLARY

AND ANALOG -TO-WLSE WIDTH EOUIPMENTCONVERTERS .

INCHES (SURROUNDEDBY COILS)

FIGURE-35.-Helium magnetometer sensor and electronics modules.

field and the optic axis of the cell. The instrument was mechanized with calibra-

tion circuitry which, by command from the data automation system (DAS),

superimposed a sequence of magnetic steps at the sensor to allow calibration of the

combined instrument output through the DAS for periodic checks of both linearity

and scale factor. The cycle was initiated at the beginning of each DAS sequence.

Cosmic dust detector.-This experiment was designed to make direct meas-

urements of dust-particle momentum and mass distribution, particularly in

regions encountered having unusually high numbers of such particles. The instru-

ment is shown in figure 3-36, and a functional block diagram is given in figure

3-37. The detector was a single assembly containing a sensor plate mounted on the

top cover of an electronics chassis. The plate had a microphone bonded to one side

and penetration capacitors on both sides to provide measurements of dust-particle

impacts. The sensor was exposed to space, but the electronics chassis was protected

beneath the thermal blanket.

Ionization chamber.-By this experiment, it was hoped to detect and meas-

ure the omnidirectional flux of corpuscular radiation in space and in the vicinity

of Mars, and to measure the ionization produced by this flux. The instrument isshown in figure 3-38. The first of the two detectors which comprised the instru-

ment was an integrating ionization chamber of the quartz-fiber variety housed in

a 12.7-cm-diameter (5-in.) stainless-steel sphere filled with argon gas at four times

atmospheric pressure. The second detector was a conventional, halogen-quenched

Geiger-Mueller (GM) type 10311 counter tube mounted in a thin-wall, stainless-

steel cylinder 12.7 cm (5 in.) long. Both detectors were mounted to a gold-plated

100



.

C A L I B R A T I O NC I R C U l T S

..I

* CALIBRATE


FIGURE -36.-Cosmic dust detector

tDAS

DIRECTF I L M

THRESHOLD

. RESET

C - EADOUT 4 PEADCOMMANDI R C U I T

I B I TRETRO C MICROPHONE

tFILM+ THRESHOLD * GAT E AND

I B I T

3 B I TSA M P L I F I E R F I R S T

PUL SE-HE IGHT ENABL EA N 0 C HIT

+----m POWER2 OWER SUBSYST EM

SUPPLY

FIGURE -37.-Functional block diagram of cosmic dust detector.

101



MARINER-MARS 1964

FIGURE -38.-Ionization chamber.

chassis housing the high-voltage power supply and the pulse-amplifying and pulse-

shaping electronics. The package was mounted part way up the cylindrical wave-

guide on the spacecraft.

A functional block diagram of the instrument is given in figure 3-39. When

the ionizing particles penetrated the sphere, the gas became ionized, The resulting

positive ions were attracted to the ground sphere and the electrons attracted to a

collector rod, tending to neutralize its positive charge. When the collector charge

was sufficientIy neutralized, an image charge was induced and attained a sufficient

magnitude that the fiber inside the sphere was pulled to the collector and con-

tacted it. A surge of electrons then passed from the collector to the fiber and

through a load resistor. The current pulse was amplified, shaped, and presented to

the DAS. Since each output represented a fixed amount of charge collected from

102




OAS

PREAMPLIFIER _-_ OUTPUT

AMPLIFIER IL ION CHAMBERDATA

G EIGER- M UELLER_

1 03 11 TUB E

PREAMPLIFIER

ION CHAMBER SUPPLY

GEIGER-MUELLER 10311 SUPPLY

AMPLIFIER SUPPLY

___ OUTPUT ]

AMPLIFIER J

POWER

SUPPLY

GEIGER-

MUELLER I0311

DATA

POWERSUBSYSTEM

]_ SPACECRAFTPOWER

I

FIGURE 3-39.--Functional block diagram of ionization chamber.

the gas, the interval between pulses was inversely proportional to the ionization

rate.

Charged particles which penetrated both the steel cylinder and the glass wall

(location of the cathode) of the GM tube and entered into the sensitive volume of

the tube ionized the gas molecules. The high electric field between the anode (a

wire running along the axis of the tube) and the electrode accelerated the dis-

sociated ions and electrons. The electrons accelerated rapidly because of their low

mass, ionized other gas molecules, and thus produced a momentary current flow

through an appropriate load resistor. The resulting pulse was amplified, shaped,

and delivered to the DAS. Since one output pulse was generated for each charged

particle that penetrated the shield and entered the sensitive region of the tube, the

pulse rate was directly proportional to the radiation flux.

Cosmic ray telescope.--A cosmic ray telescope (fig. 3-40) was included in the

science subsystem to measure certain charged particles approaching within a

conical acceptance zone. The detector system consisted of three gold-silicon, solid-

state, surface-barrier detectors arranged as a telescope, with absorbers placed

between each. The detectors were positioned so that the half-angle of the cone of

103



MARINER-MARS 1964 b.

\---ELECTRONICSW S S I S

FIGURE-4O.--Cosmic ray telescope.

acceptance of the telescope was 20". When an ionizing particle passed through a

surface-barrier detector, the detector put out a pulse of charge whose size was

proportional to the energy lost by the particle in the depletion layer of the de-

tector. A charge-sensitive preamplifier and an amplifier were connected to each

detector. A voltage pulse proportional to the charge collected from the detector

was generated at the output of an amplifier to be passed on by a pulse-height dis-

criminator. Logic circuits provided the current rates, which, together with the

pulse-height analysis data, were recorded in two 10-bit registers to be read out

alternately by the DAS. A power converter changed the 2400-hertz spacecraft

power into the various dc voltages required by the detectors and electronics.

Figure 3-41 illustrates the operation of this instrument.

Trapped radiation detector.-The trapped radiation detector (fig. 3-42) was

designed to measure the distribution, energy, and identity of magnetically trapped

particles in the vicinity of Mars. Included were three GM tubes (detectors), asilicon, solid-state, surface-barrier proton detector, electronic discriminators and

amplifiers, and a power supply. The functional operation of the instrument is

illustrated in figure 3-43. The GM tubes measured the total number of charged

particles passing through their sensitive volumes after entering at the end of the

tubes. By allowing for omnidirectional flux of higher energy particles, a directional

measurement was made of the low-energy particles. The outputs were shaped by

104



SPACE VEHICLE SYSTEM DESIGN AND T€STING OPERATlONS

PULSE RESET IMODE I!

FIGURE-41 .-Functional block diagram of cosmic ray telescope.

FIGURE -42.-Trapped radiation detector.

105



MARINER-MARS 1964f_

AMPLIFIERND OUTPUT

AMPLIFIER tND OUTPUT

AMPLIFIER IND OUTPUT

DETECTORD

I_..__ PREAM PLI F[ER

1I POSTAM PLIFIER

I _ '

POSTAM PLIFIER

RATE,M,TERIOSTAMPLIFIER DISCRIMINATORI v I AND OUTPUT j

DISCRIMINATOR /EFERENCE

_ _ RAT E LI MiTE R tOSTAMPLIFIER DISCRIMINATOR AND OUTPUT

- IOWER

SUPPLY

TEMPERATURE LRANSDUCER

DAS

DETECTOR AOUTPUT

_'_ DETECTOR 8

OUTPUT

DETECTOR COUTPUT

_--- DETECTOR DI

OUTPUT

DETECTOR D2OUTPUT

POWER

SUBSYSTEM

DATAENCODER

FIGURE 3--43.--Functional block diagram of trapped radiation detector.

means of saturating current amplifiers before being sent to the DAS. The solid-

state detector output was fed into a linear charge sensitive amplifier (preampli-

fier), followed by a series of highly stable negative-feedback voltage amplifiers

(postamplifiers). Voltage gains of the postamplifiers were different and were set to

produce pulses, corresponding to the particles detected, at two identical amplitude

discriminators.

Solar plasma probe.--To measure the densities, velocities, temperatures, and

directions of movement of low-energy protons streaming outward from the Sun, a

solar plasma probe was included in the science subsystem. The instrument (fig.

3-44) measured and recorded the magnitudes of the positive-ion currents inter-

cepted by each of the three sectors of a current-collecting electrode in each of 32

106



SPACE VEHICLE SYSTEM DESIGN AND JESTING OPERAJIONS

FIGURE-44.-Solar plasma probe.

slightly overlapping energy intervals. A high-voltage subsystem modulator, trans-

former, multiplier, and voltage divider supplied selected magnitudes of square-

wave ac to a grid so as to select the various energy intervals. Any one or all three

outputs of the three current-collecting sectors with their individual preamplifiers

(comprising the current measurement chain) could be gated into the summing

amplifier and subsequent circuitry. Two standard dc currents produced by the

calibration signal generator were gated periodically into the three sector pre-amplifiers. The digital control circuitry stepped the instrument through the proper

sequence of measurements. These operations are illustrated in figure 3-45.

The functioning of the science subsystem is illustrated in figure 3-46. An

ultraviolet photometer was a part of the set of scientific experiments initially

selected for the Mariner-Mars 1964 mission; however, because of its unavailability

in time for complete testing, this instrument was replaced by an instrument

107



MARINER-MARS1964

SENSOR lTHREE- SECTIONL IPREAMPLIFIERSI_ I SUMMING

I FARADAY Ivl (THREE) lwl AMPLIFIER

L cup Jl I1

l t CONTROL LOGIC

STANDARD

CALIBRATE

SIGNAL

CONTROL

LOGIC

.ULT,P.,EROLTAGE'" IITRANSFORMERI

TSELECTION CONTROL LOGICREGISTER

BANDPASSFILTER

LOGARITHM IC L,_ SYNCHRONOUS_J MEASUREMENT _

COMPpRLESI_ RON rv I DETECTOR I_1 DIAMOND Ivl

HIGH- 1

VOLTAGE

MODULATOR

_LcoNTROL LOGIC

ANALOG -

TO-PULSE-

WIDTH

CONVERTER

I

T4okc

OSCILLATOR

AND CALIBRATE

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FIGURE3-45.--Functional block diagram of solar plasma probe.

simulator (dummy) which served no purpose other than maintaining inertial

mass and temperature balance in the spacecraft.

In addition to the above experiments, an occultation experiment was chosen

as a Mariner-Mars 1964 planetary-science experiment. However, a special scien-

tific instrument was not required since the experiment relied on the telecommuni-

cations system and on the accuracy of the flight path. This experiment is discussed

in chapter 4 and is described in detail in chapter 7.

The ancillary equipment on the spacecraft consisted of the DAS, the plane-

tary scan system, and the narrow-angle Mars gate. The DAS, a digital computer

located in bay III, automatically controlled and synchronized the data-gathering

sequence of all science instruments and formatted all the diversified science data

into a single, continuous bit stream of ones and zeros for the data encoder, with a

separate bit stream for the video storage subsystem. Each of the science instru-

ments was sampled at a different rate and in a different format. The data were

converted from the forms presented by the instruments (such as serial, parallel,

108




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MARINER-MARS 1964

analog-to-digital conversion, and pulses to be counted) into a single, serial bit

stream to be transmitted to the spacecraft telemetry system (data encoder) for

subsequent transmission to Earth. The only functions which bypassed the DASwere the 2400-hertz power from the power subsystem and the temperature meas-

urements performed by the data encoder. The DAS received and issued com-

mands to the spacecraft as required to perform the science mission. Identification

information was supplied with the data so that each block of data bits was

uniquely identified. Thus, those science measurements represented by any block of

data could be located in time and space through the knowledge of the spacecraft

trajectory. In addition to processing scientific measurements, the DAS also per-

formed science subsystem performance measurements to aid in the engineering

evaluation of the science payload and the calibration of the science instruments.

The real-time portion of the DAS (fig. 3-47) operated continuously during

the mission, being turned on after launch and remaining on until after Mars en-

counter. This portion sequenced the cruise-science instruments, collected their

data, translated the information into constant-rate form, and transmitted the

information to the data encoder to be passed on to the transmitter and then to

Earth. The non-real-time portion of the DAS (fig. 3-48) was dormant during most

of the mission, being turned on a few hours prior to Mars closest approach. Its

primary function was to provide the sequencing for the television recording por-

tion of the encounter. Instead of being transmitted directly to Earth, data from the

non-real-time portion were stored on magnetic tape in the video storage subsystemfor later playback and transmission to Earth at a much slower rate than that

during recording. The real-time and non-real-time portions of the DAS were

completely independent of each other, each naving a separate power supply and

sequencer. The word formats for the real-time and non-real-time portions are

shown in figures 3-49 and 3-50, respectively.

The primary function of the planetary scan system (fig. 3-51) was to find and

hold the correct camera attitude for the television camera as the spacecraft passed

Mars. Its operational sequence (fig. 3-52) was designed as follows: When power

was turned on by either the CC&S command or a ground command, the scan

system entered into a planet-searching operation. The television was mounted on

a scan platform which rotated about the roll axis of the spacecraft through 180 °

of arc, searching for the planet at a scan rate of 0.5 deg/sec. Limit switches

actuated scan reversals when the platform completed each of the 180 ° rotations.

Mounted on the platform was a wide-angle planet sensor with a 50 ° circular field

of view. This sensor detected the presence of the planet and sensed its position with

110




respect to the spacecraft. When the planet came into the sensor field of view, the

scan system generated a wide-angle, planet-in-view signal and switched to planettracking in such a way that the television was oriented toward the desired portion

of the planet. When the spacecraft reached the proper position for picture taking,

a scan inhibit signal was initiated either by a narrow-angle Mars gate (fig. 3-53),

which was also mounted on the platform, or by the television instrument itself.

This signal indicated that the planet was in the television field of view, stopped the

tracking motion, and initiated the picture-taking sequence. Thus, the tape

recorder was triggered to start its recording operation.

When the spacecraft was launched (data mode 2), the science subsystem was

deenergized. On spacecraft separation from the Agena D second stage (still data

mode 2), the real-time portions of the DAS and the interplanetary instruments

were energized. During maneuvers (data mode 1), no science data were trans-

mitted, although the real-time portion of the DAS continued to sample the

instruments. The initiation of the encounter mode by either a CC&S or ground

command was designed to accomplish the following functions:

1. The television, non-real-time portion of the DAS, narrow-angle Mars gate,

planetary scan system, and tape recorder were energized.

2. The non-real-time portion of the DAS sequenced the television, but did

not issue start and stop commands to the tape recorder.

3. The protective science cover was removed from the television, narrow-angle Mars gate, and planetary scan system sensors. The scan platform moved

through 180 ° arcs until the planet was acquired. An output from the planetary

scan system through the DAS to the data encoder initiated data mode 3 (science

data). The planetary scan system then tracked the planet to insure that the plat-

form continued to point at the bisector of brightness defining the location of Mars.

When the edge of the lighted disk came into view of either the television or the

narrow-angle Mars gate, the DAS tape-recorder-start circuitry was enabled. The

DAS waited until the beginning of the next standard sequence to start the record-

ing cycle. All science data acquired during encounter were stored in the spacecraft

tape recorder. The measurements made by the real-time portion of the DAS were

sent to Earth by means of the telemetry channel and were also transferred to the

non-real-time portion of the DAS for storage. Scan platform motion was inhibited

during the science encounter (picture-taking sequence) mode.

The science encounter mode and data mode 3 were terminated when the

second end-of-tape signal occurred. A backup stop-recording circuit was also

included in the DAS. This circuit could be set after 18 pictures and reset after 22

!11



MARINER-MARS 1964

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FIGURE 3-49.--Real-time DAS word format.

pictures were taken. The reset function activated the "switch to data mode 2"

command to the data encoder in exactly the same way as in the second end-ol-tape

signal. The spacecraft then stayed in data mode 2 until either a CC&S or ground

command caused the encounter instruments and the non-real-time portion of the

DAS to be deenergized.

The playback of stored science data was initiated by either a CC&S orground command. The science subsystem was deenergized, data mode 4 initiated,

and the stored data transmitted to Earth. The pictures were transmitted individ-

ually; between each picture, the data encoder switched to data mode 1 so that

engineering telemetry could be transmitted and the condition of the spacecraft

evaluated. No new science data were obtained during data mode 4. Upon com-

pletion of picture playback, the spacecraft reverted to data mode 2.

115



MARINER-MARS 1964

REAL-TIME DATA FRAME

f! WORDS WORDS

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280 BITS 140 BITS

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TV AAC

LOGIC

:5 BIT S

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FIGURE 3-50.--Non-real-time DAS word format.

116



..v SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIO NS

*

FIGURE-51 .-Planetary scan system.

TESTING OPERATIONS

Environmental Testing Program

The environmental testing program was established by the Mariner-Mars

1964 project office to meet its requirement for flight-acceptance testing of all

proof test model^,^ flight, and spare spacecraft equipment, and for type-approval

testing of one complete set of prototype spacecraft equipment. All equipment was

required to pass the tests before being considered acceptable for flight. Since it was

felt that a system configuration provided the only true mechanical, thermal, and

electrical environment and interactions for the various spacecraft equipment,

environmental testing of the assembled spacecraft was also included. In general,

the static environments, such as thermal soak and linear acceleration, could be

readily achieved on the subsystem level, but the dynamic environments, such as

jThe proof test model (FTM) is a prototype of the actual flight spacecraft and is used to “prove” the overall

design and any design changes before they are incorporated into the flight spacecraft. Test data on proof-test-

model performance served as standards against which flight spacecraft performance could be judged.

117



MARINER-MARS1964

PREAMPLIFIER.

It"

SENSOR TEMPERATURE SIGNAL I BAS

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' u-2.4 kc

I

_'mm_F. 3-52.--Functional block diagram of planetary scan system.

vibration, thermal-vacuum, and electromagnetic-interference-produced environ-

mental interactions, were quite dependent upon the response characteristics of the

total vehicle and were, therefore, best achieved at the system level.

Type-approval tests were defined to verify designs, and the flight-acceptance

tests were defined to prove the capability of the equipment or spacecraft to with-

stand the environmental conditions expected during flight. The complete environ-

mental testing program covered such diverse environments as vibration, shock,electromagnetic interference, temperature cycling and extremes, humidity, ex-

plosive atmosphere, and magnetic field effects.

The type-approval subsystem-level environmental test levels were intention-

ally severe to compensate for material and fabrication differences in the flight

hardware. In general, these tests were conducted to provide qualification against

118



I.SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

FIGURE-53.-Narrow-angle Mars gate.

environments that were independent of the spacecraft, such as humidity, static

acceleration, and explosive atmosphere (at launch). Ten test environments were

defined. Since all tests were not appropriate for each item, they were performed

only where applicable. The type-approval tests covered all phases of equipment

preparation an d operation from fabrication of the equipment to the actual flight.

The flight-acceptance subsystem-level environmental test levels were in-

tended to be as severe as the expected environment. These tests were limited to twoenvironments in which equipment operation could be evaluated and which could

be simulated in a well-controlled manner. All flight equipment was subjected to

flight-acceptance testing in vibration and thermal-vacuum environments.

In formulating which type-approval system-level tests would be performed,

the expected environment was evaluated. Type-approval testing at the system

level consisted of full-scale vibration, shock, acoustic, space-simulation, thermal-

vacuum, and electromagnetic-interference tests on the proof test model. To com-

pensate for the statistical limitations of the small number of samples and to insure

that faults and inadequacies would be located, the imposed test conditions were,

by selection, more severe than the operational conditions. They were not, how-

ever, intended to be so severe as to exceed reasonable safety limits or to excite

unrealistic failure modes.

Flight-acceptance system-level testing was performed on each of the three

flight-quality spacecraft in such environments as vibration, thermal-vacuum, and

electromagnetic interference. The conditions used during these tests were those

estimated to be the actual operational environmental conditions.

119



MARINER-MARS 1964e

Essentially all magnetic environmental testing was conducted on the assembly

or subassembly level. All proof-test-model and flight spacecraft equipment was

examined on the basis of magnetic qualities to insure meaningful magnetometermeasurements. The effort made at magnetics control during the Mariner-Mars

1964 project was more extensive than that during previous projects.

Also included in the testing program were life tests. The ultimate aim of any

life test is to gain information about the operating characteristics of given equip-

ment as a function of time and, as a result, to be able to estimate the useful life of

the equipment. For the Mariner-Mars 1964 spacecraft equipment, it was neces-

sary to define the test objectives such that the following would be provided:

1. Analysis of such effects as the degradation of performance as a function of

time, the critical failure modes, and the influence of various test environments on

these effects.

2. Support, whenever necessary, in the diagnosis of inflight problems.

3. Indication, whenever possible, of the nature and extent of the likely deteri-

oration of flight hardware.

The units assigned for the life tests were the type-approval units. Each underwent

a wide assortment of environmental conditions ranging from simulated flight and

ambient laboratory conditions to extensive temperature cycling. Various miscel-

laneous subsystem environmental tests were also conducted whenever they were

deemed necessary or desirable.

The small number of subsystem failures or problems occurring during the

system-level flight-acceptance tests proved the adequacy of the subsystem-level

tests that had been performed. A total of 83 design changes were documented as a

result of problems encountered during the environmental tests. Even this number

could have been reduced had it not been for schedule delays which necessitated

conducting certain flight-acceptance tests prior to or concurrently with type-

approval tests. On an ideal schedule, type-approval tests would demonstrate the

need for a design change before flight hardware was fabricated. (The operation of

the equipment during the Mariner IV mission indicated that the environmental

testing program made a significant contribution to the ultimate success of the

mission.)

Miscellaneous Qualification and Developmental Testing

Many special qualification and developmental tests were also conducted to

prove the feasibility and adequacy of the spacecraft mechanical hardware. For use

120



..*

SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS0

in these tests, four test spaceframe structures (in addition to the proof test modeij

were fabricated.

1. Structural test model, composed of mass and structural simulations of all

spacecraft components including electrical cables. This model was used primarily

for structural qualification tests of the basic structure.

2. Temperature control model, composed of thermal mockups of all critical

spacecraft items with flight-equivalent external surfaces and thermal shields. This

model was used primarily in space simulator tests (fig. 3-54).

3. Developmental test model, a structural simulation of the flight spacecraft

structure, composed of mass mockups of all structurally critical spacecraft items.

Its primary use was for structural vibration qualification tests with the Agena D/

spacecraft adapter.4. Extra test model, composed of the basic octagon structure ballasted to the

proper boost weight and moment of inertia. This model was used for -4gena D/

spacecraft separation tests and other interface tests.

FIGURE -54.-Te1nperature control model during space simulator

testing.

121



MARINER-MARS 1964 D

q,

In addition to the tests performed with these spaceframes, many developmental

tests using prototype and test hardware were performed to confirm analyses, verify

problem solutions, and gather critical information essential to mission success.

System Testing Program

System testing of the Mariner-Mars 1964 spacecraft, performed on the three

flight-quality spacecraft and the proof test model, was conducted at both JPL and

the AFETR to test the spacecraft as a complete integrated system. These tests,

including the previously discussed environmental tests, always followed the as-

sembly and subsystem checkout to verify the functional integrity of the spacecraft.

One such test always preceded shipment of the flight spacecraft to AFETR and

another followed upon arrival of the spacecraft at that location. A final system

test was performed before final commitment to the launch pad, after which noflight equipment could be removed or connectors disconnected. During the tests,

the spacecraft were exercised through the entire flight sequence from launch to

picture playback in all primary and backup modes of operation; for example,

several encounter sequences were performed, including one using CC&S com-

mands and another using ground commands. The "combined system test" was the

overall name given to the preliminary on-pad tests performed at AFETR on the

launch vehicle and the spacecraft. Among the tests included in the series was an

exact simulation of the actual launch-day countdown. While these tests were per-

formed at AFETR, the proof test model was used to verify the compatibility of the

JPL Space Flight Operations Facility and the Deep Space Instrumentation

Facility with the Mariner-Mars 1964 spacecraft.

Whereas tests of the proof test model were performed to verify design, those of

the flight spacecraft verified proper operation of the equipment. The flight space-

craft acceptance criteria were: (1) The spacecraft had to operate for 250 hours in a

solar-vacuum environment without any major failures (defined as failures pro-

hibiting the successful completion of the mission) ; and (2) the spacecraft could be

shipped to AFETR only upon successful completion of the final system test atJPL.

(A test was considered successfully completed when no major misoperation

occurred during it.) The minimum qualification time for any subsystem was 100

hours, and the minimum total test time (subsystem and system) was 400 hours.

Test Facilities

During the test and evaluation program, the spacecraft were exposed to

different environments simulating launch and flight conditions. For this usage,

122



.J


three basic JPL and three AFETR environmental facilities were available. They

are as follows:

JPL SpacecraftAssemblyFacility

The spacecraft and the system test complex (discussed later in this section)

were first assembled and tested at this facility, where all hardware to be assembled

was brought together for the first time. Simultaneous testing of three spacecraft

could be conducted in the Spacecraft Assembly Facility.

JPL Spacecraft Environmental Test Facility

In order to verify that the proof-test-model and flight spacecraft could operate

satisfactorily after exposure to flight-acceptance vibration levels (determined fromthose expected of the Atlas D and Agena D during the boost phase), a vibration

exciter with a 133 446-N (30 000-1b) vector force rating, a granite oil table, and a

horizontal test fixture for vibration in all three spacecraft axes were provided.

Other equipment included acoustic chambers, shock equipment, centrifuges, and

climatic test facilities (e.g., temperature, salt, fog, and explosive-atmosphere

chambers).

JPL 7.5-meter (25-ft) Space Simulator

The spacecraft were exposed to the vacuum and thermal conditions simulat-

ing a space environment in this 7.5-meter-diameter (25-ft), 12-meter-high (40-ft)

simulator chamber. The chamber could be evacuated to a pressure of approxi-

mately 10 -6 mm of mercury. Liquid-nitrogen cold walls were used for cooling,

and arc lamps were used for simulating the Sun's heat.

Eastern TestRange Spacecraft CheckoutFacility

The Spacecraft Checkout Facility, Hangar AO, is the principal JPL test

facility at the AFETR. A complete system test complex enabled a detailed check-

out of the spacecraft before launch operations; some of this equipment was also

used during the launch operations. The operations and communications centers

are also housed in this facility.

Eastern TestRange ExplosiveSafe Facility

Located in a remote area, this facility provided a location where the hazard-

ous operations necessary in the final preparation of the spacecraft could be per-

formed just prior to the spacecraft launch operations in the launch complex. Two

123



MARINER-MARS 1964

m

ii

buildings comprise this facility: the sterilization and assembly laboratory, and the

propellant laboratory. The hazardous operations included installation of the

pyrotechnics squib and the midcourse motor, and loading of the liquid propellant.

Eastern Test Range Launch Complex

Certain monitoring and control equipment is installed in the launch complex

for use during the launch operations of a spacecraft. This equipment is housed in a

blockhouse console connected to the spacecraft by cabling and an umbilical con-

nector. Test personnel monitor and control the spacecraft during prelaunch

operations from this area. Spacecraft control equipment, such as power supplies

and signal conditioning units, is located in the launch pad building and on the

umbilical tower and is controlled from the blockhouse during launch operations.

Parabolic antennas located on both the service and umbilical towers provide anRF link between the launch pad, the Spacecraft Checkout Facility, and the Deep

Space Instrumentation Facility areas at AFETR.

Operational Support Equipment

The spacecraft operational support equipment (OSE) provided a tool for

verifying the design and flight readiness of the proof-test-model and flight space-

craft. The equipment was mechanized so that the spacecraft was exercised through

all its standard and backup modes of operation, with enough flexibility to adapt to

nonstandard conditions. Another use was in supporting troubleshooting operations

necessary to locate failures, with the capability to differentiate between OSE prob-

lems and spacecraft problems. The equipment (both electrical and mechanical)

necessary to support subsystem, system, and flight readiness testing was divided

into three categories: system-test-complex equipment, launch-complex equip-

ment, and mechanical support equipment.

System-test-complex equipment

The system test complex (fig. 3-55) consisted of approximately 45 racks of

equipment capable of evaluating and troubleshooting on both the subsystem and

system levels. Essentially the same configuration was used at both the JPL Space-

craft Assembly Facility and the Eastern Test Range Spacecraft Checkout Facility.

This equipment served to: (1) differentiate between spacecraft and OSE failures;

(2) exercise and/or activate the OSE subsystems to simulate flight-type opera-

tions; (3) monitor and evaluate the OSE subsystems during subsystem and system

tests; (4) provide self-power for subsystem test and evaluation; and (5) provide the

124



.. SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

.I

m 1I

FIGURE-55.-System test complex.

necessary input, stimulus, and output loading required for spacecraft subsystem

testing and troubleshooting. The equipment was divided into four groups, which

are described in the following paragraphs.

The first group was the direct counterpart of the spacecraft flight subsystem

and contained such units as radio, command, power, CC&S, data encoder,

attitude-control, pyrotechnics, video storage, and science subsystems. The second

group, which completed the command and telemetry group, contained the

ground command subsystem for sending commands to the spacecraft and the

ground telemetry subsystem for demodulating, decommutating, distributing, and

printing out the telemetry data. The ground telemetry subsystem assembly is

shown in figure 3-56. The third group contained the following support items: the

central recorder, the central timing system, the power distribution system, the

telemetry processing system, and the telemetry teletype encoder. A computer data

system (fig. 3-57) consisting of a medium-size general-purpose computer and a

data input system (fig. 3-58) made u p the fourth equipment group.

125



.MARINER-MARS 1964 .

FIGURE-56.-Ground telemetry subsystem assembly.

126



*.. SPACE VEHICLE SYSTEM DESIGN AND TESTING OPERATIONS

I

I I!

FIGURE-56.-Concluded.

Some ancillary equipment was also provided. Laid out in a rectangular pat-

tern with the subsystem test consoles forming the periphery, the complex was

located on a raised floor; provisions for rigidly mounting and positioning the

spacecraft were included. Contact of the system test complex with thc spacecraft

was made in three ways: (1) through the separation and the umbilical connectors,

which provided an interface with the entire spacecraft system; (2) through the RF

127



MARINER-MARS 1964

TEST DATA INPUTS

176 ANALOG CHANNELS

eOEVENT CHANNELS

144 STATUS CHANNELS

16 TRANSIENT DETECTORS

I0 COUNTERS

PCM TELEMETRY N_I

PCM TELEMETRY NO. 2

PCM TELEMETRY NO.3

TTY FORMAT TELEMETRY

TV DIGITAL VIDEO DATA

P CM P UL SE -C OD E- MO DU LA TE D

TTY TELETYPE

SYSTEM TEST COMPLEXES

II _ PATCH

J OATA INPUT LINE PRINTER

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= 4 TTY PRINTERS

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TELEMETRY DIRECTION BOX

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TELEMETRY _ "_

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I

I DIGITAL DATA CABLES

3700 ft (max)

t

i CENTRAL COMPUTER ARE_ _ PROGRAMING ]I CONSOLE

I= FOUR-UNIT

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CARD READER

-_I PLOTTER 1

DIGITAL COMPUTER

FICURE 3 57.--Block diagram of computer data system.

link; and (3) through direct-access connections to the individual subsystems of

the spacecraft.

Launch-complex equipment

The OSE configuration at the launch complex is shown in figure 3-59. The

launch-complex equipment consisted essentially of six racks of equipment: two

located in the launch-pad building immediately beneath the umbilical tower and

four located in the blockhouse. Contained within the racks were selected elements

of most of the subsystems represented in the system test complex that were re-

quired to support launch operations. In many instances, parallel equipment was

available to support operations in the Air Force Eastern Test Range Spacecraft

Checkout Facility.

128




FIGURE-58.-Data input system.

HANGAR A 0

GTS GROUND TELEMETRY SUBSYSTEM

LCE LAUNCH COMPLEX EOUIPMENT

UMBILICAL

TOWER

PARASITIC

IARD-LINE ANTENNA

R F CONNECTIONS

I L CE I

LAUNCH

BUILDINGBLOCKHOUSE

k

UBlLlCAL

FIGURE-59.-0SE configuration at launch complex.

129



MARINER-MARS 1964

The launch-complex equipment functioned to put the spacecraft in its launch

mode after the spacecraft was mated to the launch vehicle. The equipment located

in the launch-pad building consisted of science OSE, launch support power, and

line-driver amplifiers (required because of the long line cables between the launch-

pad building and the blockhouse). The blockhouse equipment consisted primar-

ily of power control, data encoder, attitude-control, and CC&S OSE subsystems.

Mechanical support equipment

Spacecraft ground handling equipment was used during the spacecraft

assembly, testing, and handling phases. The system was designed to maximize

operational simplicity and allow safe access to the spacecraft components during

all phases of operations. Special attention was paid to the height of the spacecraft

equipment. Such attention to details significantly contributed to the speed, ease,

and safety with which the spacecraft assembly, testing, and handling operations

were accomplished. In the design of the equipment, the following were required:

The equipment must provide stiff structural support for the spacecraft structure

during ground tests; the design must allow personnel to work with and around the

equipment safely and efficiently; the equipment must contain as few parts as

possible; equipment for use in the vicinity of the spacecraft must be capable of

being kept clean to prevent contamination of the spacecraft; and commercial parts

should be used in the equipment wherever possible.

The ground handling equipment consisted of: (1) a universal support ring tosupport the spacecraft octagon when it was being moved in approximately the

same manner as when it was mated to the booster adapter; (2) a spacecraft posi-

tioner to tilt and turn the spacecraft; (3) a dolly for moving the spacecraft within

the work area; (4) a universal ring spacer to raise the spacecraft when mounted on

the dolly to an optimum working height; (5) a transport trailer for interarea space-

craft transport; (6) hoisting equipment to lift the spacecraft; (7) shipping con-

tainers for use during shipment of the spacecraft around JPL, around AFETR,

and between these two locations; and (8) a purging-control assembly for purging

the spacecraft when it was under either the shipping bag (of the shipping con-

tainer) or the flight shroud.

Another piece of mechanical support equipment was a magnetometer test

fixture used to support the complete spacecraft, less solar panels, when the space-

craft magnetic field was mapped. The fixture allowed the spacecraft to be rotated

and displaced about any axis through the magnetometer sensor.

130



Q

CHAPTER 4

Flight History and Space Vehicle Performance

CHRONOLOGY OF MARiNER-MARS 1964 FLIGHTS

Moriner III Flight

After two scheduled and four unscheduled holds in the countdown, Mariner

III, the first of the Mariner-Mars 1964 flight spacecraft, was launched from

Launch Complex 13 at AFETR at 19:22:04.92 GMT 1 on November 5, 1964.

(Launch had been delayed from the preceding day, which was the first day of the

launch interval, to permit a test series on certain relays replaced in the Agena D

second stage.) All systems appeared to be functioning properly during the initial

phase of the flight. At approximately 60 min after launch, it was confirmed that

the science instruments were on, but there was no indication of solar panel power.

A command was sent to the spacecraft to turn off the gyros to conserve battery

power and thus extend the life of the spacecraft, giving additional time for possible

corrective action.

After turnoff of the gyros, both engineering and science telemetry data gave

indications that either the Agena D second stage or the shroud had not separated

from the spacecraft. The separation velocity was calculated to be that which would

be approximately normal with the shroud on. Other factors also verified that the

shroud was still on the spacecraft. The execution of a maneuver designed to shake

the spacecraft free of the shroud was not verified, since the last telemetry reception

from the spacecraft occurred prior to the time of the maneuver start. Battery

power was depleted 8 hours 43 min after launch. The spacecraft did, however,

attain Earth-escape velocity and went into orbit about the Sun.

An extensive 24-hr/day test program directed by Lewis Research Center

(LeRC), the Lockheed Missiles and Space Company (LMSC), and JPL was

started the following day. The probable failure mode was found to be a structuralfailure caused by skin separation from the fiber-glass honeycomb core of the

shroud. A LeRC/LMSC/JPL task team was then formed to design and qualify an

all-metal shroud in time to allow launching of the second Mariner-Mars 1964

spacecraft in that launch interval. An intensive effort by this team enabled the

1 Hr:min:sec Greenwich Mean Time (GMT). All times given in this document are expressed as GMT.

131



MARINER-MARS 1964

delivery of the first replacement shroud to AFETR within 17 days of the Mariner

III failure. In the new design, the fiber-glass structure was replaced by a mag-

nesium section with an inner thermal liner. As a backup effort, JPL also con-structed and tested an all-metal shroud using some parts of a previously damaged

test shroud. Meanwhile, a second all-metal shroud built by LMSC was subjected

to tests to qualify the design for flight. With the final test completed on November

26, November 27 was established as the launch date for the second Mariner-Mars

1964 spacecraft.

Mariner IV Flight

Launchto Canopus acquisition

Mariner IV, the second of the Mariner-Mars 1964 flight spacecraft, was

launched from Launch Complex 12 at AFETR at 14:22:01 GMT on November

28, 1964 (fig. 4-1). (The first launch attempt on November 27 was "scrubbed," since

insufficient time remained in the launch window to evaluate the anomaly which

caused the second of two unscheduled holds.) The ascent trajectory profile is

shown in figure 4-2, and the sequence of events for this phase of the flight is

illustrated in figure 4-3. Critical events and times of the flight from launch until

the end of the mission are listed in table 4-I. The initial launch azimuth was 90.5 °

east of true north. After liftofl, the booster rolled to an azimuth of 91.4 ° and per-

formed a programed pitch maneuver until booster-engine cutoff (BECO). 2

During the sustainer and vernier stages, adjustments in vehicle attitude and enginecutoff times were commanded, as required, by the ground guidance computer to

adjust the altitude and velocity of the vehicle at the Atlas D vernier-engine cutoff

(VECO).After Atlas D/Agena D separation, there was a short coast period prior to the

first ignition of the Agena D second stage. The Agena D engine ignited and, at a

preset value of sensed velocity increase, was cut off. At that time, the Agena D/

spacecraft combination was coasting in a nearly circular parking orbit around the

Earth in a southeasterly direction at an altitude of 189 km (117.5 statute miles)

and an inertial speed of 7.9 km/sec (4.88 statute miles/sec). The parking orbit,

illustrated in figure 4-4, was inclined 28.3°; apogee (point in orbit farthest from

the Earth) was 184.2 km (114.5 statute miles); and perigee (point in orbit nearest

to the Earth) was 172.2 kln (107 statute miles).

2 Because of the Earth's rotation, the direction in which the booster should fly is constantly changing; thus, the

Atlas D rolled to its proper bearing shortly after liftoff and performed a gradual pitch maneuver from the vertical

in the desired flight direction.

132



*(.

4FLIGHT HISTORY AN D SPACE VEHICLE PERFORMANCE

FIGURE-1 .-Launch of Mariner IV spacecraft.

After a parking-orbit coast time of 32 min 1 5 sec, as determined by the

ground guidance computer and transmitted to the Agena D during the Atlas D

vernier stage, a second ignition of the Agena D engine occurred. This engine was

cut off 95 sec later with the Agena D/spacecraft combination in a nominal Earth-

Mars transfer orbit. Injection (second Agena D cutoff) took place over the Indian

Ocean at a geocentric latitude and longitude of -26.25' and 68.82", respectively.

133



MARINER-MARS 1964

AGENA D SECOND --AGENA D SECOND

IGNITION---_ CUTOFF /SEPARAF--'4GEA/A D/ISoPNACECRAFT

_AGENA O FIRST \ T

ooo ,ooooo/ _,//.,_oo_" 400 / /y._--_,46EN,4 O FIRST IGNITION

?- 300 _'/_K_ VECOC--: 200_/ _.,,__ _-SECO (Inouticglmile=l.15stotutemiles=l.852km)

2/,-_oo ,o _BECORANGE, nouticol miles

FIGURE 4-2.---Ascent trajectory profile of Mariner IV spacecraft.

/

\\

\\

\\

\\

\\\

\6000

,_ 7 6 5

9

(I) LAUNCH _ 8

(2) BECO

(3) SECO, VECO, SHROUD EJECTION, AND _

ATLAS D / AGEN4 D SEPARATION

(4) AGENA D FIRST IGNITION

(5) ,4GEN.4 D FIRST CUTOFF (START OF PARKING ORBIT) _'-_ INJECTION

(6) AGENA O SECOND IGNITION (END OF PARKING ORBIT) ( ...... _-7

(7) AGENA D SECOND CUTOFF (INJECTION) AND _'_'"'m_ 6 [_A_E,,,,,/SPACECRAFTEPARAT,ON"I SUNCOO,S'T,ON LA'-'NC"V5(9) CANOPUS ACQUISITION I 2-4

FIGURE 4 3.--Typical sequence of events froul launch to Canopus acquisition.

134



FLIGHT HISTORY AND SPACE VEHICLE PERFORMANCE

Table 4-1.--Critical events and times of the Mariner IV mission a

Event ]

Launch ...............................................................

Date

Nov. 28, 1964

Atlas D/Agena D separat ion . .. ... .. .. ... ... ... .. .. ... ... ... .. .. ... ... ..

Agena D first ignition ........... .......................................

Agena D first cutoff . .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

Agena D second ignit ion . .. .. ... ... ... .. .. ... ... ... .. .. ... ... ... .. .. ...

Agena D second cutoff (injection) .......................................

Entrance of Agena D/spacecraft into Earth's shadow .......................

Agena D/spacecraft separation; cruise-science instruments on; data mode 2 ....

CC&S L-1 event (deploy solar panels, unlatch scan platform; preempted by ac-

tion of separation-lnitiated timer on spacecraft) .......................Exit of spacecraft from Earth's shadow .................................

CC&S L-2 event (turn on attitude-control subsystem; preempted by action of

pyrotechnics arming switch at Agena D/spacecraft separation) ..........

Sun acquisition .....................................................

CC&S L-3 event (energize Canopus sensor, turn on solar pressure vanes) ....

3 DC-21 (roll override: negative increment) transmissionsb ..................

Canopus acquisition ....................................................

QC1-1, -2, and -3 (pitch turn, roll turn, and motor-burn durations ) transmissionb

DC-29 (arm first propulsion maneuver) transmissionb ......................

DC-14 (remove maneuver command inhibit) transmissionl, ..................

DC-27 (initiate midcourse maneuver) transmissionb .........................

DC-13 (inhibit maneuver command, inhibit propulsion command ) transmissionb

3 DC-21 (roll override: negative increment) transmissions h .................

7 DC-21 (roll override: negative increment) transmissionse .................

Canopus reacquisition ... .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. .. ..

QC-1, -2, and -3 (pitch turn, roll turn, and motor-burn durations) transmissionb

DC-29 (arm first propulsion maneuver) transmissionb ......................

DC-14 (remove maneuver command inhibit) transmission b .................

DC-27 (initiate midcourse maneuver) transmissionb ........................

Start of midcourse motor burn ..........................................

Nov. 28, !964

Nov. 28, 1964

Nov. 28, 1964

Nov. 28, 1964

Nov. 28, 1964

Nov. 28, 1964

Nov. 28, 1964

Nov.

Nov.

28, 196428, 1964

Nov. 28, 1964

Nov. 28, 1964

Nov. 29, 1964

Nov. 30, 1964

Nov. 30, 1964

Nov. 30, 1964

Nov. 30, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dee. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dee. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 4, 1964

Dec. 5, 1964

Dec. 5, 1964

Dec. 5, 1964

Dec. 5, 1964

Dee. 5, 1964

Dec. 5, 1964

Dec. 5, 1964

Dec. 5, 1964

See footnotes at end of table.

Time, GMT,hr:min:sec

14:22:01

14:27:23

14:28:14

14:30:38

15:02:53

15:04:28

15:05:51

15:07:09

15:15:0015:17:35

15:19:00

15:30:57

06:59:00

09:13:00

10:45:00

10:59:07

11:02:47

13:05:00

13:10:00

13:15:00

13:45:00

14:05:00

14:35:00

14:47:13

15:22:00

15:32:00

16:02:00

22:40:00

23:04:00

23:05:01

23:05:59

23:40:00

23:57:00

23:58:00

00:02:44

13:05:00

13:10:00

13:15:00

13:45:00

14:05:00

14:25:00

16:09:11

135



MARINER-MARS 1964

B

Table 4-1.--Critical events and times of the Mariner IV mission--Continued

Event

Sun reacquisition .....................................................

DC-21 (roll override: negative increment) transmission b ....................

Canopus reacquisition .................................................

DC-7 (switch power amplifiers) transmissionb ..............................

DC-21 (roll override: negative increment) transmission c ....................

Canopus reacquisition ..................................................

DC-15 (Canopus gate inhibit override) transmissionc ........................

CC&S MT-6 (switch data encoder bit rate to 8a_ bits/sec)command .........

DC-3 (command data mode 3) transmission b .............................

DC-2 (command data mode 2, turn on cruise science) transmissionb ..........

DC-26 (turn off planetary science, cruise science, and battery charger) trans-

missionb

DC-2 (command data mode 2, turn on cruise science) transmissionb .........

DC-28 (turn on battery charger, turn off video storage subsystem) transmission I'.

DC-25 (turn on planetary science, unlatch cover) transmission h . ............

DC-24 (inhibit scan search) transmissionb ...............................

DC-28 (turn on battery charger, turn off video storage subsystem) transmission b .

DC-3 (command data mode 3) transmissionb .............................

DC-2 (command data mode 2, turn on cruise science) transmission b .........


missionb

DC-2 (command data mode 2, turn on cruise science) transmission b ........

CC&S MT-1 (lst Canopus sensor cone angle update) command ...........

CC&S MT-5 (switch transmitter to high-gain antenna) command ..........

CC&S MT-2 (2d Canopus sensor cone angle update) command ...........

CC&S MT-3 (3d Canopus sensor cone angle update) command ............

CC&S MT-4 (4th Canopus sensor cone angle update) command ...........

DC-25 (turn on planetary science, unlatch cover) transmissiond ............

CC&S MT-7 (encounter science on; backup) command (preempted by DC-25)

DC-24 (inhibit scan search) transmissiond ................................

DC-3 (command data mode 3) transmissionb .............................

Wide-angle acquisition at spacecraft (action preempted by DC-24 and DC-3)..

DC-16 (narrow-angle acquisition) transmission (preempted by action at narrow-

angle acquis it ion)

Narrow-angle acquisition at spacecraft ...................................


missionb

6 DC-2 (command data mode 2, turn on cruise science) transmissions b ........


Date

Dec. 5, 1964

Dec. 5, 1964

Dec. 5, 1964

Dec. 13, 1964

Dec. 17,1964

Dec. 17,1964

Dec. 17, 1964

Jan. 3,1965

Feb. 11, 1965

Feb. ll,1965

Feb. 11, 1965

Feb. 11,1965

Feb. 11, 1965

Feb. 11,1965

Feb. ll,1965

Feb. 11, 1965

Feb. 11,1965

Feb. 11, 1965

Feb. 11,1965

Feb. 11, 1965

Feb. 27,1965

Mar. 5, 1965

Apr. 2,1965

May 7, 1965

June 14, 1965

.July 14, 1965

July 14, 1965

July 14, 1965

July 14, 1965

luly 14, 1965

July 15, 1965

July 15, 1965

July 15, 1965

July 15, 1965July 15, 1965

July 15, 1965

July 15, 1965

July 15, 1965

,July 15, 1965

Time, GMT,

hr:min:sec

16:21:07

16:52:00

16:58:19

14:09:00

16:00:00

16:06:22

17:30:00

16:59:54

03:29:29

03:36:13

03:53:15

04:15:51

04:32:39

06:54:43

08:59:23

09:13:51

09:30:56

10:21:20

10:27:08

10:49:35

17:02:19

13:02:37

14:25:1517:27:25

15:51:45

14:27:55

15:41:49

17:10:18

22:10:29

23:42:00

00:11:57

00:17:21

00:31:42

00:32:40

00:37:00

00:42:00

00:47:00

00:52:00

00:57:00

136




Table 4-1.--Critical events and times of the Mariner IV mission--Continued

Event

Spacecraft's closest approach to Mars ....................................

Spacecraft's entrance into Earth-occultation region of Mars ................

Spacecraft's exit from Earth-occultation region of Mars ...................

CC&S MT-8 (encounter science off) command (preempted by DC-26) .......

CC&S MT-9 (cruise science off, start data playback) command ...........

Start of first picture playback ..........................................

End of f irst picture playback .. ... .. .. ... ... ... .. .. ... ... ... .. .. ... ... .

Start of 2d picture playback ...........................................

DC-28 (turn on battery charger, turn off video storage subsystem) transmissione.


missione

DC-2 (command data mode 2, turn on cruise science) transmissione .........

End of second picture playback ........................................

DC-25 (turn on planetary science, unlatch cover) transmissionf ............

2 DC-28 (turn on battery charger, turn off video storage subsystem) trans-

missionsf

2 DC-26 (turn off planetary science, cruise science, and battery charger) trans-

missionsf

7 DC-2 (command data mode 2, turn on cruise science) transmissionst ........

DC-13 (inhibit maneuver command,inhibit propulsion command) transmissionf

QC-1, -4, and -3 (pitch turn, roll turn, and motor-burn durations) trans-

missionsf

DC-17 (cycle Canopus cone angle) transmissionf .........................

DC-25 (turn on planetary science, unlatch cover) transmissiont ............

DC-3 (command data mode 3) transmissionf ............................

DC-24 (inhibit scan search) transmissionf ...............................

DC-16 (narrow-angle acquisition) transmissionr ..........................

DC-2 (command data mode 2, turn on cruise science) transmissionf .........


missiont

DC-22 (change tracks) transmissionf ...................................

DC-4 (command data mode 4) transmissionf ............................

Start of "black space" picture playback .................................

DC-28 (turn on battery charger, turn off video storage subsystem) transmissione


missione

DC-2 (command data mode 2, turn on cruise science) transmissione ..........

End of "black space" picture playback ..................................


Date

July

July

July

July 15,

July 15,

July 15,

July 24,

July 24,

Aug. 3,

Aug. 3,

Aug. 3,

Aug. 3,

Aug. 21

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 21,

Aug. 26,

Aug. 26,

Aug. 27,

Aug. 30,

Aug. 30,

Aug. 30,

Aug. 30,

Aug. 31,

Aug. 31,

Aug. 31,

Aug. 31,

Sept. 1,

Sept. 2,

Sept. 2,

Sept. 2,

Sept. 2,

Time, GMT,

hr:min:sec

15, 1965 01:00:58

15, 1965 02:19:11

15, 1965 03:13:04

1965 05:01:49

1965 11:41:50

1965 12:49:54

1965 19:26:33

1965 21:21:53

1965 03:08:33

1965 03:14:33

1965 03:20:33

1965 03:36:02

, 1965 22:22:00

1965 23:20:00

23:22:00

1965 23:28:13

1965 23:30:13

1965 23:34:13

1965 23:39:00

1965 23:44:00

1965 23:49:00

1965 23:54:00

1965 23:59:00

1965 21:06:52

1965 21:15:16

21:23:40

21:32:04

1965 19:40:00

1965 20:30:00

1965 21:10:24

1965 22:48:33

1965 23:35:26

1965 00:05:00

1965 00:44:00

1965 00:49:00

1965 01:25:00

1965 02:00:46

1965 06:17:00

1965 06:23:00

1965 06:29:00

1965 06:48:32

137



b

n

MARINER-MARS 1964

Table 61.- Cri ti ca l events and times of the Mariner IV mission-Concluded

I IEvent

Time, GMT,Date I hr:min:sec

DC-12 (transmit low, receive low) transmissione. . . . . . . . . . . . . . . . . . . . . . . . . .

Receipt of last signal from spacecraft; end of mission.. . . . . . . . . . . . . . . . . . . . . .Oct. 1, 1965

Oct. 1, 1965

21 :30:17

22:05:07

aAll commands sent to Mariner I V from the deep space stations were verified in the telemetry da ta as having

11 Command(s) transmitted and verified by Pioneer DSS 11.

c Command(s) transmitted and verified by Woomera DSS 41.

d Command(s) transmitted and verified by Johannesburg DSS 51.

eCo mmand(s) transmitted by Venus DSS 13 in uplink lock a nd verified by Echo DSS 12 in downlink lock.

f Command(s) transmitted and verified by Echo DSS 12.

been received and acted upon within 1 min after transmission.

FIGURE-4.-Early flight path of Mariner IV spacecraft.

The trajectory was well within launch tolerances. The Agena D/spacecraft com-

bination was a t an altitude of 198.4 km (123.3 statute miles) and was traveling at

an inertial speed of 11.50 km/sec (7.15 statute miles/sec).

138



Q


At ! min 23 sec after injection, the Agena D ,/spacecraft entered the Earth's

shadow. Separation occurred 1 rain 18 sec later. With the spacecraft operating in

data mode 2 (engineering and science data), the cruise-science instruments and

the real-time portion of the data automation system were turned on. The total

time spent in the Earth's shadow by the Mariner IV spacecraft was 11 rain 44 sec.

Within 1 hour after iniection, the spacecraft was receding from the Earth in

almost a radial direction with decreasing speed. This reduced the geocentric

angular rate of Mariner IV (in inertial coordinates) until the angular rate of the

Earth's rotation exceeded it.

The CC&S command to deploy the solar panels and unlatch the scan plat-

form (L-1 event) occurred on time, but no action resulted since this command

action was preempted by the action of the separation-initiated timer. The CC&Scommand to turn on the attitude-control subsystem (L-2 event) was preempted

by the action of the pyrotechnics arming switch during the already completed

separation of the Agena D. Sun acquisition was completed in approximately 8

rain, and the spacecraft went into a programed roll to furnish ea,.bradun data for

the magnetometer (i.e., so that the magnetic field value of the spacecraft could be

subtracted from the magnetometer readings). Telemetry indicated that the video-

storage-subsystem launch mode was turned off on time at separation.

The CC&S command to initiate acquisition of the star Canopus (L-3 event)

occurred on November 29, with the acquisition sequence beginning at a clock

angle of 60 ° . (The terms "clock angle', and "cone angle" were defined previously

in fig. 3-18.) At a clock angle of 119 ° (in the vicinity of the star Markab), an ac-

quirable object entered the Canopus sensor field of view and was acquired. Telem-

etry indicated immediately that this object was not the star Canopus. (Subse-

quent analysis revealed that it was probably a cluster of stars whose brightness was

augmented by earthlight reflected into the sensor optics.) Later on November 29,

the spacecraft was in a roll search in its automatic reacquisition mode. Another

acquirable object entered the sensor field of view, but the brightness of this object

was shown by telemetry to be approximately one-quarter of the expected Canopus

brightness. It was decided to proceed with Canopus acquisition by means of

ground commands the following day.

The first such roll override command (DC-21) was transmitted on November

30, and the spacecraft went into a roll search. After nearly 60 ° of roll, another star

was acquired. Comparison of the data received with the values expected resulted

in the first positive indication of the spacecraft roll orientation between the star

Regulus and the star Naos. A second command (DC-21) moved the roll reference

139



MARINER-MARS1964e

to the star Gamma Velorum. The final roll override command (DC-21) resulted

in Canopus acquisition (as determined by the light intensity observed by the

Canopus sensor and the saturated condition3 of the Earth detector) and the initia-tion of the premidcourse-maneuver cruise phase of the mission.

Premidcourse-maneuvercruise

This portion of the Mariner IV mission was essentially without incident,

except for the two roll control transients, one of which caused a momentary loss of

Canopus acquisition. The first transient occurred on November 30, but telemetry

showed no correlation with any spacecraft function or event. The second transient,

which occurred on December 2, was the first to demonstrate characteristics which

later became associated with a specific type of repeatable roll transient. Canopus

brightness had been consistently indicating its expected value. Suddenly, the

Canopus brightness for one sample rose above the expected value, and the roll

error signal was less than expected; for the next sample, the Canopus brightness

was less than the expected value, the roll error signal was higher than expected,

and the gyros turned on. Reacquisition occurred almost immediately, and the

gyros turned off. Analysis indicated that not all the roll position errors observed

could be valid. Although the problem remained unexplained, all indications were

that the spacecraft had tracked a very bright object which had passed through the

sensor field of view.

Midway through November 30, the Earth detector came out of saturation,somewhat earlier than expected. Because of the additional confidence in the deter-

mination of roll orientation afforded by the Earth detector data after the detector

came out of saturation (when it could be determined positively that the detector

was indeed sensing the reflected sunlight from the Earth), it was recommended

that a midcourse maneuver be performed as early as possible. The optimum date

was chosen to be December 4 during the pass over Pioneer DSS 11 with December

5 and 6 as alternate dates.

Tracking data then being gathered and analyzed indicated that, without

such a correction, the spacecraft would pass the upper leading edge of Mars at a

closest approach distance of about 246 378 km (153 125 statute miles). Closest

approach would then occur at 01:25:11 GMT on July 17, 1965. To alter the

trajectory so that the spacecraft would pass through a selected aiming region

centered at approximately 12 000 km (7500 statute miles) from the center of

3A condition at which the reading or output from the instrument could no longer increase regardless of whether

the input (reflected sunlight from the Earth) increased or remained constant.

140



Q

FLIGHTHISTORYAND SPACEVEHICLEPERFORMANCE

Mars, a midcourse velocity correction of 16.98 m/sec (55.7 ff/sec) was required.

[The maximum capability was 81 m/sec (267 ft/sec).] In addition to altering thedistance from the planet at the time of closest approach, this correction changed

the arrival time (time of closest approach) to 01:47:00 GMT on July 15,1965, and

thus allowed preset CC&S commands to activate properly various subsystems

near encounter.

Mictcoursemaneuver

On December 4, the sequence of planned midcourse-maneuver events was

followed through the spacecraft receipt of the ground command to initiate the

midcourse-maneuver sequence (DC-27). Almost immediately after the space-

craft responded to the command, Canopus acquisition was lost and the spacecraft

began a roll search. A successful maneuver requires that roll and pitch stabiliza-

tion be maintained by maintaining Sun and Canopus acquisition until the time

of the maneuver pitch turn. The midcourse maneuver was inhibited by a ground

command (DC-13) after it was determined that there was insufficient time to

reacquire Canopus with ground commands (DC-21's) before the start of the

pitch-turn maneuver. Sun acquisition remained stable during this period.

The spacecraft continued to roll until an object was acquired. Three roll-

override commands (DC-21's) were sent to the spacecraft, and short-duration

roll searches were made. Finally, a decision was made to discontinue the attemptto acquire Canopus during the Pioneer DSS 11 pass because of the short time

remaining in the view period of that station.

The spacecraft was then acquired by Woomera DSS 41, and a DC-21 com-

mand was transmitted which resulted in acquisition of the star Regulus. After six

additional commands, Canopus was again acquired.

The second attempt to perform the midcourse maneuver, which occurred

on December 5, was entirely successful. At the time of the maneuver, the space-

craft was at a geocentric distance of 2 033 776 km (1 264 000 statute miles) and

was traveling at an inertial speed of 3.15 km/sec (1.96 statute miles/see) relative

to the Earth. Both Woomera DSS 41 and Pioneer DSS 11 tracked the spacecraft

during the maneuver, after Pioneer DSS 11 had sent the command to initiate the

maneuver sequence (DC-27). The roll-position error signal and roll rate were

approximately zero when the command was sent, providing optimum conditions

for the roll channel at the start of the maneuver. (Canopus acquisition was main-

tained until about 1 hr after DC-27 was sent.) The ground-commanded (QCI-1,

-2, and -3) and actual (estimated) values of the maneuver were as follows:

141



MARINER-MARS 1964

0

p

Commanded pitch turn, deg ...............

Actual pitch turn, deg .....................

Commanded roll turn, deg .................

Actual roll turn, deg .......................

Commanded velocity increase, 4 m/sec (ft/sec).

--39.16

-- 39.47

156.08156.71

16.98 (55.7)

Actual velocity increase, m/sec (ft/sec) ...... 17.34 (56.9)

As planned, the angle of flight was changed less than 1/4 o. The maneuver moved

the trajectory from a 242 960-km (151 000-statute-mile) miss distance on the

wrong side of the planet Mars (ahead of the planet and above its equator) to an

approximately 9650-km (6000-statute-mile) miss distance on the side of the planet

which satisfied all science and engineering subsystem constraints (behind the

planet and above its south pole). (The accuracy of the maneuver can be demon-

strated by comparing the difference between the commanded miss distance and

the actual miss distance at encounter to the miss difference resulting from throw-

ing a baseball from Los Angeles to New York, aiming for home plate and hitting

second base.)

Postmidcourse-maneuver (interplanetary) cruise

Return of the spacecraft to a cruise configuration was uneventful. Sun re-

acquisition was accomplished, and roll search was initiated. The first acquirable

object to pass within the Canopus sensor field of view, the star Gamma Velorum,

was acquired. One roll-override ground command (DC-21) was sufficient to

reacquire the star Canopus.

After the reestablishment of a cruise-mode configuration, the spacecraft

entered a cruise phase which lasted until the start of the Mars encounter sequence

on July 14, except for the period of the early science cover deployment exercise on

February 11. At the time of reestablishment of the cruise mode, the spacecraft was

traveling primarily under the gravitational influence of the Sun in an ellipse with

the Sun as the focus. The Mariner IV elliptical interplanetary orbit was inclined

0°TaA ' of arc to the ecliptic, with perihelion (point in orbit nearest to the Sun)

near launch and aphelion (point in orbit farthest from the Sun) near Mars atencounter. The Earth-probe and Sun-probe distances are shown in figures 4-5

and 4-6, respectively, as functions of time from launch.

During the early portion of the cruise, the heliocentric velocity of the space-

craft was greater than that of the Earth; thus, the spacecraft led the Earth in

4As determined by the commanded midcourse motor-burn time of 20.06 sec.

142




"o

xE

I-u)

Q

w

rrO.

I

w

300

200

I00

Jo_

ENCOUNTER --

//

/

3OO

2011

IOO

_J

I

ENCOUNTER--,

.IJ

I00 200 300 0 I00 200

TIME FROM LAUNCH, doys TIME FROM LAUNCH, doy$

\

3oo

FIGURE 4-5.--Earth-probe distance as a function FIGURE 4-6.--Sun-probe distance as a 1unction

of time from launch, of time from launch.

orbital rotation around the Sun. Their relative positions are illustrated in figure

4-7, which contains a heliocentric plan view of the orbits of Earth and Mariner IV

during the first 35 days of flight. Slowly, however, the spacecraft began to move

out toward the orbit of Mars with decreasing heliocentric speed. On February 28,1965, the Earth finally passed the spacecraft in its orbital motion around the Sun.

At that time, the spacecraft was at a distance of 19 million km (11.8 million

statute miles) behind the Earth and 321 800 km (200 000 statute miles) above the

ecliptic plane. Throughout the rest of the flight, the Earth increased its lead in

orbital rotation about the Sun. The trajectory of Mariner IV in relation to Earth

and Mars for the period from launch to encounter is shown in figure 4-8.

Occurring during the postmidcourseomaneuver cruise phase were the exe-

cution of six CC&S master timer (MT) commands, the transmission and execution

of two ground commands to modify the cruise configuration of the spacecraft, and

two failures in the science subsystem (the ionization chamber and the plasma

probe, discussed later in this chapter and in chapter 7). The ground commands

sent to the spacecraft were as follows:

1. Power amplifier switch command (DC-7). As the distance from Earth

increased, the radio output of the spacecraft had to be increased to maintain con-

tact with the Earth. As part of the radio flight plan formulated prior to the

Mariner IV launch, a DC-7 command was transmitted to the spacecraft on

143



MARINER-MARS 1964

SUN

SCALE:

MILLIONS OF KILOMETERS

o IO 20

EARTH ORBIT_

3C

25

2O

15

I

5 5

//

//

//

I

35 doys

MARINER /]Z"

ORBIT

3O

25

_.0

I0

FIGURE 4-7.--Heliocentric plan view of Mariner IV trajectory during first 35 days of flight.

144




#MARS ORBIT

EARTH ORBITAT LAUNCH

NOVEMBER 28_ 1964

SUN

_ MAR/NER .23E

/ t _ TRAJECTORY

Oda,

/ 228

I / /

180 \

\90 150

[20 \

\MARS AT ENCOUNTER \

JULY 15_ 1965 __1

6O

FIGURE 4-8.--General relationship of Earth, Mars, and Mariner IV positions from

launch to encounter.

145



MARINER-MARS1964

December 13 to switch the radio from the cavity amplifier to the longer life and

slightly more powerful traveling-wave-tube amplifier. As expected, a slight in-

crease in spacecraft temperature due to the increased power required by thetraveling wave tube was noted.

2. Canopus sensor gate override command (DC-15). Because of the previous

problems in maintaining attitude stabilization in normal roll control prior to the

midcourse maneuver, and because of a subsequent loss of Canopus acquisition and

a reacquisition on Gamma Velorum on December 7, questions were raised con-

cerning the capability of the spacecraft to remain attitude stabilized. It was

decided to reacquire Canopus with a DC-21 command (roll override) and then

send a DC-15, which would disable (desensitize) the Canopus sensor brightness

gates used for acquisition and prevent the initiation of roll search due to observed

high brightness gate violations. These commands were sent to Mariner IV on

December 17, and, although a significant number of roll transients (approxi-

mately 40) were observed during the remainder of the mission, roll acquisition to

the star Canopus was maintained with no problems occurring.

The first CC&S command was the bit-rate switch event (MT-6) which

occurred on January 3 as expected. The data encoder responded normally,

switching from a rate of 33V3 bits/sec to 8½ bits/sec to permit the long-range

communications required for the mission. The altered bit rate was maintained for

the duration of the mission. Coincident with the rate change, a 1-dB received

signal level change was observed at Woomera DSS 41. This was found to be apeculiarity of the L- to S-band conversion system at that station and was not

associated with the spacecraft.

The first Canopus sensor cone angle update command (MT-1) was issued by

the CC&S on February 27. The cone angle was changed from the preset position

of 100.2 ° to 95.7 °. A skip in the data encoder commutation cycle which occurred

at the time of this event still remains unexplained.

As the distance from Earth increased, the received signal strength from the

spacecraft low-gain omnidirectional antenna had reached levels approaching

telemetry thresholds. A marginal received signal condition was anticipated at the

time prior to the programed CC&S switchover of the spacecraft transmitter to the

spacecraft high-gain antenna. The switchover command (MT-5) was issued on

March 5 after the Earth had entered the narrow beam of the high-gain antenna.

The received signal strength then increased about 15 dB. The spacecraft continued

to receive on the low-gain antenna.

146



,J


The second Canopus sensor cone angle update command (MT-2) was

observed in the telemetry data on April 2. The new cone angle was 91.1 ° . An

apparent decrease in the average ground-received carrier power of about 1.5 dB

was indicated and confirmed by three deep space stations during the following 2-

week interval. This behavior is still unexplained, although there is considerable

doubt that it was in any way connected with the MT-2 event. By April 14,

Mariner IV had equaled the U.S. long-distance communications record, just

under 86.9 million km (54 million statute miles), set by Mariner II.

On May 7, the third Canopus sensor cone angle update command (MT-3)

was observed in the data. In view of the previous anomalies which occurred during

the MT-1 and MT-2 events, a thorough analysis of the data was performed for

MT-3. All data indicated that the cone angle update to 86.5 ° was normal in

every respect.

The fourth Canopus sensor cone angle update to 82 ° by MT-4 on June 14

was also completely normal.

A class 2 solar flare occurred on February 5. Solar flares, which are brilliant

eruptions of hydrogen gas generally observed in the vicinity of the large irregular

sunspots, last from a few minutes to an hour or more. (During the mission, a total

of six class 1 and six class 2 flares occurred.) The class 2 flares generally last

approximately 35 min. The particle experiments on Mariner IV detected the

high-energy particles from this event, as well as a rate increase over the succeeding3 hours. Class 2 flares were also detected on April 11 and 16. Increases in magnetic

plasma activity detected by the spacecraft instruments on April 16 were possibly

caused by a solar storm resulting from the flare of April 11. No solar storm activity

was detected that might have been caused by the flare of April 16. A solar flare

detected by the spacecraft instruments on May 25 was not detected on Earth,

possibly because Earth-based observers could no longer see the same part of the

Sun as the spacecraft.

Science-coverdeployment exercise

In order to preclude the possibility that particles dislodged by science-cover

deployment at encounter would cause a loss of Canopus acquisition, it was decided

to deploy the science cover earlier than planned. On February 11 with 12 ground

commands (DC's) transmitted to the spacecraft, the deployment exercise was

performed to: (1) deploy the science cover; (2) preposition the scan platform to the

147



MARINER-MARS1964

optimum encounter position to permit useful television data should a scan plat-

form failure occur before encounter; and (3) turn off the battery charger and

enable the boost mode. All commands were received and executed normally, with

only one anomaly observed during the exercise. Command lock was lost after an

apparently normal command lockup, but was regained approximately 6 min later.

There was no explanation for this loss. Telemetry confirmed that the science cover

was deployed, the scan platform was prepositioned to within 0.7 ° of the optimum

position, and all systems were functioning without any apparent degradation.

Only about 44 min of cruise-science data were lost. The spacecraft was returned to

a cruise configuration without difficulty.

Encounter

Mariner IV approached Mars along the trailing edge and from inside the

orbit of the planet. The near encounter of the Mariner IV and Mars orbits is

illustrated in figure 4-9. The encounter sequence was initiated on July 14 with the

transmission of a command to turn on the encounter-science instruments (DC-25).

A deck skip in the data encoder observed coincident with the command execution

was not totally unexpected. Spacecraft response to the command was normal:

non-real-time power turned on, video-storage-subsystem 2400-hertz power turned

on, scan platform went into a normal search, and power levels and temperatures

increased as predicted. The CC&S turn-on of encounter science (MT-7 event)

was observed in the telemetry data, but, inasmuch as it had been preempted by

DC-25, it had no effect upon spacecraft performance.

SHADOW MARINER TRAJECTORY

ZONE OCCULTATION ZONE RELATIVE TO MARS

L

I [ I / II /I (

\-/ ii I

i / i1 / / /J I t z z

f t / CLOSEST i ! ! i

•,q..- I / APPROACH /

; _" _ _ / Ohr/ --I/2 hr/ --Ihr / --I i/2 hr / MARS ORBIT

A * --------__ _, _ A ' /

;T " "' "'i-it?NDTV / START Tv

f?' SENSOR"SEES"MARS, MARINER TRAJECTORY

(u TV READY RELATIVE TO MARS

FmURE 4 9.--Mariner IV and Mars orbits near encounter.

148




A command to inhibit the scan platform (DC-24) was sent to the spacecraft

for a predicted scan platform position of 178.9". The data indicated that the scan

platform had positioned to 178.45 ° , well within tolerances.

Beginning approximately 51_ hours after DC-25 was issued and continuing

through encounter, anomalous unexplained indications were noted in the mag-

netometer X-axis data which appeared to be cyclic with a period of approximately

16 min. This period corresponded to coincidences between the real-time and non-

real-time portions of the data automation system which count at different rates.

The anomaly had no effect upon the quality of the planetary data gathered.

Had the scan platform been allowed to search for and track the planet auto-

matically, acquisition of the planet by the scan sensor would have initiated a

switch to data mode 3 (all science data) several hours before the recording se-quence. With the scan platform prepositioned, however, the expected time of

wide-angle acquisition was so near the beginning of the television recording

sequence that insufficient time was available for television performance analysis

prior to the latest time to send commands. Therefore, a commanded switch to A,,ill

science data (DC-3) was included in the encounter sequence and was transmitted

approximately 7_/2 hours after the start of the sequence. Television sequencing

was normal.

Wide-angle acquisition occurred approximately 11_ hours later. No effect

was noted, other than the indication in the science telemetry data, because the

scan platform had been inhibited by DC-24 and the switch to data mode 3 had

been accomplished by DC-3. The backup command (DC-16) for the initiation of

the recording sequence was transmitted on the following day, July 15. It was

anticipated that this command would arrive at the spacecraft after two pictures

had been recorded by the video storage subsystem if normal-angle acquisition by

the sensors on the spacecraft had initiated the sequence.

Narrow-angle acquisition was noted in the data. The narrow-angle Mars gate

had sensed the planet at about 43.5 min before closest approach or at approxi-

mately 18 200 km (11 313 statute miles) from the center of the planet. The tape

recorder started recording television data 1 min 12 sec later. A ground command

to back up the termination of the recording sequence (DC-26) was initiated, and

commands to turn cruise science back on (DC-2's) were then sent every 5 min

until six of these commands had been transmitted. After a picture-taking sequence

which lasted 25 min 12 sec, at a distance of 13 270 km (8248 statute miles) from

the center of the planet, the scan platform field of view moved permanently off the

149



MARINER-MARS 1964

planet because of the angular movement of the spacecraft in its hyperbolic orbit

about Mars.

Although the recording sequence appeared to start normally, two anomalousend-of-tape events were observed. It was concluded that these could have been

triggered by noise in the system or could have resulted from some normal transient

in the television sequencing. (During the playback of the stored data, it was

verified that the recording sequence was normal because a failure would have

resulted in fewer than 21 pictures. In other words, only a normal sequence would

have yielded the 21 pictures plus the 22 lines of the 22d picture which were found

to be present on the tape.)

Tracking data gathered and analyzed during the encounter sequence indi-

cated that the Mars-encounter trajectory as predicted during the postmidcourse-

maneuver cruise portion of the flight was somewhat erroneous. Orbit computa-

tions made as early as 1 hour before encounter revealed that the actual magnitude

of the parameter B would not be the approximately 12 068 km (7500 statute miles)

from the center of Mars hoped to be attained by the performance of the midcourse

maneuver. (This discrepancy was previously discussed in ch. 2.)

Postencounter cruise

The command to terminate the recording sequence (DC-26) was observed in

the telemetry data on July 15, and confirmation was received that cruise science

was on and appeared normal. All engineering and science subsystems were per-forming as before encounter. The remaining CC&S event to turn off the encounter

science (MT-8 event) occurred several hours later, but, since it had been pre-

empted by DC-26, this command had no effect upon spacecraft performance.

Approximately 1 _ hours after DC-26 was observed in the data, and approxi-

mately 1 1,4 hours after closest approach to Mars (which occurred at 01:00:58

GMT on July 15), the spacecraft RF signal was lost as Mariner IV entered the

Earth-occultation region behind Mars (i.e., that region where the spacecraft

passed behind Mars as viewed from the Earth, as shown in fig. 4-9). When the

spacecraft entered the occultation region, it was known that its radio signals would

pass obliquely through the atmosphere of the planet and would be attenuated and

bent, just as a stick appears to be bent when placed in water. Thus, by measuring

and then analyzing the changes in the characteristics (frequency and strength) of

the radio signals, it was hoped to learn more about the composition, density, and

scale height of the Mars atmosphere.

150




All four deep space stations then tracking Mariner IV were able to obtain

usable data. The occultation occurred about 8 min later than the time estimated

prior to closest approach. Less than 1 hour later, the spacecraft left the occultation

region and the spacecraft RF signal was reacquired. Analysis of the telemetry data

subsequent to the exit indicated that no apparent changes in the state of the space-

craft had taken place during the occultation and that performance was normal.

As Mariner IV left the vicinity of Mars, the gravitational pull of the planet

altered the heliocentric orbit of the spacecraft to such an extent that the perihelion

distance changed from 148 322 648 to 166 774 861 km (92 183 125 to 103 651 250statute miles). Because the spacecraft passed underneath the planet, a heliocentric-

orbit-plane change of 2.67 ° occurred between the preencounter and postencounter

orbits. As was previously illustrated in figure 2-16, the postencounter orbit does

not intersect the Earth's orbit. The minimum distance between the orbits, which

represents the smallest possible closest approach distance between Earth and the

spacecraft, equals 16 335 373 km (10 152 500 statute miles). The spacecraft will

-**-: close encounter '":_" Ea .... In September 1 n,.7 a:o.,.... n a ...... _*_- " . ,,,, at a u .... nce of approxi-

mately 47 264 375 km (29 375 000 statute miles).

Televisionpicture playback

Postencounter cruise was terminated, and television picture playback was

initiated by the CC&S event MT-9 on July 15, approximately 81_ hours after the

spacecraft left the occultation region. The switch to data mode 4 (stored science

data playback) was confirmed. An anomaly was noted with MT-9 when two

events were registered in an event register rather than the expected single event.

Coincident with MT 9 was a CC&S cyclic command. It is believed that the two

events were sufficiently separated in time that each (rather than just one) caused

an increment in the event register.

After approximately 68 min of engineering data, the first data mode 4 data

were observed in the telemetry. Playback proceeded in a normal manner through

the termination of the playback on August 3. During this period, the data from

each of the pictures were transmitted to Earth twice, with the exception of the last

few lines of the 22d picture. Data recovery rate was quite high, and the perform-

ance of the spacecraft was normal throughout the playback sequence.

The first complete transmission of the stored data ended.July 24, and a second

transmission began automatically. This second playback was necessary to insure

that any data missing from the first transmission might be recovered and to pro-

151



MARINER-MARS1964

vide duplication of the remainder of the data for comparison purposes. Spot checks

of the data from both transmissions confirmed that the data were repeatable, so it

was decided to return the spacecraft to a cruise mode at the end of the second play-

back. The only anomaly occurred during the second playback when two events

were recorded in a data encoder register at the first end-of-tape signal. The track

change was normal, however, and the extra end-of-tape event was attributed to

dirt or foreign material on the end-of-tape foil.

Shortly before the playback phase was terminated, a second recording

sequence was considered to provide information on the behavior of the television

electronics for uniformly black pictures at each of the various gain settings. With

this possibility in mind, it was decided to terminate the playback in such a manner

that some portion of the Mars data would not be erased should another recording

sequence be performed. The playback was terminated in the 18th line of the 22dpicture, thus protecting all the complete pictures on the second track.

Since an electrical short in the 2400-hertz power supply for the video storage

subsystem would result in a "catastrophic" failure of the spacecraft power sub-

system, it was decided to send a DC-28 command on August 3 to turn off this

power supply. Data mode 1 (engineering data) was confirmed. Because this

command also disabled the boost mode and turned on the battery charger, it was

followed by a DC-26 command to turn off the battery charger and reenable the

boost mode. Then, a DC-2 command was sent to the spacecraft to inhibit the play-

back motor, switch to data mode 2 (engineering and science data), and turn on

cruise science.

Final 8 weeks

After return to a cruise configuration, all cruise-science instruments were

verified to be operating as before encounter, with all engineering subsystems per-

forming normally. The spacecraft continued to perform well throughout the final

8 weeks of the mission. One anomaly was noted: the cosmic dust detector instru-

ment began to return apparently abnormal data sporadically. Problems occurred

which caused the postponement of command sequences twice, but investigation

verified that in each case the problems were ground based rather than associatedwith the spacecraft.

Various tests and operations were conducted during the remainder of the

postencounter cruise phase, in which it was possible to maintain contiguous telem-

etry coverage. Among these were inflight calibrations of the spacecraft radio and

command subsystems, which were considered necessary to any attempt to re-

152




acquire the spacecraft in 1967; and the "black space" encounter sequence to

provide television calibration data, which was considered necessary to provide an

insight into the possible fogging of Mars pictures resulting from some unknown

phenomenon in the television electronics. On August 30, the command to turn on

the encounter science (DC-25) was transmitted, and the scan platform search

mode was initiated. A DC 3 transferred the data encoder on the spacecraft to data

mode 3 (science data), a DC-24 inhibited the scan platform motion at 148.43 °,

and a DC-16 actuated the narrow-angle acquisition logic and initiated the record-

ing sequence. The following day, a DC-2 was transmitted to transfer the data

encoder to data mode 2 (engineering and science data), after which the ends of the

second and first tracks of the tape were indicated in the data. The data encoder

was transferred to data mode 2 by the data automation system after the tape waspositioned at the end of the second track (the 22d picture). The science subsystem

power and the scan-platform and video-storage-subsystem 400-hertz power were

turned off by a DC-26, and the video-storage-subsystem tape recorder changed

tracks after receipt of a DC-22.

A total of 101_ pictures was recorded during this sequence. After the space-

craft video-storage-subsystem track change, the first picture was ready for play-

back. Playback was initiated by ground command (DC-4), and data mode 1 data

(engineering data) appeared in the telemetry approximately one-half hour later.

Data mode 4 data (science data) were observed shortly afterward, and playback of

the first picture began. The stored data were played back until the first five pic-

tures, which included all the television gain settings, had been completed on

September 2. The sequence to return the spacecraft to its cruise configuration was

then begun. A DC-28, a DC-26, and a DC-2 were transmitted, and the first data

mode 2 data (engineering and science data) were observed in the telemetry,

indicating the completion of the television calibration sequence. The engineering

and science data indicated that all subsystems were performing as they had prior

to the initiation of the sequence.

Among the flight operations conducted during the final 8 weeks, a maneuver-

inhibit command (DC-13) and a minimum-turn and motor-burn command

(QCI-1, -2, and -3) were transmitted to Mariner IV on August 26 to protect

against an inadvertent midcourse maneuver which could take place if certain

specific failures of the CC&S were to occur prior to the 1967 reacquisition period.

Telemetry indicated that all commands were successfully received.

On August 27, a ground command to update the Canopus sensor cone angle

to the first optional position (DC-17) was transmitted. All previous cone angle up-

153



le

¢

MARINER-MARS 1964B

dates were commanded automatically by the CC&S, but no master timer (MT)

event was provided for more than the four updates previously executed during the

postmidcourse-maneuver cruise phase. The new cone angle setting was 77.8 °.

On August 6 and again on September 3, the spacecraft-received carrier power

with command modulation applied was below the worst-case command threshold

for 10-kW transmitters at the prime deep space stations. The level remained low

thereafter. For a significant portion of the period after September 7, the space-

craft-received carrier power without command modulation applied was suffi-

ciently close to the absolute RF threshold to result in a drop in the ground-

received carrier power below the absolute telemetry threshold each time the space-

craft was "locked up" two-way, causing a loss of all data. To prevent this, two-way

tracking was forfeited, and the spacecraft was locked up only periodically for short

durations to prevent the spacecraft from transferring to receive by means of the

high-gain antenna.

On September 3, the average ground-received carrier power had reached the

worst-case telemetry threshold and had then dropped steadily as the spacecraft

high-gain antenna pointing error increased. On October 1, as predicted, the

carrier level was fast approaching the absolute telemetry threshold. On that day, a

command to transmit and receive by means of the low-gain antenna (DC-12) was

transmitted from the 100-kW transmitter at Venus DSS 13. This transfer pre-

cluded the cycling of the receiver between antennas by the radio subsystem logic

circuitry, thus permitting access from the 100-kW transmitter if desired. TheMariner IV spacecraft was thus in the proper configuration for a projected

attempt to reacquire the spacecraft in 1967. Telemetry from Mariner IV was lost

at 22:05:07 GMT on October 1, 1965, at approximately 309.2 million km (192.2

million statute miles) from the Earth, marking the end of the mission.

Futureoperationsfor Mariner IV

The DSIF had determined that, if the Mariner IV transmitter was returned

to the omnidirectional low-gain antenna, it would be possible to track Mariner IV

as an RF source completely around its orbit of the Sun. Invaluable celestial-

mechanics data for reducing the present uncertainty in the astronomical unit

value and the ephemerides of the Earth and Mars could thus be obtained.

During the latter half of 1967, the spacecraft will again be within telemetry

reception range of the Earth. If the transmitter is still operating at that time and if

tracking schedules permit, it may be possible to recover additional cruise-science

information.

154




MISSION PLANNING AND DECISION PHILOSOPHIES

Launch

For the Mariner-Mars 1964 missions, two launch pads were used to obtain an

acceptable launch probability of success. The dual-pad requirement reduced the

hazards to mission success from "catastrophic" damage to the pad, either by

missions immediately preceding the first Mariner-Mars 1964 spacecraft launch or

by the first launch itself. This was important since the nominal launch period was

only about 27 days and since 21 days are normally required to refurbish a pad and

erect a second launch vehicle. Two pads thus permitted the immediate shift of

launch operations to the second pad if the vehicle on the first pad encountered

problems or was launched.

Primary operations during the launch phase were directed toward the ac-

complishment of the primary objectives: injection of the spacecraft successfully on

a Mars trajectory, deployment of the solar panels, and acquisition of the Sun.

Without the accomplishment of these objectives, the mission would be an immedi-

ate failure. (During the Mariner III mission, the shroud failure precluded all

three of the primary launch objectives and the spacecraft failed upon depletion of

the battery.) Secondary objectives during this phase were: turn-on of cruise science,

increase of the radio transmitted power (RF power up), removal of CC&S relay

holding current, turnoff of video-storage-subsystem launch mode, attainment ofmagnetometer calibration roll, acquisition of Canopus, and assessment of the

capability of the spacecraft to perform the midcourse maneuver. Each of the

objectives, with the exception of injection (over which the Spacecraft Performance

Analysis and Command (SPAC) Team, discussed in ch. 6, had no control), had to

be examined; the procedure by which it could be verified had to be determined;

and appropriate plans for any corrective action had to be formulated.

Before launch, a standard sequence of events during Canopus acquisition was

established, since it was anticipated that star identification might pose a serious

problem and because the celestial geometry near launch made it probable that the

first star acquired would not be Canopus. The adopted sequence allowed the

spacecraft to acquire any object which fulfilled the Canopus sensor brightness

logic intensity requirements and to become roll stabilized to that star. All data

which might provide evidence as to the roll orientation of the spacecraft were

then gathered and evaluated. Based on this evaluation, a recommendation for

command action could be formulated and then implemented during the next

Pioneer DSS 11 pass.

155



MARINER-MARS 1964

.o

Midcourse Maneuver

The midcourse maneuver was perhaps the most demanding phase of the

mission, for the SPAC and FPAC Teams (discussed in ch. 6) in particular.

Normally, the accuracy of injection is such that a midcourse correction is required

for a successful mission, and the required correction parameters must be furnished

by ground command. Careful monitoring of the performance of the spacecraft

during all phases of the midcourse maneuver is required since a nonstandard mid-

course maneuver may result in, at best, an unacceptable planetary-miss distance

and, at worst, a total spacecraft failure.

A capability existed for termination of the inidcourse maneuver by ground

command before rocket motor ignition, should an anomaly jeopardizing the suc-

cessful execution of a maneuver arise. Barring gross failure, a terminated sequence

would have no significant effect since the maneuver could be rescheduled and

performed successfully regardless of the number of previous aborts. For example,

the first attempt to correct the trajectory of Mariner IV was terminated because of

an unexpected loss of roll attitude shortly after the initiation of the maneuver

sequence. However, the following day (December 5) the maneuver was completed

without incident. Thus, the operations philosophy for the midcourse-maneuver

phase was determined by the abort capability inherent in the spacecraft design.

For the Mariner-Mars 1964 missions, it was fully expected that the maneuver

would be executed successfully without deviation from the predetermined plan.

Every consideration had been given to insure that the spacecraft would enter the

midcourse sequence in the proper state.

Postmidcourse-Maneuver (Interplanetary) Cruise

The nominal flight sequence of events was designed to be totally automatic

and to require no action from the ground for a normal flight. Such a design

relieves the mission from any dependence upon the ground-to-spacecraft com-

munications link, provided the CC&S on the spacecraft operates correctly. The

ground command function is necessary to the success of the mission only in the

event that the CC&S fails. This was the "cornerstone" of the functional redun-

dancy philosophy employed in the Mariner-Mars 1964 spacecraft design. A strict

reliance on the functional branch of the spacecraft may, however, lead to a flight

sequence less than optimum. Thus, during the Mariner IV mission it was decided

to depart from the nominal sequence at those times when it was advantageous to

do so. The major departures are discussed in this section.

156



t FLIGHTHISTORYAND SPACEVEHICLEPERFORMANCE

Early in the Mariner IV mission, a recommendation was made and approved

to switch by ground command the spacecraft transmitter from the cavity amplifier

to the longer life, slightly more powerful traveling-wave-tube amplifier. This

switch represented the implementation of a plan formulated prior to launch. As

the distance from the Earth increased, the radio output of the spacecraft had to be

increased to mli_tain contact with the Earth. A DC-7 command was transmitted

to Mariner IV on December 13, and the transfer was successfully completed.

The traveling-wave-tube amplifier has a predicted lifetime far in excess of the

minimum Mariner-Mars 1964 requirements, and its known failure modes lead to

immediate "catastrophic" failure. The cavity amplifier, on the other hand, shows

definite aging effects which limit its useful lifetime to approximately the length of

the Mariner IV mission. Use of this amplifier during launch was necessary since

the traveling-wave-tube amplifier did not have the low power required for that

phase of the mission. Had the traveling-wave-tube amplifier failed at some later

time during the mission, the cavity amplifier would still have been available for the

remainder of the flight.

Another factor which contributed to the decision was the capability of the

spacecraft to transfer automatically from one power amplifier to the other if the

amplifier output stopped below a fixed point. Since the most likely failure mode of

the traveling-wave-tube amplifier is "catastrophic" failure, automatic switching

to the cavity amplifier would have taken place had the former failed. The most

likely failure mode of the cavity amplifier is gradual degradation; therefore, ifcommand capability had been lost, the spacecraft could have reached a state

where the output was so degraded that useful data could not be returned and yet

not sufficiently degraded to cause automatic switching to the traveling-wave-tube

amplifier.

After the transfer, the traveling-wave-tube amplifier operated as expected,

with one exception: from December 22 to 31, the helix current of the amplifier was

varying more than expected and generally tended to increase. This trend was

discovered to be a characteristic of the amplifiers used in space applications and

was considered normal.

Because of roll transients which had occurred, questions were raised concern-

ing the capability of the spacecraft to remain attitude stabilized in normal roll

control. When Canopus acquisition was lost and the star Gamma Velorum was

acquired on December 7, it was decided to allow the sensor to remain acquired to

Gamma Velorum by means of the high-gain antenna and to formulate a plan of

157



MARINER-MARS 1964¢

action. (Roll orientation is critical during cruise only when transmitting or receiv-

ing; loss of acquisition then is accompanied by a loss of downlink communications

from the spacecraft.)

Transmission of a DC-15 command to the spacecraft would mean the

removal of the Canopus sensor brightness gates used for acquisition and would

thus prevent the initiation of roll search due to the observed high brightness gate

violations (found to be caused by transients resulting from dust particles illumi-

nated by the Sun). Since the spacecraft would respond slowly compared with the

speed of the violations, the sensor would remain oriented properly toward Cano-

pus regardless of the transient brightness seen. The only remaining mechanism for

this type of loss of acquisition with the gates removed would be that the sensor

would follow another object besides Canopus if the object were brighter than

Canopus and moving slowly enough through the sensor field of view to allow

spacecraft roll response.

The major drawback to this action would be that, since the gyro control unit

was controlled through the gate logic, there could be no gyro turn-on by ground

command should Canopus acquisition be lost. If the brightness gates were dis-

abled and the spacecraft was receiving by the high-gain antenna, it would require

662/_ to 1331/_ hours (the time required for the spacecraft logic to switch the

receiver back to the low-gain antenna automatically, thus restoring command

capability) to reacquire, should Canopus acquisition be lost. If Sun acquisition

were also lost, reacquisition within the required time might not be possible. How-ever, since the probability for such a double loss of acquisition was quite low, the

risk was accepted.

Therefore, a DC-15 was transmitted on December 17, preceded by a DC-21

(roll override) to reacquire Canopus. Although a significant number of roll tran-

sients (approximately 40) were observed in the telemetry during the rest of the

mission, none caused the loss of Canopus acquisition. It was decided to remain in

the DC-15 condition when the transmitter was switched from the low-gain

antenna to the high-gain antenna because of the continuing occurrence of roll

transients. A more severe problem was presented by an apparent requirement to

switch the receiver to the high-gain antenna several weeks later. However, analysis

of the spacecraft RF characteristics in the transmit high-gain, receive low-gain

mode showed that an RF interference effect between the signals from the two

antennas into the receiver produced the unexpected advantage that command

capability could be maintained from the three prime tracking stations during the

remainder of the mission for all but a relatively brief period. Thus, the major

158



\


disadvantage associated with DC-I 5 was eliminated, and a decision to maintain

the DC-15 condition for the remainder of the mission was made.

Shortly after the switch of the spacecraft transmitter from the low-gain to the

high-gain antenna, it was found that the spacecraft-received signal strength and

the ground-received signal strength were varying in a manner which indicated

that some amount of RF interference was being experienced. This was felt to be a

problem only when the Earth was nearly alined with the boresight of the high-gain

antenna and the spacecraft was either transmitting or receiving by means of the

low-gain antenna. It was known that RF leakage by the high-gain antenna was

approximately of the same order of magnitude as the normal signal by the low-

gain antenna. The path length from the ground antenna to each of the spacecraft

antennas then determined the phase relationship of the two interfering signals,

causing either constructive or destructive interference.

An early switch of the spacecraft transmitter to the high-gain antenna had

been considered. However, analysis of the signal characteristics and other data

showed that the first null in the interference pattern (which would last from 3 to 5

days) would still yield signal strengths above the absolute telemetry threshold, and

that the telemetry data received would be of sufficient quality to allow normal

spacecraft data analysis. After the null, the received signal strength from the

spacecraft would begin to increase as the changing relative path length from the

two antennas to the Earth reached a point where the interfering signals began tocome into phase with each other. The CC&S-controlled switch of the spacecraft

to the high-gain antenna was scheduled for just after the time of the predicted peak

signal strength. Thus, it was decided to take no command action to effect an

early switch.

The CC&S-controlled transfer occurred as expected on March 5. Coincident

with the switch of the transmitter was a 15-dB increase in spacecraft-received

signal strength. Subsequent analysis indicated that the change of state of the cir-

culator switch controlling the antenna switch also changed the phase relationship

of the interfering signals on the Earth-to-spacecraft link, effectively changing the

position of the interferometer pattern.

Originally it had been predicted that command capability by means of the

spacecraft low-gain antenna could be maintained through April 8. Predictions,

however, included a very deep null in the interference pattern from March 24 to

April 1 for the 10-kW transmitters and from March 26 through 30 for the 100-kW

transmitter (Venus DSS 13). This null would have precluded command action

iust prior to loss of command capability by means of the low-gain antenna. Except

159



1

MARINER-MARS 1964

for the command blackout period during the predicted null, command action

would be possible throughout the postencounter period by using the 100-kW

transmitter. Thus, the maximum loss of both uplink and downlink communica-

tions would be the period between the Goldstone Deep Space Communication

Complex passes. However, if the spacecraft was switched to receive by means of

the high-gain antenna, a loss of Canopus acquisition could result in a command

blackout of up to 1331/_ hours (with DC-15 in effect), the inaximum time until the

spacecraft automatically switched its receiver back to the low-gain antenna in the

event of a loss of uplink communications.

It was first decided to construct the spacecraft in such a manner that the

transfer of the spacecraft receiver to the high-gain antenna would take place auto-

matically on March 26. To accomplish this, no uplink RF lock could be estab-

lished until two consecutive CC&S cyclic pulses occurred. At the second pulse, the

receiver would transfer from one antenna to the other and, at each subsequent

pulse, would transfer again as long as the uplink RF lock was withheld. Thus, to

abort the transfer to the high-gain antenna, all that would be required was to

establish uplink RF lock with the spacecraft. Just prior to the cyclic pulse which

would have effected the transfer, it became apparent that the decrease in signal

strength received by the spacecraft was less than predicted. As additional telem-

etry became available, it was concluded that the null would probably still leave

a substantial command margin for even the 10-kW transmitter. A two-way RF

lock was established prior to the cyclic pulse, thus inhibiting the transfer of the

spacecraft receiver.

Analysis indicated that the error in the prediction of the spacecraft-received

signal strength was due to the fact that the circulator switch controlling the

antenna transfer had different characteristics in flight than those assumed prior to

launch. Predictions of the command margin for the remainder of the mission

showed a positive margin for all but a brief period during May and June, when

command capability could still be maintained from the 100-kW transmitter.

Therefore, it was decided that the spacecraft should continue to receive signals by

the low-gain antenna throughout the mission, pending a final review as the 10-

kW-transmitter command blackout approached in May.

Although the spacecraft-received carrier power using the 10-kW Earth-based

transmitters dropped as expected in May, it was recommended that the space-

craft still continue to receive signals by means of the low-gain antenna because:

(1) few failures could be hypothesized that were correctable by ground command

160




and that could not wait until the 100-kW transmitter could be used; and (2) the

significant risk of losing Canopus acquisition (since DC-15 was still in effect), and

therefore all command capability, would be present if the transfer were made.

Thus, the 100-kW transmitter would be kept ready for command transmission on

short notice. Periodic command loop lockups with both the 10- and 100-kW

transmitters were made during the remainder of the mission.

Early Science Cover Deployment Exercise

At the recommendation of the Encounter Planning Working (EPW) Group

(discussed later in this section), a science cover deployment exercise was per-

formed early to prepare the spacecraft for encounter. The science cover was

deployed on February 11, 1965, for the following reasons:1. To minimize the possibility of losing lock on the star Canopus because of

slow-moving dust particles being shaken loose from the spacecraft by the inertia

of the spring-operated science cover when deployed by CC&S command MT-7.

2. To reveal possible but unexpected problems with the encounter sequence

in the actual flight environment, and to allow adequate preparations to enhance

the SPAC Team's capabilities for taking prompt corrective action.

3. To permit positioning of the scan platform to increase the probability of

planet acquisition and thus enable useful television data, especially in the event of

a scan system failure.

4. To provide information, such as temperature and power frequency shift,

which could be useful in assessing power subsystem operation at encounter.

5. To minimize the risk of a science cover deployment system failure by

actuating the system earlier in flight.

A series of operational tests was conducted to establish the proper times for com-

mand transmission in order to avoid any potential degradation to the system and

in order to confirm that timing errors and scan platform positioning uncertainties

were well within tolerance.

A command sequence was adopted which included all of a normal encounter

sequence, except the exercise of the video-storage-subsystem tape recorder in

either a record or reproduce mode. This exercise was not pertbrmed since the

risks involved were sufficiently large to compromise any possible value to be gained

by its accomplishment. In order to minimize the remaining risks, the command

sequence included a number of commands in a preliminary sequence designed to

verify the correct operation of the spacecraft as much as possible before sending

161



MARINER-MARS 1964

Ii

any commands that might have irrevocable consequences. The fact that a minor

roll disturbance did occur during the science cover deployment exercise confirmed

the wisdom of the decision to conduct it at the early date.

Encounter

The basic design of the spacecraft provided for a completely automatic

planetary encounter sequence with no ground action required. If the ground

command option were to be available, however, there existed a number of possible

backups to the spacecraft automatic functions and, in addition, a number of

alternate modes which might be superior in a given situation to the normal mode

of operation. In order to provide the best possible sequence of events for the en-

counter, an Encounter Planning Working (EPW) Group was formed in January1965 to operate in parallel with the SPAC Team for the purposes of investigating

all spacecraft encounter-related operations and modes and recommending a

detailed encounter sequence of events. Appropriate SPAC Team, J PL technical

division, and spacecraft subsystem representatives composed the EPW Group,

whose objectives were as follows:

1. To develop an encounter sequence which provided the maximum as-

surance of obtaining useful television and occultation data.

2. To determine the operational mode providing the greatest assurance of

maintaining attitude stabilization throughout encounter, thereby assuring real-

time planetary field data, particle experiment data, and occultation data.

3. To provide for full utilization of backups for all critical functions.

4. To select appropriate alternate modes for any functions that were both

time and functionally critical.

Based upon the findings of the EPW Group, an encounter sequence of events was

formulated by the director of the SPAC Team and was approved.

The basic procedures of the approved sequence were as follows: (1) turn on

the scan platform and at the same time start the television camera shutter operat-

ing; (2) stop the scan platform at the best angle for taking pictures; (3) start the

tape recorder at the proper time so that the pictures would be recorded; and (4)

stop the tape recorder at the proper time so that the pictures would not be erased.

The spacecraft was designed to produce this sequence automatically.

However, it took 25_ min for round-trip communications with the space-

craft; therefore, if the automatic sequence failed, little time would be left to

analyze the situation, send corrective commands, and wait 25_ min to see

162



o

FLIGHTHISTORYAND sPACEVEHICLEPERFORMANCE

whether the corrective commands were effective. Therefore, the optimum en-

counter plan was designed to anticipate possible troubles with enough lead time

for appropriate action. An initial set of commands would be sent to the spacecraft

to start the scan platform and the television shutter and then to stop the platform

at the optimum angle. If these commands were not effective, the automatic

sequence would then act as a backup. If the automatic sequence should also fail,

enough time would remain to recycle the events for another attempt at the

sequence. When Mars came into the field of view of the television lens, a command

would be sent to turn on the tape recorder. This command would serve as a

backup to the automatic sequence. The same procedure would be used to stop the

tape recorder at the proper time.

In preparation for the encounter, tests were conducted involving nominalencounter operations and possible encounter operations with a failure-mode con-

dition. These tests were designed to exercise space flight operations (SFO) Sys-

tem analysis and operations personnel (described in ch. 6) in the performance

of their duties and to develop and maintain the proficiency required for proper

participation during the planetary-encounter portion of the Mariner IV mission.

All such testing was conducted from April 13 to July 12, 1965.

The major problem in encounter operations was the transmission delay

between the Earth and the spacecraft. Approximately 13 min was required from

data transmission at the spacecraft until data presentation to the SPAC Team;

more than 12_ min were required from command initiation at the DSIF until

command execution at the spacecraft. Thus, immediate response to anomalous in-

dications in the data could only reduce the effective reaction time in an emergency

situation to a minimum of 25_ min. Conversely, normal command action as a

part of the sequence of events could be taken only if it were assumed that nothing

abnormal had occurred in the 13 min prior to command initiation and that

nothing abnormal would occur during the 12 _ min before command execution.

Since most commands to be transmitted were time critical or functionally

critical, it was important to prevent accidental loss of command lock and, if such a

loss occurred, to provide for a minimum-time reacquisition of the command loop.Only deep space station personnel could act on the former, but the latter could be

achieved by maintaining the command subcarrier at the slightly offset, minimum-

lockup-time frequency normally used for initial lockup and not for command

transmission. This offset was not large enough significantly to increase the proba-

bility of dropping lock at encounter signal levels. Whereas this provision was not

required during the actual encounter sequence, the flawless operation of the corn-

163



MARINER-MARS 1964

mand detector and decoder in accepting and executing all commands verified the

practicability of this approach for such critical operations.

Only minor departures were made from the recommended encounter se-quence of events during the actual Mariner IV encounter phase.

Postencounter

Planning for the postencounter phase consisted of determining, evaluating,

and selecting those activities which were to be attempted before the communica-

tions threshold with the spacecraft was reached. The activities considered were

those needed to prolong communications with the spacecraft during 1965, to

enhance the reacquisition of the spacecraft in 1967, and to exploit the utility of the

spacecraft as an engineering test instrument subsequent to the achievement of the

primary mission objectives. The basic ground rules used in making decisions

relative to the postencounter activities were that the accepted activity had to: (1)

provide enhancement of the data already received (e.g., the television calibration

sequence); (2) protect the accumulation of a maxinmm amount of deep-space

science data in 1965 (e.g., by updating of the Canopus sensor cone angle, which

would allow Canopus acquisition until 1966); or (3) enhance the possibility of

reacquiring the spacecraft in 1967 (e.g., by using the midcourse-maneuver inhibit

sequence and by switching the spacecraft transmitter to the low-gain antenna).

Special spacecraft tests which did not jeopardize the successful achievement of the

above objectives would also be permissible.Shortly before the playback phase of the mission was terminated, it was

recommended that a second recording sequence be considered. Such a sequence

could provide information on the behavior of the television electronics in provid-

ing uniformly black pictures at each of the various gain settings. With this possi-

bility in mind, it was further recommended that playback be terminated in such a

manner that some portion of the Mars data would not be erased should another

recording sequence be performed. By stopping the tape at some position other

than at the foil marking the end of the tape, from 1 to 10_ pictures could be

protected, since in a recording sequence the tape would run only until the end-of-

tape foil had been crossed twice. It was decided to terminate playback in the 18th

line of the 22d picture, thus protecting all the complete pictures on the second

track.

The calibration sequence was begun on August 21, when the star Altair was

at a proper cone angle to be photographed. While this star, a light source of low

magnitude, would not be resolved by the television subsystem, the effects of the

164



i


light source would be apparent in any pictures, thus providing calibration data

for the television optics as well as for the electronics. A normal two-way RF lock

was established with the spacecraft, command modulation was applied, and the

command detector locked up. All telemetry indications were normal, so the first

command in the sequence, a DC-25 (encounter science on), was transmitted. It

was planned to follow the planetary encounter sequence of events closely with, of

course, the automatic events. However, shortly after transmission, the telemetry

indicated that the carrier power at the spacecraft was fluctuating severely. Several

samples indicated momentary drops of lock by the command detector as the

carrier level at the spacecraft dropped below threshold. Since these momentary

out-of-lock indications appeared to indicate either a DSIF failure or a spacecraft

failure, it was decided to terminate this calibration sequence.

Since the DC-25 command was being executed at the spacecraft, an emer-

gency command sequence had to be used for returning the spacecraft to a

cruise mode. Redundant commands were included to avoid the possibility that

one of the momentary command dropouts would fall just prior to or during the

receipt of any of the commands, thus inhibiting execution of the command. The

spacecraft was, however, returned to cruise with no anomalies occurring.

Although the following day was also acceptable for the exercise using the star

Altair, no exercise was to be scheduled unless assurance could be given that no

failure or problem had occurred in the spacecraft radio or command equipment

and that earlier problems at the deep space stations involved had been resolved.

Since insufficient time remained to give this assurance, it was decided to forgo the

exercise until a full analysis could be made of spacecraft and DSIF performance.

This analysis revealed the problems to have been caused by deep space station

anomalies, and these problems were subsequently resolved.

The second attempt at the calibration sequence occurred on August 30 and

was successful. This try differed from the first only in that no star of sufficient

magnitude was available as a light source. Therefore, the data gathered applied

only to the television electronics and not to the optics. The use of the small 10-kW

transmitter at Echo DSS 12 was possible, since during this period the RF inter-

ferometer effect aided, rather than degraded, communications capability. The

sequence proceeded without incident. A nonstandard time to the first end-of-tape

foil was due to the tape position at the start of the sequence; tile playback of the

encounter data had been terminated just before the foil on the tape signaled the

end of tape, thus allowing the calibration sequence to record over only half the

planetary data, with the video storage subsystem automatically inhibiting after

165



MARINER-MARS1964

the foil had been passed twice. The spacecraft transferred back to data mode 2

(engineering and science data) after the tape was positioned at the end of the

second track (the 22d television picture). The stored data were then played back

until the first five pictures (including all the television gain settings) had been

completed. The spacecraft was then returned to a cruise configuration.

The transmission of a DC-13 (midcourse-maneuver inhibit) to the spacecraft

had been considered prior to planetary encounter. This command would protect

against an inadvertent midcourse maneuver if a failure occurred in the CC&S on

the spacecraft. Since the likelihood of this occurring before encounter was very

small and since the maneuver could have been inhibited by ground command,

even if the failure occurred, it was decided not to send the command prior to the

encounter. After encounter, however, the major portion of the mission objectives

had been completed, and new objectives--to track the spacecraft in its orbit

around the Sun and to attempt to reacquire it in 1967--had been added. There-

fore, it was necessary to condition Mariner IV so that it would be in an optimum

configuration. If an inadvertent midcourse maneuver should occur, the helio-

centric orbit of the spacecraft would be altered. In order to prevent this occur-

rence, it was necessary to send a DC-13 on August 26. As an additional safeguard,

a command (QCI-1, -2, and -3) for minimum roll and pitch turns and for a mini-

mum motor-burn time was also sent at that time.

The spacecraft was receiving on the high-gain antenna, which, because of

its narrow beam, must be pointed directly at the Earth to receive a signal. How-

ever, there was no guarantee that this antenna would be pointing at the Earth in

1967, even if it was assumed that Mariner IV was still locked on Canopus. There-

fore, the spacecraft was switched by a DC-12 on October 1 to receive on the low-

gain omnidirectional antenna in order to improve chances for reacquisition in

1967.

Future Operations for Mariner IV

The second phase of Mariner IV operations will consist of periodic contacts

with the spacecraft by the Deep Space Network (DSN) by using advanced com-

munications techniques. During this period, commands to update the Canopus

cone angle will be sent if and when system capability is available. (Telemetry

reception capability is not expected.) Two-way experiments from which some

Doppler data may be extracted will be attempted by using the 100-kW trans-

mitter at Venus DSS 13. Additional trajectory data may be provided which will

yield further information on the inherent accuracy of DSN and possibly improve



1,


our values of certain astrodynamical constants. Finally, the spacecraft will act as a

far-field calibration device for DSN. The second phase of operations will officially

be terminated when Mariner IV telemetry data can once again be received by a

standard configuration of DSN (expected in 1967). This date depends upon the

operational mode chosen (low-gain omnidirectional versus high-gain antenna

mode) and the DSN performance capabilities available at the time.

The operational plans for the third phase of operations are much more com-

plex than for the first two phases and are still under study. This basic opportunity

to reacquire the spacecraft is unique in that it presents the possibility of exploring a

heretofore unexplored locale in space. Also, very valuable information for space-

craft engineering technology concerning longevity characteristics of spacecraft

may be extracted. Indeed, the continued operation of Mariner IV in 1967 could

well affect our basic approach to system design philosophy. Information on failure

modes and mechanisms could be obtained on a restricted level by discrete exercise

of various redundant provisions inherent in the spacecraft design. Redundancy

exists for many of the critical elements required for success of the 1967 reacquisi-

tion effort. In fact, most of the redundant aspects of Mariner IV have not yet been

utilized. Inherently, however, the greatest advantage of the reacquisition effort

lies in the basic scientific importance of continued sampling of interplanetary

conditions for an extended period during a time of relatively high solar activity.

The preparation and support directly associated with the Mariner IV reacqui-sition effort have been officially designated the Mariner IV project.

MARINER IV SUBSYSTEM PERFORMANCE

Structure and Mechanisms

Because the functions of all the structural and many of the mechanical items

on the spacecraft were passive during and after launch, little telemetry was

transmitted about their performance. Thus, the inflight performance of theseitems, as well as that of the electronic packaging and cable harnessing, could only

be deduced from other flight information. All such information indicated satis-

factory performance.

Telemetry was recovered during the flight on the operation of the following

mechanical devices:

1. Separation-initiated timer and pyrotechnics arming switch. Telemetry

event indications verified that either the separation-initiated timer or the pyro-

167



MARINER-MARS1964

_t

technics arming switch energized the pyrotechnics subsystem. It was not deter-

mined which one armed the subsystem since Agena D/spacecraft-separation

telemetry data were not available for this operation. The first available telemetryfrom the data encoder counter, received about 3 min after spacecraft separation,

indicated normal solar panel deployment and unlatch of the scan platform, both

actuated by the separation-initiated timer.

2. Solar panel deployment springs. Telemetry indicated that each of the

four solar panels had been deployed by the solar panel deployment springs to

within 20 ° of the fully open position. Temperature and power measurements later

confirmed the deployment.

3. Cruise dampers. From the accuracy of the midcourse maneuver, it could

be deduced that the solar panels did deploy and latch in the fully open position

and that any panel excursions during the motor firing were adequately damped.

4. Scan actuator. During the science cover deployment exercise on February

11, the scan actuator functioned normally for 127 min (approximately 11 scan

cycles), the average cycle lasting 11 min 52.9 sec. During the 108 rain of actuator

operation at encounter, which was again normal, the average cycle lasted 11 min

53.2 see. For the 66 min of actuator operation during the aborted picture-taking

sequence of the star Altair on August 21, the average cycle was 11 min 53.8 sec.

The operating period for the completed television calibration sequence on August

30 was 139 min, with an average cycle lasting 11 min 53.5 sec.

5. Scan inhibit switch. The data encoder event register indicated that thescan inhibit switch operated perfectly when the pinpuller latching the scan plat-

form was fired at spacecraft separation. The subsequent successful operation of

the scan actuator verified this indication.

6. Science cover. On February 11, the science cover was unlatched by a

solenoid initiated by the pyrotechnics subsystem. This operation was verified by

a change in data number of the spacecraft identification channel (fig. 4-10) and

by a decrease in scan platform temperature.

Radio

The radio subsystem operated continuously with no evidence of any mal-

function or degradation of performance. However, the following anomalies were

observed:

1. Variation in traveling-wave-tube helix current after the switch to the

traveling-wave-tube amplifier on December 13 (fig. 4-11). This variation was

found to be a characteristic of these components and was therefore considered

normal.

168



,I

" FLIGHT HISTORY AND SPACE VEHICLE PERFORMANCEQ

I00

9O

BO

7O

60

0 40 80

__._SCIENCE COVER

DEPLOYED ON

FEBRUARYII

120 160 200 240 260

DAYS FROM LAUNCH

FIGURE 4-10.--Data number as a function of days from launch, as registered on

spacecraft telemetry identification channel for science cover position.

/56

Z

54

52

_o326 366

1964

SWITCH TO TRAVELING - WAVE- TUBE .

AMPLIFIER ON DECEMBER 13

JV

40 80 120 I(SO 200

1965

DAY OF YEAR

52.59 8.40

m

40.Zl ID ?.70

_ az

43.82 7.00240

FIGURE 4-11.--Data number as a function of day of year, as registered on spacecraft telemetry

identification channel for traveling-wave-tube helix current.

169



J

MARINER-MARS 1964e

2. Interferometer effect. Since the "effective" isolation provided by the

circulator switches, which switched the transmitter or receiver to operate with the

high-gain or the low-gain antenna, was finite, the transmission or reception of the

same signal over both antennas was possible. The relative amplitude of the com-

posite signal varied with look angle to the spacecraft, and this variation, called the

interferometer pattern of the two antennas, could have a constructive or destruc-

tive effect. For Mariner IV, the pattern was such that the maximum occurred at

the look angles during encounter. Fortunately, the interferometer pattern pro-

vided higher gain than normal for the critical portion of the mission and yet had

low gain for command reception capability in the event the spacecraft lost attitude

control. The spacecraft-received carrier power plotted as a function of date in

figure 4-12 demonstrates the interferometer effect. The Earth-received carrier

power is shown in figure 4-13.

-I00 ,

PREDICTED

E -=_o

nn

hi

0

n.-u.I -120

cl[

aILl

tLI

_ -130

U_

0UJ

(/) -140

0 ACTUAL

• TOLERANCE

e_e e ' sHIGH GAIN

"

b * " " i

i-1501 I • •

28 NOV 31 DEC 9 FEB 21 MAR 30 APR 9 JUNE 19 ,JULY

DATE, 1964-1965

FIGURE 4-12.--Spacecraft-received carrier power as a function of date, showing

interferometer effect.

28 AU'_

170



$

4


-130

"10

a."u.I - 140

0o.

LIJ

-150r_

C_b.I

m

I.iJobJ

n,, -160I

nrI--

LIJ

I f I I I

PREDICTED0 ACTUAL

. • TOLERANCE

.. f_.. I•• - . "" "" "•

"---'-'J • ACTUAL POINTS ARE DATA FROMLOW GAIN

28 NOV

PIONEER DSS 11, CORRECTED

FOR CALIBRATION BIAS

! I1-

31 DEC 9 FEB I 21 MAR 30 APR 9 JUNE/_MT-5

DATE, 1964-1965

FICURE 4--13.--Earth-received carrier power as a function of date.

!

19 JULY 28 AUG

3. Frequency shift by CC&S cyclic command. Both receivers at the Tidbin-

billa Station dropped lock momentarily at the time of the CC&S cyclic command

on December 23. After that occurrence, transients were observed every time a

cyclic command occurred while in one-way lock, and analysis indicated that this

had been the case since launch. The problem was traced to an auxiliary oscillator

that, while operating in one-way lock, would shift frequency slightly when a cyclic

command pulse was received, because of additional loading on the power supply.

The frequency shift appeared in the spacecraft RF carrier signal as a drop in

received signal strength.

4. Best lock frequency change. The best lock frequency is that onto which the

spacecraft receiver automatic-phase-control loop will lock in minimum time. Be-

cause of aging and temperature differences, the Mariner IV best lock frequency

changed from its prelaunch value. However, by use of two different methods to

update the frequency during the flight, the ground stations repeatedly obtained

two-way lock with the spacecraft within a few seconds after the signal arrived at

the spacecraft.

171



P

MARINER-MARS 1964

Data Encoder

The data encoder subsystem operated normally throughout the mission. As

predicted before launch, several position skips (deck skips) and position resets(deck resets) in the telemetry commutation cycle occurred coincidentally with

changes in the spacecraft power profile. These skips and resets were due to the

susceptibility of the commutator to electrical transients and they had no effect on

the mission other than a change of reference time for science frame count and deck

synchronization. The only unexplainable deck skips occurred shortly after launch

while the spacecraft was passing through the Van Allen radiation belts and once

coincidentally with CC&S command MT-1 (first Canopus cone angle update).

An extra event was observed in a data encoder register in addition to the

single event normally associated with CC&S command MT-9 (turnoff of cruise

science, start of picture playback). This event was believed to be due to the non-

simultaneous closure of the two relays associated with MT-9 and a CC&S cyclic

command pulse which was coincident with it.

Video Storage

The functional specification for the video storage subsystem stipulated two

prime requirements: storing 20 television pictures and playing them back at an

error rate of less than 1 in 105 bits. The subsystem actually recorded and stored

21 full pictures and approximately one-tenth of the 22d picture. Thus, the first

prime requirement was fulfilled. By comparison of redundant information from

one picture to another and of pictures from the first playback to the corresponding

pictures from the second playback, the error rate was found to be satisfactory.

Among the anomalies which occurred were two abnormal false shutter indications

observed on July 15 on a data automation system (DAS) telemetry channel after

the start of picture taking. The origin of these indications is not known, but no data

were found that link them definitely to the video storage subsystem. No anomalies

affected the performance of the subsystem.

The second playback of the Mars planetary data was stopped before comple-

tion so that, after the television calibration sequence of August 30, valid Mars data

would remain on tape for a recorder playback test during the proposed reacquisi-

tion of Mariner IV in 1967. During the sequence on August 30, 10 _ pictures were

recorded and the first 5 were played back, thus protecting all the Mars data on the

second track. Characteristics of the subsystem during playback of the calibration

sequence were the same as those during playback of the planetary data.

172



¢,


Command

Analysis of spacecraft telemetry indicated that all 85 commands (by October1, 1965) sent by the command subsystem were received and correctly acted upon

by the other spacecraft subsystems. It can thus be concluded that all 85 commanbs

were received with no bit errors, though it is not possible to perform a statistical bit

error test on a spacecraft in flight. The few anomalies which occurred were re-

solved without difficulty.

Attitude Control

Sun acquisition was maintained by the attitude-control subsystem without

incident throughout the mission, except, as expected, during the midcourse-ma-

neuver sequence. A significant problem was caused by roll transients believed to

be the result of bright flashes external to the spacecraft and detected by the Ca-

nopus star sensor. It was theorized that a dust particle of minute size, floating close

to the Canopus sensor and shining brightly in the sunlight, might have so in-

creased the brightness registered in the attitude-control subsystem that the space-

craft released its lock on Canopus and began to track the dust particles. As the

particles drifted out of the field of view of the sensor, the spacecraft was left in a

star-search mode. After this phenomenon had repeated itself several times, it was

decided to initiate an alternate control mode by the transmission of a ground

command (DC-15) on December 17 to disable (desensitize) the brightness gates

and the automatic reacquisition logic in the Canopus star sensor. The capability

to react to excessive brightness had been included only to prevent Earth lock, and

by December 17 the Earth was moving well out of the way. From that time on,

there was never a loss of Canopus acquisition in spite of approximately 40 observed

roll transients.

There were no other anomalies observed which might have jeopardized the

success of the mission, although problems were encountered in the solar-pressure-

vane system operation. The first solar-pressure-vane measurements were coinci-

dent with the turn-on of the adaptive mode actuators at the time of the CC&S

L-3 event (initial star acquisition). It was then observed that the vanes had been

deployed beyond the nominal position angle of 35 ° below the plane of the solar

panels and were much nearer than expected to the plane of the panels. This vane

position difference was due to an unsuspected source of friction present during

ground testing, but not present in the space environment. Nevertheless, there was

a net restoring torque of about 1.1 dyne-cm/deg about the pitch axis and 1.7 dyne-

cm/deg about the yaw axis. The +X and -X vanes (yaw) worked properly in

173



b

MARINER-MARS 1964

the adaptive mode by canceling out an initial unbalanced torque of 25 dyne-cm,

probably due to the skewed angle of the high-gain antenna. The positions of the

+X and -X vanes are shown in figure 4-14. The +Y and - Y vanes are defi-nitely in a failed mode, probably because the actuators, which were designed to

rotate the vanes about their deployed positions to cancel out unbalanced solar

torques, locked up electrically.

The thermal actuators were designed to provide damping such that, after

the stepping motor actuators had canceled out the unbalanced torque, the limit

cycle would be damped out within the deadband of the gas subsystem. The +X

and -X vanes were expected to operate in the thermal mode to damp out the

limit cycle. After the unbalanced torque was reduced to about 5 dyne-cm, the

disturbance torques were quite variable. The maximum restoring torque of the

spacecraft over the -t-0.5 ° limit cycle was 0.85 dyne-cm and the vane damping

was 0.30 dyne-cm. It was not possible for the thermal actuators to perform their

function under these circumstances. These disturbance torques were probably due

to the normal leakage of the gas valves.

Eu

i

_" o

m

-20

oI,-

-5o

/ANGULAR POSITION MEASURED FROM

INITIAL POSITION OF DEPLOYMENT

SUCH THAT A POSITIVE ANGULAR

CHANGE PRODUCES A CONTINUOUSWAVE RESTORING TORQUE

.... t .....

. 8z

_oI-

0

zo<[

-4

326 346 366 20

1964

-- - X VANE

I

___. + X VANE

40 60 80 IO0 120 140 160 180 200 220 240 Z60 280

1965

DAY OF YEAR

FzOtJRE 4-14.--Solar-pressure-vane angular position and torque imbalance as functions of day

of year.

174



4,

b


Central Computer and Sequencer

Power was initially applied to the CC&S at approximately 4 _ hours before

launch, and operation continued uninterrupted after turn-on. On August 5, after

251 days of operation, the clock had lost a total of 188.5 sec, resulting in an accu-

mulated clock error (fig. 4-15) of only -0.00087 percent.

The unexplained anomalies which occurred during the flight that may have

been associated with the CC&S were the data encoder deck skip coincident with

CC&S-command MT-1 (first Canopus cone angle update) and the extra event

in the CC&S register following MT-9 (turn-on of cruise science, start of picture

playback). None of the anomalies affected spacecraft operation, however.

Power

Operation of the solar panels began when the Sun was acquired. Since the

spacecraft was in the Earth's shadow for a shorter time than expected, no zener

diode operation was required. Differences in solar panel currents immediately

m"One

n,-uJ

v

o¢.9

uJ

1

IE

OO

-120

-80

-40

• °--- SLOPE = -0.00085 N

_ _ SL?PE = -0.00_)82%

MT- 6"-'-1 _ i_--- SLOPE_-0.00078%

o I 128 NOV 31 DEC 9 FEB 21 MAR 30 APR 9 JUNE 19 JULY 28 AUG

DATE, 1964-1965

FIGURE 4-15.--CC&S accumulated clock error as a function of date.

/

TOCT

175



MARINER-MARS 1964

after Sun acquisition resulted from variations in panel temperatures; however,

after I hour the temperatures equalized and the currents merged. For the remain-

der of the mission the individual panels performed as predicted, and no abnor-

malities or degradations were encountered: Because of the relatively small pitch

turn of approximately 39 ° during the midcourse maneuver, the solar panels

supplied all the spacecraft power, and the battery was not required. The con-

sumption of solar panel power during the mission is shown in figure 4-16.

With the exception of a higher-than-expected battery voltage, operation of

the power subsystem was completely normal and predictable throughout the

flight. With the battery charger on, by December 2 the battery voltage had risen

to 34.8 volts de. The charger was turned off on February 11, and on February 22

the voltage reached a new high of 35.0 volts de. After that, it increased steadily,

except for a brief period during which the changing of spacecraft loads resulted ina lower voltage. When the antenna changeover was made on October 1, the

voltage read 37.2 volts de. A plot of battery voltage as a function of the day of the

year is given in figure 4-17. Two theories were advanced to explain the battery

voltage increase. They are as follows:

1. The increase was a normal consequence of the 0-g gravitational field, the

temperature effects, and the small (1.6-mA) charge current produced by the

battery voltage telemetry transducer.

750

600

430

150

I I

-- CAPABILITY

O POWER CONSUMED

TR_VEL; t_1_ -WN AvE-TUBE _

IF BATTERY CHARGER OFF _._I_OUNT(:_,I

o ooPLAY_BAC K ._

0 I

28 NOV 31 DEC 9 FEB 21 MAR 30 APR 9 JUNE 19 JULY 28 AUG 7 OCT

DATE, I cJ64-1965

FIGURE4-16.--Solar panel power as a function of day of year.

176




3el

>. 37

I--

.-I

o> 36

>,-

bJI--

I.-

< 35m

34326 366 40 80 120 160 ZOO 240 280 :320

1964 1965

DAY OF YEAR

FIGURE 4-17.--Battery voltage as a function of day of year.

2. The battery case had cracked, and the electrolyte was leaking out slowly

and evaporating, producing an open cell which the battery voltage telemetry

transducer interpreted as increased voltage.

Since the voltage did not exceed 38.0 volts de, a limit at which the second theory

would become the more likely one, a positive determination of which theory is

correct does not appear possible. Numerous attempts to resolve the problem

proved unsuccessful.

Pyrotechnics

All indications for pyrotechnics subsystem events were received as expected.

Pyrotechnics current indications were seen at the CC&S L-1 event (deploy solar

panels, unlatch scan platform), but these indications were not considered abnor-

mal since pinpullers have a possibility of shorting after they have been fired. A

nominal-voltage-to-solenoid event was received as part of the planetary encounter

sequence, an indication that one channel of the pyrotechnics subsystem was still

nominally charged and functioning.

Propulsion

Inflight telemetry coverage of the propulsion subsystem performance was

excellent throughout the mission. Nitrogen tank pressure remained constant up to

the time of the midcourse maneuver, an indication that the system was leaktight.

177



MARINER-MARS 1964

The propellant tank pressure and oxidizer pressure also maintained a constant

value up to the time of the midcourse maneuver. Motor ignition and thrust ter-

mination were verified, the burning time being the predicted 20.06 sec. The de-

livered velocity increment was calculated to be within 5 percent of that of the

commanded maneuver. Even this discrepancy might be attributable to computa-

tional error and thus does not necessarily indicate an error in the execution of the

maneuver. The propulsion subsystem remained leaktight throughout the mission,

presenting no anomalous torques to the spacecraft and providing a reliable second

midcourse-maneuver correction capability.

A pressure rise of approximately 27 579 N/sq m (4 lb/sq in.) noted in the

fuel tank on February 3 was not totally unexpected, since a similar problem was

noted on most previous propulsion subsystems. An incompatibility between the

fuel tank bladder material and the hydrazine fuel was the probable cause of the

increase. No detrimental effect to the propulsion subsystem or the spacecraft took

place as a result of this rise, and no further increase was noted.

Temperature Control

Monitored temperatures remained within allowable limits throughout the

flight. The flight temperatures were generally lower than prelaunch predictions,

as demonstrated by a typical plot--that for the solar panel temperatures--in

figure 4-18. These lower values resulted because space simulator tests prior to the

mission were conducted at higher solar intensities than those encountered during

\

_'NOMINAL '

0 FLIGHT DATA 60

I J

_o_

o

240 28o _o40 80 120 160 200

TIME FROM LAUNCH, doys

I I I 1 I I I I 1 I I l 1 I l I I

DATE, 1964-1965

FIOURE4-18.--Solar panel temperature as a function of time

from launch.

178




the flight. The flight results indicated that all temperature-control hardware

functioned normally and that no unexplained temperature anomalies occurred.The absorptivity standard provided valuable information regarding the degrada-

tion of surfaces in space and the problems resulting from space simulator testing.

Temperature histories for the four samples tested are given in figure 4-19. The

solar absorptance determinations were not successful, however, because of the

magnitude of the radiation and conduction losses.

Scieace

Television

The television instrument was turned on and operated twice during the

Mariner IV mission: during the science cover deployment exercise on February

18C

16C

14C

rEC

_. IOCo

a

4C

2C

2_

k_E ....

22C

aQ:

20C

I- 18C

16C

14C

tgC

I I o

F_i

PRE

. BASED ON

LABORATORY

S UR FA CE P RO PE RT Y

MEASUREMENTS; DOES "1NOT INCLUDE LOSSES

1

_THERMOMETER No.

, S

3

\

, J

TT!ER_R N_

! _ PREDICTION BASED ON

LABORA TORY SU RFA CE

;_ _ -PROPERTY MEASUREMENTS;

IO< DOES NOT INC LU OE LOSSES

CALIBRATION

SHIFT

, ".,,

I04

i I I L I i i i

(a) WHITE SAMPLE

o A CT UA L

-- PREDICTED

i

i -'-P.--

, i i

( b) B LA CK SAMPLE

o A CT UA L

PREDICTED

i i i i

I I I

_THERMOMETER No.

__ PRED,CT,ON"SEDON• PROPERTY MEASUREMENTS;

_. DOESo_ NOT IN CLLOE LOSSESIs\

4

i i I

240 _TH'ERMO_AETEI_ NO,

_/_1 _IN

160 PREDICTION -- O

I 'ASEDONr',,,,140 LABORATORY SURFACE _

r P RO PE RT Y M EA SU RE ME NT S ; _

10012080 DIOES NOT IN CLUDE LOSSES _,_ _ _'_z_

60 i i i i i J i

DATE, 1964-1965 DATE. 1964-1965

(C) ALUMINUM SILICONE SAMPLE

O ACTUAL

PREDICTED

I I i i

i , , , i

(d) POLISHED ALUMINUM SAMPLE

o A CT UA L

-- PREDICTED

i i i

61626

JULY

FIGURE 4-19.--Absorptivity-standard sample temperature histories.

179



MARINER-MARS 1964

ll and during the Mars encounter phase on July 14 and 15. The command to

deploy the science cover (DC-25) also turned on the television. When the switch

to data mode 3 (science data) was made, real-time telemetry data on the televisioninstrument were received. Reduction of the data from 60 frames provided the

following results: The data indicated that the shutter exposure time was 200 msec

and that no planet-in-view signal had been received or generated. The gain

control computer remained inhibited in the minimum-gain state, and the video

level was less than 2 volts. All received indications were normal for the science

cover deployment exercise, and no anomalies were detected. The turnoff of the

television instrument was timed to insure that the shutter was in a closed position.

The television was turned on by a DC-25 for the encounter operations. The

instrument had operated for approximately 7a/_ hours prior to the switch to data

mode 3. The shutter filter indication was proper, as were the exposure time of 200

msec, the lack of a planet-in-view signal, the video indication of less than 2 volts,

and the minimum-gain position of the computer. On frame 41 of the DAS frame

counter after wide-angle acquisition, a planet-in-view signal was indicated. All

other television data remained the same. It is felt that the television received the

planet-in-view signal from DAS with the Mars gate as the originator. On the next

frame, frame 42,.the data indicated more than 2 volts of video signal. Frame 43

indicated that the television had switched out of the photometer mode and into the

normal mode, and the inhibit was released from the gain control computer. The

computer remained in the minimum-gain state since the video signal was large

enough, but it was not sufficiently large to cause the shutter exposure time to

switch to the backup time of 80 msec. The first picture was recorded during

frame 43.

The data remained the same with the filter color alternating until frame 59,

where the video signal level was less than 2 volts. In frame 60 the level was again

over 2 volts. Frame 59 had indicated a green filter, and the subsystem was designed

so that the gain would change after a green picture with less than 2 volts of video.

The gain, however, did not change and stayed at minimum until frame 68, where

it stepped up one level. (In frame 67 the video signal level was again less than 2

volts.) The gain switched again in frame 70 and switched to maximum in frame

72. The video signal indications remained less than 2 volts until frame 72, where

the level again was over 2 volts. The spacecraft then switched to data mode 2

before data could be received on frame 73.

The television pictures received during the sequence demonstrate, along with

the telemetry data, that the television instrument operated perfectly.

180




Helium magnetometer

The performance of the magnetometer during all phases of the mission wasnominal, and no anomalies were observed. During the period before Canopus

acquisition, the magnetometer was used in the determination of the spacecraft

orientation, and during this roll period the X and Y components of the magnetict-1.field of the spacecraft were detcrmined. Whcn the gyros came on, ,._e magnetom-

eter indicated a shift of several data numbers; therefore, the magnetometer was

useful in establishing the time the gyros came on or went off. Since the perform-

ance of the instrument during and after encounter was nominal, the planetary

measurements obtained can be accepted with a high degree of confidence.

Cosmic dust detector

Bidirectional impacts noted in the cosmic dust detector data shortly after

launch continued for approximately 3 hours. Since the instrument would be de-

tecting spacecraft-generated noise at this time, and since discrimination between

real impacts and the spacecraft-generated noise would be extremely difficult, it

was concluded that spacecraft operation under these conditions was normal.

Starting December 6, these impacts were again noted in the data and con-

tinued for 26 hours. Investigation showed that, during this exact period, an inter-

mittent are had developed in the plasma probe high voltage. The cosmic dust

detector data were in agreement with those data accompanying the plasma probe

anomalies.

The thin-film capacitor sensors failed to register any dust particle impacts

during the first 146 days of flight (through April 23). It is now felt that the par-

ticles impacting the Mariner IV sensor were most probably silicates or cometary

ice fragments, and it is quite conceivable that the threshold mass sensitivity of the

film to these particles was too high for the instrument to have registered such

impacts.

On April 24, a film hit was recorded coincidentally with a microphone event,

indicating normal film operation; another occurred on April 27. Each of the two

films on the sensor (one on each side) recorded a hit: one on April 24, and the

other on April 27. These two events were considered real since they occurred

during the period of maximum flux experienced up to that time and the micro-

phone particle momentum data were saturated, indicating particles of sufficient

mass to activate the films. Approximately 235 cosmic dust impacts were recorded

during the mission.

181



J

J

MARINER-MARS 1964

Ionization chamber

The ionization chamber instrument performed nominally through the first

70 days of the mission (through February 5). A profile of radiation intensity was

faithfully recorded by the instrument as the spacecraft passed through each of the

Van Allen radiation belts, through the boundary of the magnetosphere, and into

the relatively stable interplanetary region. On February 5, a solar flare occurred,

and both the ionization chamber and the GM tube recorded peak activity of the

flare. As the flare subsided and the ionization chamber returned to the normal

background level, the GM tube indicated a higher rate than expected and on

February 20 the rate rose sharply, an indication of a spontaneous self-sustained

discharge. On March 17, the tube rate dropped to zero and the ionization Cham-

ber was not returning any data. No response from either sensor was observedafter that time. After investigation, it is felt that the power supply, which is com-

mon to both the ionization chamber and the GM tube, failed or was shorted.

Cosmicray telescope

The cosmic ray telescope performed satisfactorily throughout the entire

flight. Because of a shift in a zener diode, the pulse-height analyzer peak shifted

down three channels during the first 5 or 6 weeks of flight. After this period, no

more shifting was observed. The small shift does not, however, affect the data

obtained.

Trapped radiation detector

With the exception of the period from February 5 to March 5, the perform-

ance of the trapped radiation detector was normal. During this period there was

an increase in bad data points which was attributed to the reduction of the low-

gain antenna signal strength. When the switch to the high-gain antenna was

completed, the problem disappeared.

Solar plasma probe

The plasma probe functioned correctly until December 6, 1964, when the

load resistor of the high-voltage power supply failed. This failure eliminated the

internal calibration signal for the energy windows and changed the values of

some of the windows. A further consequence was that the energy windows after

the failure depended on the telemetry bit rate. For this reason no useful data were

obtained from the instrument during the remainder of the high-bit-rate data

transmission. With the establishment of the low transmission rate on January 3,

182




1965, the values of the energy windows were such that approximately half the

data returned from the plasma probe provided useful information. Data from theperiod January 3 to March 3, 1965, and from the Mars encounter have been

analyzed; analysis of the remainder of the interplanetary data is still in progress.

Ancillary equipment

After turn-on of science power after launch, the real-time portion of DAS

functioned continuously (with two short commanded interruptions), precisely

correct with no anomalies occurring. Every bit of data produced by every instru-

ment was sent to the data encoder at exactly the correct instant in exactly the

correct form. Approximately 200 million bits of data were processed by the real-

time portion. One flip-flop changed state well over 2 trillion (2 × 1012) times with-

out missing a beat.

The non-real-time portion was operated twice during the mission: first for

the science cover deployment exercise on February 11, and the second time at

planetary encounter. Every function was performed in exactly the correct manner.

The only events that could, even remotely, be considered nonstandard were two

extra end-of-tape events monitored through the real-time data. In all probability,

these events were noise pulses on the end-of-tape line which triggered the sensitive

monitor flip-flop. Functionally, these pulses were ignored even though monitored,

since DAS was designed not to act on an end-of-tape pulse prior to the beginningof the 19th picture.

The planetary scan system was energized twice during the flight also. On

February 11, the scan system responded properly to the DC-25 (encounter

science on) command by immediately entering the search mode. The correct

initial search direction was obtained, and the system made a total of 10 search

cycles. The decision was made to use a ground command to stop the scan motion,

a DC-24; it was based on the desired television pointing direction in terms of

spacecraft clock angle and the scan speed calculated from the data received for

the first six scan cycles. Upon receiving DC-24, the platform stopped 0.72 ° away

from the desired position.

On July 14, the scan system responded to a DC-25 (encounter science on) by

initiating a planet search operation in the correct direction. Because of the suc-

cessful use of a DC-24 on February 11, it was decided to use this same command to

preposition the platform, using the spacecraft automatic planet-tracking sequence

as backup. Upon receipt of DC-24, the scan platform motion was inhibited at a

clock angle of 178.45 °, or 0.72 ° away from the desired position. On July 14 at

183



MARINER-MARS 1964#

23:42:00 GMT at the spacecraft, the wide-angle sensor detected the edge of the

planet and the scan system initiated a planet-in-view signal. The planet came into

the television camera field of view 23 rain later. Data indicate that the planet

entered the wide-angle sensor field of view about 5 min earlier than expected.

184



CHAPTER 5

Tracking and Data Acquisition

SCHEDULED SUPPORT FOR MARINER-MARS 1964 MISSIONS

Tracking and telemetry coverage for the Mariner-Mars 1964 missions was

provided by the Air Force Eastern Test Range (AFETR) and the JPL Deep Space

Network (Deep Space Instrumentation Facility (DSIF), Ground Communication

System (GCS), and Space Flight Operations Facility (SFOF)). Backup coverage

for the Agena D tracking and telemetry requirements was provided by the

Goddard Space Flight Center.

Air Force Eastern Test Range

The AFETR provided coverage during the Atlas D and Agena D flight

portions of the missions. The five areas of support were: (1) metric data from

launch to 1500 meters (4920 ft), (2) engineering sequential data, (3) telemetry

data, (4) communications, and (5) data processing. The AFETR locations which

provided this support were the following: (1) Patrick Air Force Base, Fla.; (2)

Cape Kennedy, Fla.; (3) the Florida AFETR Annex instrumentation sites at

Cocoa Beach, Melbourne Beach, Vero Beach, and Williams Point; (4) Grand

Bahama Island, Bahamas; (5) San Salvador Island, Bahamas; (6) Grand Turk

Island, British West Indies; (7) Antigua Island, British West Indies; (8) Ascension

Island, South Atlantic Ocean; (9) Pretoria, Republic of South Africa; and (10)

the range instrumentation ships Coastal Crusader, Swordknot, and Twin Falls. The

range instrumentation ships were positioned in the Indian Ocean and the South

Atlantic Ocean to support the launches so as to optimize the coverage capability

throughout the launch period.

The three basic preinjection requirements for near-real-time data were initialacquisition prediction data for the Deep Space Instrumentation Facility (DSIF),

orbital elements of the parking orbit, and the initial estimate of spacecraft injection

conditions. To calculate DSIF look angles as acquisition aids and to satisfy the

need for raw data to contribute to the accuracy and reliability of the spacecraft

orbit-determination operations, the AFETR tracked the C-band (4000 to 8000

Me) radar beacon in the Agena D stage. (There was no requirement for AFETR

185



MARINER-MARS1964

tracking of the spacecraft.) Until separation, the orbits of the spacecraft and the

• Agena D were, of course, the same. At separation, a relative velocity of approxi-

mately 0.6 m/sec (2 ft/sec) was imparted to the spacecraft by springs in the

separation system; however, it did not significantly alter the total momentum.

The processing of AFETR raw data after spacecraft injection into the transfer

orbit was dependent upon telemetry identification of certain events. Relative

weighting of the various data types (e.g., range and angles with respect to DSIF

data) was a task requiring more information than was available to AFETR. The

AFETR was required to determine the Agena D orbit and furnish raw tracking

data to JPL during the launch phase. (Raw data are defined as azimuth, eleva-

tion, and range points not altered by smoothing or weighting. One exception

concerning altering of the data was the correcting of the raw data of the range

instrumentation ships for the motion of the ships.)An evaluation made by LeRC of the Atlas D/Agena D performance capa-

bilities resulted in the requirement that the telemetry system of the launch vehicle

be tracked during certain phases of the missions. In addition, range safety con-

siderations necessitated obtaining launch vehicle telemetry data during the boost

phase. The telemetry could be received by a station equipped for S-band and also

a station designed to receive Agena D telemetry. The spacecraft transmitter was

continuously radiating from the time of launch, and the telemetry signal modu-

lated the 98-kc subcarrier of the Agena D telemetry system. Both links were

exploited by AFETR to satisfy the telemetry coverage requirements.

Since the requirement for the return of data from downrange stations to Cape

Kennedy within 36 hours of the first launch could not be satisfied by the usual

data-return capabilities, a special plan was devised by the U.S. Air Force by which

especially dispatched aircraft were supplied downrange for the pickup and

delivery of these data. In addition, a telemetry reader station was established at

the AFETR Pretoria Station in South Africa whereby data displayed on oscillo-

graph-type recorders were analyzed by launch vehicle engineers and the results

were telephoned to AFETR.

Most of the communications support instrumentation provided was at

AFETR. All Florida mainland instrumentation sites were linked by teletype andvoice lines. A submarine cable through Antigua Island connected downrange

stations with Cape Kennedy. Existing high- and very-high-frequency radio links

tied the range instrumentation ships and the aircraft to the land stations. Con-

nections with the radar sites at Johannesburg and Ascension Island were made by

teletype and voice circuits.

186



TRACKINGAND DATAACQUISITION

Metric, engineering sequential, and telemetry data were all recorded against

coded time. All major AFETR locations used standard 17-digit modified binary

codes. Data recorded on film, strip charts, and tapes were reduced and processed

at Patrick Air Force Base. Telemetry data were duplicated and processed at Cape

Kennedy.

Goddard Space Flight Center

Backup C-band radar tracking coverage was provided by Goddard Space

Flight Center at two stations of the Manned Space Flight Network: Bermuda for

the early Agena D phase, and Carnarvon, Australia, for the post-Agena D retro-

maneuver period. The Space Tracking and Data Acquisition Network station atTananarive, Malagasy Republic, provided FM/FM telemetry coverage of the

Agena D. Bermuda recorded telemetry data on the Atlas D link (229.9 Mc) and

the Agena D link (244.3 Mc). Carnarvon recorded telemetry data on the 244.3-

Me Agena D link only.

Telemetry coverage requirements consisted of: (1) telemetry data gathered at

the participating stations for the life of the Agena D telemetry transmitter by one

telemetry link; (2) certain real-time readouts; (3) magnetic tape recordings,

direct-write recordings, and telemetry operators' logs; and (4) oscillograph record-

ings at Carnarvon. Certain mission computing and coordinating tasks were

required of Goddard Space Flight Center. Launch trajectory data were supplied

to Goddard computers at the AFETR. Approximately 1 week prior to the day of

launch, nominal pointing data were transmitted to the participating stations.

During the transmitting lifetime of the vehicle, these computers at the AFETR

updated and displayed the data as required. In addition, acquisition messages

were generated and transmitted to the participating stations. Radar data from

Bermuda and Carnarvon were reformatted into standard 38-character Gemini

format and transmitted to the AFETR in near-real time. Acquisition aids pro-

vided pointing information to the radar and telemetry RF inputs during the life of

the telemetry transmitted from the spacecraft. As the downrange radar source,Carnarvon was responsible for phasing with the Twin Falls range instrumentation

ship and with Pretoria during acquisition.

For telemetry reception, each of the three locations used two receivers.

Limited real-time readouts were provided by Carnarvon. Signals were recorded

on 2.54-cm (1-in.) 14-track magnetic tape at Bermuda and Carnarvon, and on

1.27-cm (0.5-in.) 7-track tape at Tananarive. Participating stations used existing

187



MARINER-MARS 1964e

full-time voice, teletype, and high-speed data circuits. No critical coverage was

required during the missions; however, special coverage was required on thevoice, teletype, and high-speed data circuits from Cape Kennedy, Goddard Space

Flight Center, and Bermuda during the launch phase.

Deep Space Network

The JPL Deep Space Network (DSN) consists of the Deep Space Instrumen-

tation Facility (DSIF), and space communications (deep space) stations based

around the world; the Ground Communication System (GCS), which connects all

parts of DSN by radio, telephone, and teletype; and the Space Flight Operations

Facility (SFOF), the command and control center located atJPL. The DSN is the

NASA facility for two-way communications with unmanned space vehicles travel-

ing 16 090 km (10 000 statute miles) from Earth and beyond.

Deep Space InstrumentationFacility

Deep space stations.--From the point of view of the Mariner-Mars 1964

project, the capabilities and facilities provided in DSIF were under the manage-

ment and control of the DSIF System Manager. (The DSIF management organi-

zation is illustrated in appendix A.) DSIF space communications stations (or deep

space stations, DSS's) are situated approximately 120 ° apart in longitude so that a

spacecraft is always within the field of view of at least one of the ground antennas.

For the Mariner-Mars 1964 missions, DSIF consisted of the following operational

stations (fig. 5-1): Pioneer DSS 11, Echo DSS 12, and Venus DSS 13 of the Gold-

stone Deep Space Communication Complex in California; Woomera DSS 41 and

Tidbinbilla DSS 42 in Australia; Johannesburg DSS 51 in South Africa; Robledo

DSS 61 in Spain; and Cape Kennedy DSS 71, the Spacecraft Monitoring Station

in Florida. Designated as prime stations for the missions were Pioneer DSS 11,

Woomera DSS 41, and Johannesburg DSS 51. The other deep space stations

served in a backup capacity: Echo DSS 12 backed up Pioneer DSS 11 ; Tidbinbilla

DSS 42 backed up Woomera DSS 41; and Robledo DSS 61 backed up Johannes-

burg DSS 51. Venus DSS 13, with its high-power transmitter, was made available

for backup use during the encounter phase of the mission when maximum com-

munications distances were involved.

JPL operates the U.S. stations, and the overseas stations are staffed and

operated by government agencies of the respective countries, with the assistance of

U.S. support personnel. The setup and equipment of a typical deep space station

188



.. TRACKING AND DATA ACQUlSlTlON

d I, ,

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FIGURE5-1 .-Location of deep space stations that supported Mariner-Mars 1964 missions.

for the Mariner-Mars 1964 missions are shown in figure 5-2. Capabilities and

operational characteristics of several of the stations are given in table 5-1.

DSIF functions include tracking (locating the spacecraft; measuring its

distance, velocity, and position; and following its course), data acquisition (gather-

ing information from the spacecraft), and command (sending instructions from the

ground to guide the spacecraft in its flight to the target, and telling the spacecraft

when to perform required operations and when to turn on the instruments for

performing the scientific experiments of the mission). Support for the Mariner-

Mars 1964 missions consisted of obtaining angular position, Doppler, and telem-

etry da ta during the postinjection phase and transmitting required commands to

the spacecraft. The DSIF was required to track the spacecraft, enabling it, in turn,

to supply raw tracking data for the determination of spacecraft orbits. These

determinations were necessary in generating predictions and for calculating the

required midcourse maneuver. First, an early orbit was determined to allow cal-

189



MARINER-MARS 1964

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191



MARINER-MARS1964

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culation of look angles for subsequent tracking. In general, DSIF initial acquisi-

tions were performed with the aid of preflight prediction data graphs or tabula-

tions and inflight prediction messages based on the actual orbit as determined by

the AFETR. Subsequent acquisitions were made with prediction messages based

on orbits calculated to satisfy the need for a final premidcourse-maneuver orbit.

Second, a final premidcourse-maneuver orbit had to be determined with sufficient

accuracy to permit the maneuver to be made within the accuracy requirements.

Raw tracking data in the form of two- and three-way Doppler and antenna point-

ing angles were provided for orbit determination.

Prior to launch, it was necessary to specify the amount of noise that could be

expected so that an a priori orbit-determination capability could be predicted as

the launch azimuth and launch days were varied. The estimates of the data noise

were used to establish the data weights which determined the orbit-determinationaccuracy capability.

A deep space station normally tracks a spacecraft from horizon to horizon,

while maintaining a two-way frequency lock between the ground and spacecraft

transmitters and receivers. One of the Mariner-Mars 1964 project requirements

was for dual spacecraft coverage. Thus, the deep space stations incorporated the

capability of simultaneously transmitting near-real-time data received from one

spacecraft, while recording and storing telemetry data from a second spacecraft

for subsequent transmission to SFOF.

Station reports were transmitted to the SFOF from each deep space station

every 30 min from launch to the midcourse maneuver and at 1-hour intervals

after the midcourse maneuver. Coverage was equivalent to 24 station-hours/day

for most of the mission. During the injection and the midcourse maneuver,

additional coverage was provided by the prime stations up to the full view period

of each station. The full 24-station-hour/day coverage of the Mariner-Mars 1964

missions was forfeited only during certain critical periods for other spacecraft

(namely, the Ranger VIII and Ranger IX spacecraft).

A brief description of the deep space stations which supported the Mariner-

Mars 1964 missions follows.

1. Cape Kennedy DSS 71, Spacecraft Monitoring Station. Located at Cape

Kennedy, Florida, this station is the initial spacecraft and telemetry link in DSN.

It functions to verify, by means of RF communications, spacecraft performance

and telecommunications compatibility with DSN during the prelaunch checkout,

launch countdown, and initial tracking phases of a mission. It is not, however,

used as a complete spacecraft checkout facility. The station participates in com-



1,


plete operational tests involving the spacecraft and SFOF. Spacecraft telemetered

data are sent from Cape Kennedy by teletype and high-speed data lines to SFOF,

and commands from SFOF are sent by teletype to this station. If necessary,

telemetry signals may be simulated so that SFOF-supervised test sequences, such

as the midcourse maneuver, may be exercised.

This station provides a stable source for measurement by telemetry of space-

craft-received power and a stable measurement of spacecraft-transmitted power

throughout the period on the launch pad up to the actual moment of launch.

Spacecraft frequencies at launch are measured, and the data are transmitted to the

usual first-acquisition station at Johannesburg, Republic of South Africa (Johan-

nesburg DSS 51). After launch, the Spacecraft Monitoring Station continues to

provide telecommunications with the spacecraft as far as 322 km (200 statute

miles) downrange at the AFETR. At that point, the functions of spacecraft track-

ing and communications are assumed by the other deep space stations.

Since long-range tracking is not conducted from this station and because of

high angular tracking rates during the launch phase, a standard 25.9-meter-

diameter (85-ft) DSIF antenna (described later in this section) is not used. Instead,

two manually steerable 1.2-meter-diameter (4-ft) parabolic reflector antennas are

provided. These antennas have azimuth-elevation mounts; i.e., they can be

moved up and down 90 ° in elevation between the horizon and zenith (the point

directly overhead) and pivoted around the vertical axis starting from true north.

One of the antennas is used for monitoring the spacecraft downrange in the

receive mode only. The second can be used to permit prelaunch two-way RF-link

calibration of spacecraft subsystems, to permit long-term power monitoring of the

spacecraft RF system, or to provide the station with the capability of supporting a

second spacecraft mission. The transmitter consists of the standard DSIF syn-

thesizer-exciter and the standard 5-watt amplifier driver used as the final ampli-

fier. It is used with a diplexer to allow simultaneous operation of both the trans-

mitter and the receiver on the same antenna.

Standard DSIF analog and digital instrumentation and recording systems

are provided, as is a ranging system for checkout of the capability to measure

range accurately by use of an automatic coded signal in conjunction with Dopplerinformation.

2. Goldstone Deep Space Communication Complex. The Goldstone Deep

Space Communication Complex, named for nearby Goldstone Dry Lake, is

located in the heart of the Mojave Desert in California approximately 72 km

(45 statute miles) from Barstow. The three deep space stations of the Complex

193



MARINER-MARS 1964

assigned to support the Mariner-Mars 1964 missions were Pioneer DSS 11, Echo

DSS 12, and Venus DSS 13, each named for the project in which it first partic-

ipated. Each station has its own antenna and tracking system.Pioneer DSS 11 was the first to begin operation (in 1958) and served as a

model for all subsequent deep space station designs. This was also the first deep

space station to be changed to S-band frequency operation. The 25.9-meter-

diameter (85-ft) antenna at this station is completely equipped in the S-band

configuration, and annex buildings and wing additions to the main control build-

ing accommodate the S-band operating and control equipment. Station addi-

tions also contain the ground support equipment used for primary control during

spacecraft missions.

Echo DSS 12 is the administration center and operations headquarters at

Goldstone. The communications building is the control center for Goldstone

communications.with SFOF and other deep-space stations, as well as for inter-

station communications at Goldstone. In the control building is the station control

and monitor console from which the station manager conducts station operations

during a mission. This console has special instruments to show the status of all

operating systems: antenna; receiver; transmitter; station instrumentation;

tracking data, telemetry, and command data handling; plus any special equip-

ment required for a mission. The 25.9-meter-diameter (85-ft) antenna at Echo

DSS 12 is equipped for receiving information from the spacecraft and transmitting

commands to it.Venus DSS 13 is used as a research-and-development facility to develop

very-high-power RF transmitters and new systems for the DSIF and also serves

as a backup for other DSIF operations. Advances in DSIF techniques are proved

and prototypes of all new equipment are thoroughly tested at this location before

being duplicated for installation at overseas stations. The main 25.9-meter-

diameter (85-ft) antenna differs from the standard DSIF antenna in that it has

an azimuth-elevation mount. The position angles of spacecraft location relative

to the antenna are measured in azimuth and elevation coordinates with a tracking

rate of up to 2 deg/sec. The antenna uses a Cassegrain cone feed system, a power-

ful 100-kW transmitter, and an extremely sensitive narrow-bandwidth receiver.

Venus DSS 13 also has a 9-meter-diameter (30-ft) reflector antenna, a i/7-scale

model of the 64-meter-diameter (210-ft) advanced antenna system at Mars DSS

14 (under construction at the time of the missions). The 9-meter-diameter

(30-ft) antenna is used for testing the design and operation of the feed system for

the large antenna.

194



t


3. Woomera DSS 41 and Tidbinbilla DSS 42. Operated and maintained by

the Australian Department of Supply through its Weapons Research Establish-

ment, these stations are each equipped with a standard DSIF polar-mounted

25.9-meter-diameter (85-ft) antenna. Woomera DSS 41 is located near Island

Lagoon, a normally dry salt lake about 22.5 km (14 statute miles) southeast of

the village of Woomera in the state of South Australia. Tidbinbilla DSS 42 is

located in the Tidbinbilla Valley about 40 km (25 statute miles) from Canberra,

Australia's capital city. The first assignment for Tidbinbilla DSS 42 was partic-

ipation in Mariner-Mars 1964 tracking operations.

4. Johannesburg DSS 5!. Johannesburg DSS 51 is staffed by personnel of

the National Institute for Telecommunications Research of the South African

Council of Scientific and Industrial Research. Equipped with a 25.9-meter-diameter (85-ft) polar-mounted antenna, the station is located near the Harte-

beestpoort Dam in the foothills of the Magaliesberg, about 64 km (40 statute miles)

north of Johannesburg. Dual spacecraft coverage for the Mariner-Mars 1964

missions was to have been provided by means of an L- to S-band conversion system

covering one spacecraft and either a portable telemetry package or a modified

angle channel of the S-band receiver covering the other spacecraft.

5. Robledo DSS 61. Operated and maintained by personnel of the Spanish

National Institute of Aerospace Technology, the Robledo Station is located near

Robledo de Chavela about 64 km (40 statute miles) west of Madrid, the capital

city of Spain. Equipped with a 25.9-meter-diameter (85-ft) polar-mounted

antenna, the first assignment of this station was participation in Mariner-Mars

1964 tracking operations.

Standard DSIF equipment and operations. In order to overcome the great

loss of energy by a signal from a spacecraft that occurs because of the tremendous

distances the signal must travel, DSIF uses antennas designed for high gain, or

very high concentration of received signal power, and powerful transmitters that

send out a very strong signal. Standard DSIF ground transmitters operate at

power levels of 10 kW. An advanced capability for transmission at 100 kW is

available at Venus DSS 13. A spacecraft transmitter, on the other hand, is verylimited in power because of size and weight restrictions. Continuing development

will increase transmitter outputs for contemplated probes to the edge of the

solar system.

Until recently, the deep space stations operated at L-band frequency: 890

MHz for Earth-to-spacecraft transmission and 960 MHz for spacecraft-to-Earth

transmission. A changeover was made to a higher RF range called S-band: 2110

195



MARINER-MARS1964

to 2120 MHz for Earth-to-spacecraft transmission and 2290 to 2300 MHz for

spacecraft-to-Earth transmission. Major modification of equipment and proce-

dures was required at all stations. Approximately 2 hours is required to change theoperating frequency at any station.

The Doppler principle is well known for its use in determining the relative

speed with which a celestial body or star and the Earth are approaching or

receding from each other (radial velocity). The Doppler shift is the apparent

change in frequency of a signal reflected from or emitted by a moving object as

the object moves toward or away from the observer. This effect may be compared

with that of a train whistle which appears to sound high in pitch as the train

approaches and then lower in pitch as it passes. This principle was adapted for

use in determining spacecraft velocity. Early spacecraft used one-way Doppler:

i.e., measuring the difference between the frequency of a signal transmitted from

the spacecraft and the frequency as it is received on the ground, which is propor-

tional to the radial velocity between the Earth and the spacecraft. Because of

inexact knowledge of the transmitted frequency, the accuracy of the measurement

of spacecraft velocity using one-way Doppler was limited to about 27 m/sec (90

ft/sec). Two-way Doppler developed for the DSIF has increased this accuracy

to better than 2.54 cm/sec (1 in./see). In two-way Doppler, a signal is transmitted

from the ground to a "turnaround" transponder (receiver-transmitter) on the

spacecraft, where it is converted to a new frequency in an exact ratio with the

ground frequency, and the signal is then retransmitted to the ground. Since thefrequency of the signal sent from the ground can be determined with great preci-

sion, the resulting Doppler information and velocity calculations are very accurate.

By two-way Doppler calculations alone, the position of a spacecraft at a distance

of several million statute miles can be determined within 20 to 50 statute miles.

A JPL-developed electronic ranging system uses an automatic coded signal in

conjunction with Doppler information to provide range measurements with an

accuracy better than 13.7 meters (45 ft) at lunar and planetary ranges.

Because of the Doppler shift and other effects, the frequency of the signal

received on the ground from the spacecraft varies widely, which means that the

receiver tuning must be changed continually. Both the spacecraft and DSIF

ground receivers use a phase-lock method of signal detection, which maintains

an automatic frequency control and keeps the receiver locked with the received

frequency.

The performance of a receiver is measured by its capability to pick up the

weak signal from the spacecraft transmitter and separate it from surrounding



TRACKlNG AND DATA ACQUlSlTlON

noises (static) originating not only in rhe Earth’s atmosphere, but from lunar,

solar, and gaIactic sources as well. DSIF receivers have a very low threshold (the

point at which the receiver can no longer detect the signal, just as in human

hearing the lower limit at which the ear no longer responds to a sound). And,

just as internal body sounds (such as that of blood coursing through the head)

interfere with the lowest external soiind discernible to the human ear, radio

receiver sensitivity is affected by internal electronic noise in the system itself. TO

help overcome this problem, advanced methods of ul tra-low-noise signal ampli -fication have been developed. DSIF S-band receiver s y s t e m use a traveling wave

maser amplifier. The maser is basically a synthetic ruby crystal immersed in

liquid helium to keep it at a very low temperature, and it operates with a

“pumped-in” source of microwave energy to augment the strength of the incom-ing signal without generating much internal system noise.

FIGURE-3. -Standard DSIF25.9-meter-diameter (85-ft)

antenna.

197



MARINER-MARS1964

im

B

The standard DSIF antenna (fig. 5-3) is a 25.9-meter-diameter (85-ft)

parabolic reflector; i.e., a perforated metal mirror that looks like an inverted

umbrella and is often called a "dish." The antenna and its supporting structure

stands 10 stories high and together weigh approximately 272 000 kg (300 tons).

Various parts of the antenna can be seen in figure 5-4. About 3630 kg (4 tons)

of electronic and operating equipment are an integral part of the antenna struc-

ture. The antenna is steerable; i.e., its "beam" or major radiation pattern can

REFLECTOR SURFACE

ONE

REFLECTOR BACKUP

ELECTRONICS

SUBREELECTOR

GEAR WHE£L

SKID

FIGURE5-4.--Skeletal diagram of standard DSIF 25.9-meter-diameter

(85-ft) antenna.

198




be readily shifted to follow a spacecraft. After a deep space probe "escapes"

from the Earth, it travels in an orbit or path similar to other celestial bodies and

"rises" and "sets" on the horizon like the Sun. The predicted or actual course

of the spacecraft is determined by the same methods astronomers use in locating

heavenly bodies; i.e., the angular position of the spacecraft relative to the star

background is defined by a set of imaginary circles (coordinates) corresponding

somewhat to Earth longitude and latitude. Each DSIF antenna is oriented to a

set of local coordinates that are used to measure the antenna pointing angles bywhich the spacecraft is located. As the antenna follows the spacecraft in its path

across the sky, a system of polar coordinates is used to measure the hour angle

(representing angular direction referenced to a station's local meridian circle;

i.e., the angle between the antenna and the spacecraft measured westward from

the DSIF antenna in the 24 hours of one rotation of the Earth) and the declina-

tion angle (representing angular direction referenced to the celestial equatorial

circle; i.e., the angle, in degrees north or south, between the celestial equator

and the spacecraft).

The gear system that mo.ves the antenna is polar mounted (fig. 5-5) ; i.e., the

axis of the polar or hour-angle gear wheel is parallel to the polar axis of the Earth

and points precisely at the North Star. This gear sweeps the antenna in an hour-

angle path from one horizon to the other. The declination gear wheel, the smaller

of the two gears, is mounted on an axis parallel to the Earth's equator (perpen-

dicular to the polar axis), which enables the antenna dish to pivot up and down.

These wheels can be moved either separately or simultaneously. The arrangement

of the gears allows the beam of the giant reflector to be pointed in almost any

direction in the sky. The motion of the antenna is controlled by the servo system.

Separate servo systems "drive" the polar wheel and the declination wheel. (Like

the driver of an automobile, the operators of the servo system control and operate

elements equivalent to those of the automobile and, in that sense, _drive" the

antenna.)

The antenna receives the strongest signals from a point directly in front of it.

Therefore, it is necessary to keep the antenna pointed in the direction of the space-

craft to receive its signals. To accomplish this, the servo system normally operates

in what is called a slave mode: angle information for pointing the antenna at

specific times is supplied to the station by a computer printout from SFOF, and the

computer and the antenna servo system operate together in an automatic loop to

keep the antenna trained on the spacecraft. Pointing-angle information based on

computer-calculated, predicted trajectory data may be supplied to the station in

199



MARINER-MARS1964

B

FIGURE 5-5.--Polar mounting of the DSIF antenna.

advance of the actual launch of the spacecraft, and this is then verified by actual

trajectory data from early passes of the spacecraft over the deep space stations.

Computer angle information may also be verified by nullifying error signals from

the receiver angle-tracking channels. (Error signals are voltages that indicate theangle between the spacecraft and the exact center of the beam of the antenna.)

With accurate information on the time and position at which the spacecraft will

appear in the antenna field of view, no time is lost in locating the spacecraft.

(Tidbinbilla DSS 42 is also equipped with a broad-beam acquisition antenna and

receiver which may be used in the early phases of a flight to direct the narrow-

beam antenna onto the spacecraft.)

2OO



,L TRACKINGAND DATA ACQUISITION

All parts of the antenna are so precisely balanced and alined that, in spite of

its weight, the antenna can be rotated at rates up to 1 deg/sec. A collimation

tower, on which is mounted a test antenna, a transmitter-receiver unit, and optical

targets, is located 1.6 to 3.2 km (1 to 2 statute miles) from the antenna of each deep

space station for use in testing and adjusting antenna alinement and operation. A

collimation tower simulates spacecraft signals in this usage. Visual checking of

antenna boresighting (adjusting line of sight, similar to alining gunsights) is ac-

complished by using an optical tracking package mounted on the 25.9-meter-

diameter (85-ft) antenna. This package consists of a television camera, a 35-mm

film boresight camera, and an optical telescope. Radio stars of known position may

be tracked by the antenna to verify pointing accuracy and other performance

factors.Tracking and telemetry data receipt.--The dish of the antenna collects the

RF energy fed into the sensitive DSIF receivers. The antenna, with an area of

almost 558 sq m (6000 sq ft), can detect extremely weak RF signals. (The signal

transmitted from Mariner IV as it approached Mars reached the DSIF antennas

at a calculated level of 10 -l° watts; the standard television signal reaches the

home receiver at a level of 10 -7 watts.)

In general, shorter RF connections between the antenna signal feed system

and the receiver mean greater antenna efficiency. Antennas designed for S-band

operation have a Cassegrain cone feed system (fig. 5-6) mounted at the center of

the reflector, which allows very short connections. This system is similar in design

to a Cassegrain telescope used in optical astronomy. Radio waves collected by the

main dish bounce up and hit a subreflector mounted on a truss-type support that

extends about 10.9 meters (36 ft) from the center of the dish. The subreflector

focuses the waves into a feed horn in the Cassegrain cone. The signal is then fed

directly from the feed horn to the low-noise maser amplifier, so maximum ampli-

fication of the weak signal occurs before it is contaminated by the electronic noise

of the rest of the receiver system.

The S-band phase-lock receiver has four separate receiving channels: two

reference channels (called sum channels) for Doppler information, spacecrafttelemetry, and television signals; and two channels that carry angle-tracking

signals for antenna pointing. The information in the sum channels is dispersed by

distribution amplifiers in the receiver system to proper destinations in the telem-

etry instrumentation and data handling systems in the control room.

Command transmission.--The accuracy of the trajectory of a deep space

probe is controlled by transmitting command signals that initiate roll, pitch, and

201



h

MARINER-MARS 1964

SIHONI

202



w


yaw maneuvers, as well as propulsion, ignition, and timing sequences, which are

determined by computations made from tracking data. Signals are also sent to the

spacecraft to change data rates; to change the type of telemetry information being

transmitted; to turn the transmitter on or off or change its power; to reorient the

spacecraft or its antennas; or even to switch antennas, receivers, and transmitters.

Sending a command to the spacecraft is somewhat the reverse process of

receiving a signal from it. The transmitting station is equipped with a 10-kW

transmitter. The exciter and controls of the transmitter are in the control room;

the RF power amplifier and associated equipment are mounted on the antenna.

The power level of the signal put out by the exciter is very low (on the order of a

few watts). This is amplified in the power amplifier so that the signal radiated

from the antenna is very strong (at least 10 000 W). The transmitter is normallyused with a diplexer, a device which allows simultaneous operation of both a

transmitter and a receiver at different frequencies on a single antenna and feed

system.

The commands to be sent to the spacecraft originate in the SFOF control

center at JPL. The necessary information is sent over teletype link from SFOF to

the participating deep space station. Since an incorrect command could result in

possible damage to the spacecraft, extreme precautions are taken to insure ac-

curacy. Command information from the SFOF is usually sent three separate times

over the teletype link to the station and is also verified by telephone. Ground

command and control equipment at the station includes read, write, verify equip-

ment that carefully checks a command before it is sent and as it is being sent to the

spacecraft. This equipment reads and verifies the teletype message, transforms the

command into a signal for radio transmission, and monitors the transmitted RF

signal bit by bit. If any bit proves incorrect, transmission is automatically stopped

to make the necessary correction. Very often, especially if the command is to be

stored in the spacecraft memory equipment for later execution, the command as

received by the spacecraft is telemetered back to the ground and checked again

with the transmitted command. A special-purpose computer is used to execute

these check routines.Data handling.--Signals processed by the receiver at a deep space station are

sent to ground instrumentation and data handling equipment in the control room.

This equipment includes paper-tape and magnetic-tape recorders and ultraviolet

oscillographs. The tracking-data handling equipment records angle measurements

of antenna position, Doppler frequency measurements, range measurements, and

time. These data are recorded on paper tape for immediate teletype transmission

203



im

MARINER-MARS1964

to the SFOF for use in determination of spacecraft orbit, calculation of maneuver

parameters, command decisions, and prediction of arrival time at the target.

Telemetry signals from the Mariner IV spacecraft came in on the receiversum channel time-multiplexed; i.e., the signals from the various measuring instru-

ments on the spacecraft were carried on one composite RF signal sequentially

(time multiplexed). This composite signal was "unscrambled" by decommutators

in the ground telemetry subsystem so that each signal was identified by a channel

number. A digital method of signal coding was used for transmission of data from

the spacecraft to the Earth. With the digital coding and phase-shift modulation

techniques employed, it was possible to increase the efficiency of data trans-

mission from the spacecraft and to simplify data handling at the ground station,

because digital signals can be formatted for direct inputs to computers and for

teletype transmission.

Detected unscrambled signals are recorded on magnetic tape so that complete

permanent records of all telemetry data from a spacecraft are available for later

data processing at SFOF. Certain selected spacecraft telemetry signals are dis-

played at the station as they are received for the use of operating personnel in

maintaining contact with the spacecraft. The Mariner IV digital data were

exhibited on special displays.

Because the quantities of data produced during a mission are enormous and

are constantly growing as space projects become more sophisticated, increasing

use is being made of onsite data processing in the DSIF to relieve the burden bothon the communications lines to the SFOF and on the SFOF data-processing sub-

system. Systems at the Goldstone Deep Space Communication Complex and

Tidbinbilla DSS 42, controlled by general-purpose digital computers, convert and

reduce some of the unscrambled spacecraft data to digital format for transmission

by high-speed data lines directly to the SFOF computers. Video data are split off

from the sum channel by the receiver and are sent to special video equipment in

the control room at the station for processing.

In addition to processing and recording spacecraft-telemetered data, each

station also processes and records data generated by ground equipment, on such

parameters as received signal strength, transmitted power, station-equipment

condition, and calibration voltages. This information is processed by the digital

instrumentation system, which uses general-purpose digital computers that accept

and process both analog and digital signals. All ground data are recorded on

digital magnetic-tape recorders, and certain selected data are recorded on

punched paper tape for transmission over teletype circuits to the SFOF.

2O4



TRACKINGAND DATA ACQUISITION

All taped information sent to the SFOF is labeled and identified by data type,

time received, station, and spacecraft number. Because time reference is a critical

factor in tracking determinations and in other DSIF functions that depend upon

the timing of electronic phenomena, the time of receipt of telemetry data is

recorded to an accuracy of at least 0.01 see. All data received during a mission are

recorded on magnetic tape for a permanent record and for the use of scientists and

engineers in evaluating the results of a mission.

DSIF acquisition procedures, which include antenna pointing, receiver

tuning, transmitter tuning, ranging lock, and telemetry decommutation, are so

precisely timed and coordinated that it is possible to start recording data from I to

10 rain after radio contact with the spacecraft is established and to start trans-

mitting data to SFOF within 4 to 16 min.

Ground communicationsystem

Interstation communications and those with the SFOF are by telephone and

teletype through the DSN Ground Communication System. For the Goldstone

Deep Space Communication Complex, a multiplex microwave link is used to

facilitate the handling of the vast amount of data that must be transferred from

these stations to the SFOF. This link carries multiplex channels for voice and

teletype transmission, two circuits for high-speed digital data transmission, and

one video channel. Communications to and from Woomera DSS 41, Tidbinbilla

DSS 42, Johannesburg DSS 51, and Robledo DSS 61 include full duplex teletype

circuits, high-speed data circuits, and voice circuits. These stations are linked

directly to the SFOF by high-speed teletype for digital-data transmission by

means of the Australian COMPAC cable (both Woomera and Tidbinbilla), the

NASA Worldwide Communications Network (NASCOM, managed and operated

by the Goddard Space Flight Center), and the transatlantic telephone cable,

respectively. Communications to and from Woomera DSS 41 and Tidbinbilla

DSS 42 pass through the NASA switching center in Adelaide, South Australia,

and those to and from Robledo DSS 61 normally pass through London, England.

Data obtained by the deep space stations are transmitted to the SFOF in realor near-real time by teletype and high-speed data circuits and are also recorded on

magnetic tape at each deep space station to be mailed later to the SFOF. Teletype

is the primary means of transmitting tracking and telemetry data to the SFOF and

of sending predictions and other data to the stations. Teletype transmission is at

the rate of 60 words/min. Analog and video circuits were also made available

during the encounter phase of the Mariner IV mission. Voice circuits are used for

205



#

MARINER-MARS 1964ii

transmission of high-priority communications other than data. Technical control

of all communications facilities throughout the DSN is exercised by the SFOF.

Onsite communications are handled by telephone, local paging system, and

closed-circuit television. In addition, high-speed data "hot" lines for fast, real-time

transmission of tracking information during critical periods of a mission and tele-

type transmission are also provided between stations of the Goldstone Deep Space

Communication Complex.

Communications capabilities between the various deep space stations and

SFOF are shown in figure 5-7 along with AFETR communications capabilities.

Although high-speed data lines were available for the Mariner-Mars 1964 mis-

sions, it should be pointed out that they were never considered fully operational

because of inherent anomalies in the system.

TIDBINBiLLA

DSS 42

WOOMERA

DSS 4i

ROBLEDO

DSS 6t

JOHANNESBURG

OSS .51

GOLDSTONE

DEEP SPACE

COMMUNICATIONCOMPLEX

m _ -- m

SWITCHING

CENTER,

ADELAIDE,

AUSTRALIA

LONDON.

ENGLAND

O

SPACE

FLIGHT

OPERATIONS

FACILITY,

PASADENA,

CALIFORNIA

AFETR

NOTES:

AFETR COMMUNICATIONS CIRCUITS

WERE USED ONLY PRIOR TO THE

MID COU RS E MAN EU VE R

COMMUNICATIONS CAPABILITIES

WERE VARIED IN ACCORDANCE

WITH THE HIGH- AND LOW-ACTIVITY

PHASES OF THE MISSION

_ FULL DUPLEX

TELETYPE

(SIMULTANEOUS

TRANSMISSION

IN BOTH

DIRECTIONS)

HAL F D UP LE X

TELETYPE

(TRANSMISSION IN

ONE DIRECTION

ONLY ON ANY

ONE CIRCUIT)

VIDEO

VOICE

-----I_ HIGH-SPEED DATA

-_oP,,ll_.----- FACSIMILE, FULL

DUPLEX TELETYPE

FIGURE 5-7.--GCS communications lines for DSN.

206




Space Flight Operations Facility

The DSN Space Flight Operations Facility (SFOF) at JPL served as the

command and control center for the Mariner-Mars 1964 space flight operations.

Various equipment and personnel were housed within this building for perform-

ance of the following operations:

1. DSN control of the deep space stations.

2. DSN ground communications operation control of the Ground Communi-

cation System and of the internal voice, teletype, and high-speed data circuits andequipment.

3. Data processing, monitoring, display, distribution, and storage (library).

4. Space flight operations direction.

5. Space flight operations evaluation and spacecraft command analyses in

real and near-real time.

The equipment and personnel 1 within SFOF which were part of the Space

Flight Operations System for the Mariner-Mars 1964 project are discussed in

chapter 6.

MARINER IV TRACKING AND DATA ACQUISITION SUMMARY

At liftoff of the Mariner IV spacecraft, 15 metric, 32 engineering sequential,

and 8 documentary cameras were operating at Cape Kennedy. Nine land-based

radar stations and three range instrumentation ships provided tracking coverage

during the mission. Continuous coverage was obtained from liftoff until 11 rain 44

sec after liftoff. Continuous telemetry coverage was provided from 7 rain before

liftoff until 12 rain 8 sec after liftoff.

Cape Kennedy sites and DSS 71, the Spacecraft Monitoring Station, locked

onto the spacecraft at liftoff. Approximately 14 sec later, Patrick Air Force Base

acquired and began sending trajectory data to Cape Kennedy. Bermuda began

tracking approximately 3 rain after liftoff. Approximately 10 min later, Antigua

acquired Mariner IV and sent data for 1 rain. Bermuda data were then back on

the line. The flight trajectory was reported as nominal when Grand Turk began

transmitting data 21 rain after liftoff. The "suitcase" telemetry deep space station

at Fort Dauphine, Malagasy Republic, was locked onto the spacecraft for 4 rain,

1This includes both mlssion-dependent equipment and personnel (specifically for the Mariner-Mars 1964

project) and mission-independent equipment and personnel (for use on other projects as well as the Mariner-Mars

1964 project).

207



,m

MARINER-MARS1964 o

and Pretoria maintained lock and sent tracking data for 4 min also. The range

instrumentation ship Twin Falls then acquired the spacecraft.

Woomera DSS 41 was the first deep space station to acquire the Mariner IVspacecraft. This was accomplished approximately 49 min after launch at 15:10:48

GMT on November 28. At that time, a signal level of -120 dBm was reported,

and at 15:14:00 the level was -87 dBm. Spacecraft RF power was reported to be

up at 15:12:42. Sun acquisition began at 15:23:00, and the spacecraft was on solar

power 1 min later. Sun acquisition and two-way lock were confirmed at 15:30:57.

At 15:35:00, the signal level at Woomera DSS 41 was -90 dBm, and the near-

Earth trajectory phase of the mission was considered successfully covered.

Woomera DSS 41 continued tracking with normal received signal level and with

good two-way tracking data being obtained through the end of its view period at

00:46:56 on November 29.

Johannesburg DSS 51 acquired the spacecraft approximately 7 hr 10 rain

after launch at 21:31:40 GMT on November 28. A normal signal level and good

tracking data were received. Pioneer DSS 11 acquired Mariner IV at approxi-

mately 16 hr 20 min after launch at 06:41:41 on November 29. Passes over

Woomera DSS 41, Johannesburg DSS 51, and Pioneer DSS 11 during this period

lasted approximately 9.5 to 10.5, 10 to 10.5, and 11.5 to 12 hours, respectively.

The Earth track of the spacecraft is shown in figure 5-8.

On November 30, Pioneer DSS 11 transmitted three DC-21's to Mariner IV

during the second pass of the spacecraft over that station. These commands posi-tioned the spacecraft in roll attitude to allow the first acquisition of the star

Canopus by the Canopus sensor. (The times that these and all other commands

were transmitted and the deep space stations which transmitted them were given

previously in table 4-I.)

On December 5 (after an abortive attempt at a midcourse maneuver the

previous day, when the spacecraft lost its lock on Canopus), the Pioneer DSS 11

receiver, demodulator, decommutator, and teletype encoder were in good lock

with a signal level of -143 dBm at 06:44:00 GMT. Two-way lock was reported at

07:00:30. With all stations and systems reported ready at 13:03:00, the initial

command of the midcourse-maneuver sequence was transmitted 2 rain later. The

midcourse maneuver was successfully completed at 16:54:57 on December 5.

Woomera DSS 41 acted as backup and tracked the spacecraft while Pioneer DSS

11 was sending the commands for the 20-sec motor burn.

On December 13, during the initial period of the cruise phase, Pioneer DSS

11 transmitted a DC-7 to Mariner IV to transfer the radio from the cavity

208



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MARINER-MARS1964m

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amplifier to the traveling-wave-tube amplifier in order to increase the signal level

received by the ground receiver. A DC-21 (roll override) and a DC-15 (Canopus

sensor gate override) were successfully transmitted by Woomera DSS 41 on

December 17, and the DC-15 condition remained in effect for the remainder of

the mission.

During the period beginning December 17 and ending at Mars encounter on

July 15, the DSIF engaged in extensive tests and exercises designed to optimize

operational procedures during the encounter phase of the mission. Some of these

tests and exercises were as follows:

1. Nominal encounter test (April 20). Using the proof test model (PTM)

spacecraft as a data source, this test Served to exercise the Track Chief and SFOF

personnel and procedures.

2. Command loop lockup exercises (between April 28 and July 15). This

series was performed to exercise command procedures and operations personnel

and to reaffirm continually the condition of the command subsystem on the space-

craft.

3. Data mode 4 data transmission test from Johannesburg DSS 51. This test

was conducted to ascertain whether a reasonable picture could be constructed

from real-time mode 4 data transmitted from Johannesburg DSS 51 to SFOF on

the partial RF communications link under specific conditions of radio propaga-

tion. Magnetic tapes containing recorded mode 4 data were played back at

Johannesburg DSS 51 as the data source.

4. Roll position simulation (May 5). A test was conducted to verify whether

or not spacecraft roll orientation data could be properly displayed and interpreted

at SFOF. Spacecraft movement about the roll axis was simulated by varying the

received signal level at the deep space station and SFOF. During the actual en-

counter, the primary source of the roll orientation data would be the spacecraft

signal strength received at the deep space station.

5. Alternate-plan encounter test. The alternate plans in the event of an

anomaly at encounter were exercised to evaluate response time with respect to

data analysis and command recommendations.

6. Nominal encounter test (May 18). Again using the PTM spacecraft, this

test was conducted to familiarize personnel with the proposed encounter sequence

and to determine the feasibility of the procedure with regard to command

response, timing, and data evaluation.

7. Two-way lockup tests. In preparation for the occultation experiment,

several two-way lockup tests were conducted to determine the time required to

210



t


achieve two-way lock with the spacecraft and to determine the advantages and

disadvantages of using the 10-kW (versus the 100-kW) transmitter during the

period when the spacecraft left the occultation region at Mars.

8. Preliminary encounter readiness test (June 1). This test was conducted to

verify the readiness of the technical analysis teams in the SFOF.

9. Encounter backup-mode tests (June 15, 24, and 25). These three tests

served to verify the use and handling of the commands for the backup mode at

encounter.

10. Occultation experiment operational tests (June 16, 23, and 24). Three

tests were conducted to exercise: (a) generation and transmission of real-time,

exit-occultation predictions for Pioneer DSS 11 and Echo DSS 12; (b) real-time

computation and display in the SFOF of Doppler residuals; (c) use of open-loopreceivers and recorders at Pioneer DSS 11 and Echo DSS 12; (d) loading of track-

ing-data punched paper tapes at critical occultation periods; (e) tracking-data

sample rate changes during critical periods; and (f) transfer of spacecraft signal

lock between Pioneer DSS 11 and Venus DSS 13.

11. Operational readiness tests. Two such tests were conducted prior to en-

counter to verify readiness of all personnel, procedures, and equipment.

Tracking by the three prime stations was continuous during the early cruise

phase, except for 4 days in January and 16 days in February when Johannesburg

DSS 51 was released for Ranger VIII spacecraft tracking operations. Mariner IV

was unmonitored for about 7 hours during each of these 20 days. When, on

January 31, Woomera DSS 41 was also released for Ranger VIII operations,

Tidbinbilla DSS 42 assumed prime status for Mariner IV.

During the cruise phase in February, Pioneer DSS 11, Tidbinbilla DSS 42,

and Johannesburg DSS 51 were tracking the Mariner IV spacecraft. During the

pass of the spacecraft over the Goldstone Deep Space Communication Complex

on February 20, Venus DSS 13 tracked instead of Pioneer DSS 11, since the latter

was assigned to track the Ranger VIII spacecraft that day. Johannesburg DSS 51

was assigned to track Ranger IX during 16 days in March.

At the beginning of April, the same stations were tracking Mariner IV as weretracking at the beginning of February. On April 27, Woomera DSS 41 began

sharing tracking operations with Tidbinbilla DSS 42, alternating with three passes

for Woomera and then four passes for Tidbinbilla. Tracking continued in this

pattern until, on July 29, Woomera and Tidbinbilla both tracked one pass.

Woomera DSS 41 then assumed tracking operations until the pass over Tidbin-

billa DSS 42 on July 6, when both stations tracked the spacecraft. On July 5, Echo

211



MARINER-MARS1964U

DSS 12 also began tracking Mariner IV; on July 6, Robledo DSS 61 joined the

operations. Thus, during the Mars encounter phase, Pioneer DSS 11, Echo DSS

12, Woomera DSS 41, Tidbinbilla DSS 42, Johannesburg DSS 51, and RobledoDSS 61 were all tracking the spacecraft. The actual encounter with the planet

Mars on July 15 was recorded by Pioneer DSS 11 and Echo DSS 12, with Venus

DSS 13 in a standby status to provide command transmission if needed. During

the encounter, the 100-kW transmitter at Venus DSS 13 provided the uplink

signal which enabled the spacecraft to establish successfully the uplink lock when

it left the occultation region.

Prior to the encounter sequence, 11 commands were transmitted by DSIF: 2

by Johannesburg DSS 51 and 9 by Pioneer DSS 11. All commands were executed

on time and functioned normally. The RF signal from Mariner IV was lost by the

DSIF at 02:31:12 GMT on July 15 because of the entrance of the spacecraft into

the occultation region at Mars. At 03:25:06 on the same day, the RF signal was

reacquired when the spacecraft left the occultation region.

The first picture data were received by the DSIF at 13:01:58 GMT on July

15. All picture data were recovered during the tirst picture playback, which was

completed at 19:26:33 on July 24. The second picture playback began at 21:21:53

on July 24 and ended at 03:36:02 on August 3.

Late on July 24, Echo DSS 12 ceased tracking and Pioneer DSS 11 continued.

Then, on July 29, Echo DSS 12 resumed tracking and Pioneer DSS 11 (:eased.

After the spacecraft pass on July 24, Woomera DSS 41 discontinued tracking for 4days. Tidbinbilla DSS 42 and Johannesburg DSS 51 both ceased tracking on July

31. Although Robledo DSS 61 had discontinued tracking operations on July 24, it

resumed these operations on July 30. On August 3, Venus DSS 13 was responsible

for transmitting the commands which effectively turned off the transmission of the

second-playback television data from the spacecraft and resumed the transmission

of engineering and science data. Tracking operations during August and Septem-

ber were conducted by Echo DSS 12, Woomera DSS 41, and Robledo DSS 61.

On September 29, Johannesburg DSS 51 tracked for one pass because of

maser problems at Robledo DSS 61, which caused that station to cease tracking to

resolve the problems. Thus, on October 1, Echo DSS 12, Woomera DSS 41, and

Robledo DSS 61 were tracking the Mariner IV spacecraft. Echo DSS 12 recorded

the end of the mission at 22:05:07 GMT on October 1, when the spacecraft signal

was lost because of a DC-12 transmitted by Venus DSS 13 to switch the space-

craft transmitter from the high-gain to the low-gain omnidirectional antenna.

212



t

L

CHAPTER 6

Space Flight Operations

The Mariner-Mars 1964 Space Flight Operations (SFO) system consisted of

the Earth-based mission-independent facilities and personnel of the DSN and the

AFETR specifically assigned to support the Mariner-Mars 1964 missions, and the

mission-dependent equipment, personnel, and procedures required to direct and

evaluate the space flight operations for the missions. The overall operational

organization for space flight operations is given in figure 6-1. The SFO system

was, in many instances, distinct from the mission-independent organization ofSFOF. The space flight operations organization for the Mariner-Mars 1964 mis-

sions is illustrated in figure 6-2. From the point of view of the Mariner-Mars 1964

project, the capabilities and facilities provided by the SFO system were under the

management and control of the SFO system manager. (The SFO management

organization is described in app. A.) The Mariner-Mars 1964 project manager

was in full charge of all activities connected with space flight operations, but the

SFO director had immediate primary control.

SYSTEM FUNCTIONS AND RESPONSIBILITIES

The primary objectives of the SFO system were to track the spacecraft,

evaluate the tracking (Doppler and antenna pointing angles) and telemetry

(engineering and scientific measurements from the spacecraft) data, process and

disseminate the data, determine appropriate commands to be sent to the space-

craft, and provide proper command transmission. Secondary objectives were to

evaluate and compare tracking and telemetry information with that obtained from

ground observations during the flights and to provide information and experience

on space flight operations and mission performance for use in future mission

planning. The following capabilities were required of the SFO system:

1. To determine the trajectory and perform any necessary corrections to it.2. To receive, record, and interpret telemetry data continuously for the entire

flight so as to monitor spacecraft performance, detect any "alarm" conditions, and

perform any necessary command functions.

3. To describe spacecraft position and science sensor orientation in a variety

of coordinate systems; in particular, to generate a variety of geometrical relation-

ships between the spacecraft instrument sensors and the planet Mars.

213



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SPACE FLIGHT OPERATIONS

SFOF I

OPERATIONS

I MANAGER

• DATA I

PROCESSINGPROJECT

ENGINEER

[

._ COMPUTERPERATORS

ANALYSIS _.

AREAPROGRAMER

T.s ___PERATORS

COMPUTINGSUBSYSTEMOPERATORS

--7

I COMMUNICATIONS [

COORD NATOR I

I I I I

I TECHN,CALDATA1OORDINATOR DISTRIBUTOR

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L_I_ Il TECHNICIANS OPERATORS

TPS TELEMETRY PROCESSING SUBSYSTEM

II S°PP°RTYSTEM

I LEADERSONNEL I

FICURE 6-2.--Space flight operations organization for Mariner-Mars 1964 project.

4. To maintain sufficient communications between the SFOF and the DSIF

in order to obtain simultaneously telemetry data from two spacecraft and tracking

data from one spacecraft from each of the prime deep space stations.

These capabilities were incorporated in the SFO system design formulated by

the DSIF operations manager, the SFOF operations manager, the data processing

project engineer, the communications coordinator, and the directors Of each of

three technical analysis teams. The technical analysis teams assisted in the defini-

tion of the standard mission, recommended courses of action during nonstandard

events to optimize the value of the mission, and performed the intragroup andintergroup technical liaison required to achieve these objectives. These teams were

as follows:

1. Flight Path Analysis and Command (FPAC) Team, whose responsibilities

were: (a) use of tracking and pertinent telemetry data to obtain the best estimate of

the actual spacecraft trajectory and, supported by DSIF, interpretation of the data

supplied by the deep space stations; and (b) generation of spacecraft commands

215



P

MARINER-MARS1964,I

t

affecting the flight path, using the support of other technical analysis teams to the

degree required. Functions included computation, maneuver analysis, orbit deter-

mination, tracking data analysis, and trajectory calculation.

2. Spacecraft Performance Analysis and Command (SPAC) Team, whose

responsibilities were: (a) determination of spacecraft performance, and (b) sub-

sequent recommendation of the required commands to be sent to the spacecraft to

achieve mission success. Functions included those related to engineering me-

chanics (such as monitoring of spacecraft temperature, pyrotechnics, scan actuator,

and science cover) and propulsion analysis.

3. Space Science Analysis and Command (SSAC) Team, whose responsi-

bilities were: (a) representation of the experimenters in the mission operations; (b)

examination and analysis, to the extent necessary, of the data from the spacecraft

science instruments and any other sources to keep appropriate personnel informedas to the status of the science instruments and to recommend appropriate action to

improve the scientific worth of the mission; and (c) control of the flow of, and the

mathematical operations performed on, the data pertinent to the experiments

during the period between data acquisition and data transmission to the appro-

priate scientists.

A full complement of personnel was continuously maintained in the SFOF during

the high-activity phases of the missions, such as operational tests, launch, mid-

course maneuver, and encounter; during normal cruise-phase operations, moni-

toring was still continuous, but a minimum staff was maintained.The central point of the SFO system during the missions was the mission

control room in the SFOF. Here all elements of the SFO system were brought to a

single focus. Two communications and control consoles provided voice contact

with any of the deep space stations and with any of the operational areas in the

SFOF. Two closed-circuit television monitors in each console displayed pictures

from any of approximately 60 closed-circuit television cameras located in the

SFOF. The mission control room is contained within the SFOF operations area,

as is a mission status board displaying current significant facts about the progress

of a mission.

DATA PROCESSING SUBSYSTEM

After elaborate routing through the various communications lines, the

Mariner IV telemetry and tracking data were sent to the SFOF for final process-

ing and display to the science and spacecraft performance analysts. Tracking

216



SPACEFLIGHTOPERATIONS

data from the AFETR were processed, and spacecraft look angles and prediction

data were transmitted to the appropriate deep space stations in accordance with

established procedures. The tracking data from the deep space stations were

processed to establish the orbit of the spacecraft, and additional prediction data

were then transmitted to the appropriate deep space stations. These data estab-

lished the orbit of the spacecraft and determined the requirements for the mid-

course maneuver. Midcourse-maneuver commands were prepared and trans-

mitted to the appropriate deep space station (Pioneer DSS 11) for transmission

to the spacecraft.

The mathematical processing of incoming data constituted the major effort in

data handling in SFOF. The type of incoming data (whether telemetry or track-

ing) and the ultimate users determined the type of computation required. Theprincipal groups which used spacecraft or spacecraft-related data and the type of

data each used were as follows:

SPAC Team: engineering telemetry

SSAC Team: science telemetry

FPAC Team: tracking

DSIF net control group: DSIF status

Mission and operations control

group: summary of all data and

their status

Two parallel multicomputer systems (strings) were incorporated to provide

on-line data processing, display, and command/control functions for the Mariner-

Mars 1964 missions. Each string contained an IBM 7040 input/output processor,

an IBM 7288 multiplexer, an IBM 7094 main processor, and two IBM 1301 disk

files. Teletype data received at SFOF were entered directly into the IBM 7288

multiplexer. Raw telemetry data received over microwave link or by telephone

(high-speed data) underwent preliminary processing in the telemetry processing

subsystem and were then entered into the IBM 7288.

The telemetry processing subsystem (fig. 6-3) converted the telemetry data to

a format compatible with the IBM 7288 high-speed subchannel and with the IBM

7094-compatible magnetic tape. The conversion process was accomplished either

in real time using the high-speed-data communications system or in nonreal time

using data recorded on magnetic tape. Data from all sources were then converted

by the IBM 7288 into the proper format for the actual data processing, which

included data identification, conversion to engineering units, data alarm monitor-

ing, and conversion to the proper format for printing and plotting.

The following were fed into the IBM 7040 input/output processor: all data

from teletype, telephone lines, microwave channels, and the telemetry processing

217



MARINER-MARS 1964,e

HIGH-SPEED I

DATA I

MAGNETIC

TAPE I

I

III j.A.,...__ MARS 1964

I FORMATTER

,]i[

TELEMETRY

DEMODULATOR

POP PROGRAMED DATA PROCESSOR

I

, i']

FIouR_. 6-3.--Telemetry processing subsystem.

subsystem, as well as all requests, parameters, and data from the user areas by way

of an inquiry station or card readers. The inputs were identified and separated by

mission number and data type,, including tracking, telemetry, and administration

data. All incoming data were written on raw data tape in a sequential mode, with

proper identification to allow separation and processing in the IBM 7094 in non-

real time. If overlaps occurred in DSIF coverage and two stations sent identical

telemetry data to SFOF, the choice of transmission to be inserted in the IBM 7040

was determined by use of the data-processing control console. The transmission

from the rejected station was recorded, but was not available for further real-time

processing.

The IBM 7094 then performed functions in both on-line and off-line modes.

The raw data table was sorted into a master data table for analysis routines. Final

reduced data prints and plots of all data were stored in disk files for immediate

processing during the operational phase and for later retrieval for subsequent

analysis and processing. Forms and tabulations for user groups and selected

parameters for status display were distributed. Thus, through the use of the IBM

7094, telemetry data from the telemetry processing subsystem and raw data tapes

recorded by the IBM 7040 were processed, and the midcourse-maneuver com-

mands and DSIF predictions were computed and generated for transmission to thedeep space stations.

System tests of the data processing subsystem were conducted by using a

Mariner-Mars 1964 spacecraft model. In one such test, the spacecraft, monitored

by the system test complex, transmitted telemetry data in real time for the DSIF

ground telemetry subsystem (demodulator, decommutator, and teletype encoder)

218



QSPACEFLIGHTOPERATIONS

in the system test complex. The ground telemetry subsystem processed the data

and produced a teletype paper tape record of the spacecraft telemetry according

to a defined format. On a non-real-time basis, the teletype paper tape record was

hand carried from the system test complex (at the JPL Spacecraft Assembly

Facility) to the SFOF for processing by the data processing subsystem. This sub-

system, using SPAC and SSAC basic IBM 7040 telemetry computer programs,

processed the spacecraft telemetry data by using the teletype paper tape record

inputs. Video pictures were processed by the IBM 7094, a film recorder, and the

JPL photography laboratory.

DATA FLOW

Real-Time Data

The nature of the Mariner IV space-flight operations was such that real-time

data flow was of primary concern. Control of this flow and of data processing was

necessary so that the proper data were received and processed at the proper inter-

vals. These data were received in real or near-real time and were automatically

entered in the data processing subsystem. These data were then operated upon and

displayed online in the user areas as rapidly as operational priorities and user pro-

grams permitted. Data were classified as real time if they were transmitted by

microwave, telephone line, or teletype within 5 rain (from the Goldstone DeepSpace Communication Complex) or 10 min (from the other deep space stations)

from the time of receipt at the deep space station. If buffered in a link (including

the data processing subsystem) for more than 5 min but less than 30 min, data

were classified as near-real time.

Non-Real-Time Data

Data received by the data processing subsystem in the form of either magnetic

tape recordings or delayed transmission from a communications link (more than

30 min after receipt of the data at the deep space station) were classified as non-

real-time data. The main characteristics of these data were that their processing

was delayed and the results were prepared off-line from the data processing sub-

system. There was no necessity for a feedback path from the analysis area or for

very rapid throughput and display. Data from all of the sources were entered

directly or by magnetic tape in either of the two available IBM 7040 input/output

processors, which performed the same input functions on non-real-time data as

performed on real-time data, but which recorded the collected and formatted

219



MARINER-MARS1964o

e

input data on magnetic tape only. These tapes were then processed collectively on

the IBM 7094 main processor at prescheduled intervals, and magnetic tapes were

generated to drive the off-line display devices.

General Paths

The typical data flow to, within, and from the SFOF is shown in figures 6-4

to 6-6. The flow from the SFOF was comprised of acquisition and tracking infor-

mation as well as commands for the DSIF, general status information, and space-

craft performance data. The incoming data circuits were routed through the

SFOF communications terminal to the IBM 7288 for processing by the IBM 7040.

These data were also made available on teletype machines and closed-circuit

television in the user areas.

OPERATIONAL MODES AND COMPUTER PROGRAMS

Flight status and data type determined the data-processing-subsystem mode

of operation and the control programs. The six operational modes available (table

6-I) provided different data throughput and failure recovery times as required for

various mission conditions. The flow of data was controlled from the data process-

ing control console. All switching of computer subsystem and input/output

equipment and control of computer program priorities were initiated at this con-

sole. Control functions were based on equipment performance and on operationalrequirements as specified by the SFO director. The IBM 7094 computer programs

were controlled by a percentage time-sharing scheme. The percentages were fixed

by the SFO director and were based on user preflight requests and the standard

sequence of events. The seven user areas in SFOF contained computer input/out-

put equipment for performing data analysis and/or command/control functions

in the data processing subsystem. Typical user area equipment is shown in figure

6-7.

The computer programs described in table 6-II provided the required data

in the proper format in the areas of flight-path analysis, telemetry monitoring,

spacecraft performance analysis, science analysis, command generation, and con-

trol and operational information. The data processing programs included those

necessary to accept and log incoming data, to monitor spacecraft telemetry and

tracking data, and to perform alarm functions in real time. Real-time analysis

programs were necessary in the determination and analysis of spacecraft perform-

ance, the evaluation of science experiments in nonreal time, and the analysis of

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FROM

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FmURE 6-7.--Typical user area equipment.

the flight path in near-real time. Non-real-time analysis programs were not re-

quired in the actual space flight operations, but complemented the detailed

analyses. The latter category included those programs which determined physical

constants and locations from tracking data and performed detailed analyses and

reconstructions of video data. Inputs from non-real-time analysis programs into

the real-time analysis programs may be used during a mission, provided suchinput or substitution of parameters is deemed necessary and appropriate to the

fulfillment of mission objectives.

The Mariner-Mars 1964 mission-dependent data-processing programs were

divided into the following categories:

1. Real-time operational monitoring and processing programs including all

IBM 7040 computer programs.

228



W

t


2. Near-real-time operational space-flight analysis programs processed in the

IBM 7094 computer for operational flight-path analysis and spacecraft and science

instrument performance analysis.

3. Non-real-time space-flight analysis and research programs having mul-

tiple options and functions.

TELEMETRY DATA FORMS AFTER PROCESSING

Mariner-Mars 1964 telemetry data appeared in the following forms after

processing:

1. Raw data. Although a somewhat inconvenient format for data analysis,

raw data were thought to be a readily available, reliable source of real-time data.

2. IBM 7040 formatted data. Use of the IBM 7040 computers permitted the

telecommunications system telemetry channels to be displayed on a separate

format and to be converted to data number, by far the most convenient form. It

:*.,as decided to display the IBM 7040 data as data numbers instead of engineering

units to allow easier real-time evaluation during the flight.

3. JPEDIT. JPEDIT was designed to be a non-real-time output of the total

data stream in a convenient format for performance analysis, data records, etc.

Its information content was supposed to be the same or better than that of the IBM

7040 formatted data.

4. EDPLOTM. This program was designed to take the data used for

JPEDIT and plot them in the form of data numbers as a function of time for per-formance analysis, data records, etc.

5. SSDM. The output of this program, a condensed version of JPEDIT, was

designed to give an output only when a data number changed in a specificchannel.

6. MDL. The master data library (MDL) was implemented so that a list of

the best data available from all sources, including selected tracking station func-

tions as well as spacecraft data, would be available at one location. The MDL is

described in detail later in this section.

Raw data displayed on 60-word/min printers proved to be the most reliable

source of real-time data during the mission. The raw Teletype data could be con-verted to a usable form in a matter of seconds. These data, usually available

several seconds ahead of the IBM 7040 formatted data output, were sometimes

used even when the formatted data (converted to data number, the most con-

venient form) were available. It was found that the JPEDIT data contained

consistently more erroneous indications than the IBM 7040 formatted data, and

thus this type did not play a significant role in mission operations. Since their data

229



0

MARINER-MARS 1964

source JPEDIT did not meet expectations, EDPLOTM and SSDM data were not

effectively used either.

MARINER-MARS 1964 MASTER DATA LIBRARY

In August 1964, a program was initiated to design, develop, and implement a

system to produce a master data library (MDL) of all data received and recorded

during the Mariner-Mars 1964 missions. The MDL would provide the best source

of tracking and telemetry data received and recorded by DSIF, as well as discrete

DSIF instrumentation performance parameters, thus producing a history of the

missions from which postflight analysis could be accomplished on the spacecraft

subsystems, the scientific instruments, the spacecraft trajectory, and the perform-

ance of DSIF.As implemented, MDL was comprised of three types of data "tables": telem-

etry, tracking, and comment. The telemetry data table was a series of two types o

digital recorded magnetic tapes generated by the IBM 7094 computer programs:

station master merge tapes and composite master merge tapes. The station master

merge tapes, generated by the "merge" program for each station, contained the

best telemetry data and ground instrumentation performance parameters re-

corded during each station's tracking. These telemetry data were derived from the

three data sources processed: the demodulator input, the demodulator output, and

the Teletype output. The selection of the data source was made by the merge pro-

gram, based upon the data quality and continuity of each data stream within a

recorded source.

The composite master merge tapes, also generated by the merge program,

contained a continuous and sequential stream of the best telemetry data derived

from the composite of deep space stations receiving and recording telemetry data

(station master merge tapes) during each day of the mission. The overlap of telem-

etry data between deep space stations was eliminated by data source selection,

based upon data quality and continuity of each recorded source. Ground instru-

mentation performance parameters were not contained on these tapes.

The tracking data table was developed from both real-time and non-real-time

data. Real-time data (sampled only at low rates) were received from the deep

space stations by Teletype communications into the SFOF and processed through

the data-processing subsystem in real time. Non-real-time data (sampled at both

low and high rates) were received from the deep space stations and were also proc-

essed through the data-processing subsystem. Both types of data were maintained

230




on digital recorded magnetic tapes in formats compatible with the tracking-data-

processor/orbit-data-generator programs, which were in themselves user com-

puter programs. The raw tracking data as recorded on Teletype paper tapes were

maintained as part of the MDL.

To explain peculiarities and/or anomalies which occurred in both the telem-

etry and tracking data tables, the supporting space-flight operations logs, DSIF

operations logs, and MDL data-processing logs were microfilmed and maintained

in the comment data table. Actual tracking and telemetry data were printed in

tabular outputs and, together with the MDL data processing logs, provided a

history of the MDL processing.

MARINER iV DATA PROCESSING AND RECOVERY SUMMARY

Data Recovery Summary

Basically, two types of data were recovered during the Mariner IV mission:

tracking data and telemetry data. The spacecraft transmitted approximately

3×108 bits of data to the Earth: 107 bits of television data, 97×106 bits of

engineering data, and 193×106 bits of science data. The number of telemetry

measurements as a function of time is presented in figure 6-8.

During the 307 days which elapsed from the first day after launch (November

29, 1964) to the antenna switchover (October 1, 1965), approximately 9500 hours

of tracking were performed at about 10 hr/day/station for three stations on the

average. Horizon-to-horizon two-way data for two passes were taken each week by

each station. Approximately 550 000 points of 60-sec data, 320 000 points of 10-see

data, and 3 200 000 points of 1-sec data were taken. The 60- and 10-sec points

were sent in real time to the data processing subsystem at SFOF. The 1-sec points

were sent in non real time for Mariner IV MDL postprocessing. Of the estimated

total of 106 points transmitted in real time, about 300 000 were precision two-way

points.

Approximately 98 percent of the real-time data was processed and logged by

IBM 7040 computers, and approximately 90 percent of the non-real-time data wasrecovered by IBM 7094 postprocessing. The 2 percent loss in real-time processing

was due to communications outages and computer outages, and the 10 percent

loss in non-real-time postprocessing was due primarily to the assignment of time

with data and the inability of the JPEDIT program to identify preambles. Of the

data received at the deep space stations, from 99.5 to 100 percent has been

recovered for the Mariner IV MDL.

231



MARINER-MARS 1964

60.0

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TOTAL MEASUREMENTS --_ /

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//_y _-- SYNC IDENTIFICATION

AND STATUS

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PICTURE AND

_b_ STORED DATA

I

1964 1965

P.SNOV 240EC 2FEB 14MAR 23APR 2JUNE 12JULY 21AUG 30SEIDT 9NOV

OATE

FIGURE 6-8.--Number of telemetry measurements as a

function of date.

Data Processing Summary

During the high-activity phases of the mission (launch, midcourse maneuver,

and encounter), redundant systems were used wherever possible. At SFOF, all

operational computers were put in use in two computer strings. One IBM 7094

computer was used primarily for flight-path programs, and the other was used for

engineering and science programs. All programs were loaded into both strings so

that an immediate backup was available for any given program.

Because of certain anomalies in handling tracking data during the early hoursof the mission, it was decided to back up the data processing subsystem by punch-

ing cards with tracking data information to be read directly into the IBM 7094

tracking data processor and merged with data coming into the IBM 7094 by way

of the IBM 7040. During the first 30 days of the mission, only about 10 to 15 per-

cent of the data was processed through normal channels (IBM 7288-IBM 7040-

232




IBM 7094). When the punched card information was added, this figure increased

to about 85 to 90 percent. Modifications made to the mission-independent editor

in the IBM 7094 increased data recovery to about 85 to 90 percent through the

data processing subsystem, and it remained at that level or better throughout the

rest of the mission.

During the cruise phase, only the IBM 7040 computer was used in real time.

The output consisted of "quick-look" formats for science and engineering data,

high-speed teleprinter formats of selected engineering measurements, and 76.2- by

76.2-cm (30- by 30-in.) plots of selected science and engineering measurements.

During that time, the IBM 7040 also generated a log tape consisting of all incom-

ing raw data. No real-time processing of tracking data was performed. One of the

IBM 7040 computers was replaced with an IBM 7044 during the cruise phase.

The IBM 7094 user programs were run on a production-type schedule. The

IBM 7040 log tape was removed each morning and edited by the IBM 7094.

_e { ..... nrt onerln rlnff nro_rnrn_ were then rim on the edited data. The com-.............. o--ee .... _ v-_ ....................

puter printout was duplicated by document control and then delivered to the user

areas. User program output for a 1-day period was delivered to the user areas by

noon of the following day.

During the encounter and the beginning of picture playback, full data process-

ing capabilities were used. Both computer strings were exercised: one for flight-

path programs and the other for engineering and science programs. After playbackof the first picture, the computer coverage was reduced to two IBM 7040 com-

puters making redundant recording tapes. IBM 7094 video processing, performed

each morning, produced two outputs: a computer printout and a magnetic tape.

The computer printout consisted of a number of formats which displayed picture

data as well as summary data concerning the picture. The picture data were dis-

played in a 200- by 200-decimal element matrix and a printout of the binary

serial bit stream. The magnetic tape produced by IBM 7094 video processing was

made to be compatible with the film recorder.

Continuous orbit runs were made on one computer string, and periodic orbit

runs were made on the other computer string from 12 hours before closest ap-

proach to the exit of the spacecraft from the occultation region. From 30 rain

before the spacecraft entered the occultation region until it left the region, residual

data were generated by determining the difference between the incoming data

and the predicted data. These were then transmitted to the SFOF from Echo DSS

12 and plotted in near-real time on the FPAC team's 76.2- by 76.2-cm (30- by 30-

in.) plotter. Unexpectedly large residuals resulting from difficulties with orbital

233



#

MARINER-MARS1964

calculations at encounter caused part of the plotting exercise to be of limited value;

however, it was possible to verify the existence of the planet's atmosphere and

make some preliminary estimates of its character in near-real time from observa-tion of the plots. A delay in locking up the closed loop at Echo DSS 12 when the

spacecraft left the occultation region made it impossible to plot residuals at that

time.



6

CHAPTER 7

Scientific Results and Conclusions

FIELDS AND PARTICLES EXPERIMENTS

Introduction

The Earth and other inner planets are immersed within clouds of plasma

(ionized gas), magnetic fields, and a flux of high-energy particles and electro-

magnetic radiation that come primarily from the Sun. Even though not all the

accelerated particles that fill the space between the planets are derived from the

Sun, most are accelerated to high energy by solar processes. The magnetohydro-

dynamic and high-energy particle flux in interplanetary space accounts for a host

of terrestrial phenomena, including magnetic storms; auroral displays; u,_rup_:.... t,on:

of radio, telephone, and telegraph communications; and the Van Allen belts.

The Sun itself generates the full spectrum of electromagnetic radiation. Of

the solar particles, the most energetic are protons (hydrogen nuclei) and electrons

in the million-electron-volt (MeV) energy range. These are born in great ex-

plosive outbursts of energy from the Sun known as solar flares. Even an ordinary

flare expends energy equivalent to millions of hydrogen bombs exploding all atonce. A major flare can release a flood of energy equal to more than a billion

hydrogen bombs exploding over a 30-min period: the equivalent of about 1019

kW-hr of energy. Exposed to such energy, particles reach relativistic velocities

that exceed one-tenth the speed of light, and masses of ionized gas or plasma are

ejected to travel across interplanetary space with velocities reaching 1126 km/sec

(700 statute miles/see).

The Sun's corona (which extends for unknown millions of miles into space) is

made up chiefly of hydrogen gas, which, at a temperature of 1 000 000 ° K (about

1 800 000 ° F) near the Sun, is mostly ionized into protons and electrons. A few

alpha particles (helium nuclei) may also be present in the corona. The density of

the Sun's corona starts at about 3× 107 protons and electrons/cm a (4.9 × 108/in. _)

near the Sun and decreases to a value of 5 to 20 protons and electrons/cm _ (82 to

328/in. 3) in the vicinity of the Earth. There is an almost continuous, high-velocity

outflow of extremely hot ionized gas caused by solar activity which results in

magnetohydrodynamic expansion of the corona itself. This outflow, the so-called

solar "wind" or solar plasma, has continuously varying energy, and the changes in

235



*

MARINER-MARS 1964

FIGURE-1 .-Configuration of magnetic field lines of force in interplanetary space.

its density, velocity, and direction account for variations in the “weather” in inter-

planetary space.

When the solar wind or the plasma ejected as a result of solar flares travels

across space, it “drags” the Sun’s magnetic-field lines of force along with it. These

lines of force are endless, each drawn through interplanetary space in a roughly

spiral configuration radiating from the Sun for as far as the solar wind blows. The

more-energetic particles from solar flares move along these magnetic-field linesof force. Therefore, as may be seen in figure 7-1, the Mariner IV spacecraft

detected these particles only when crossing the magnetic-field lines of force that

connect to the position of the flare on the Sun. Even the energetic particles called

cosmic rays, which come to Earth from outside the solar system, move along paths

which are determined by the configuration of interplanetary, magnetic-field lines

of force of solar origin. The strength and direction of the solar magnetic-field lines

236



SCIENTIFICRESULTSAND CONCLUSIONS

of force across space are also important factors affecting the "weather" in inter-

planetary space.

When the solar wind arrives in the vicinity of tile Earth, it encounters the

Earth's magnetic field. The airfoil-shaped region created by the solar wind's inter-

action with the Earth's magnetic field, compressing it and confining it, is called the

Earth's magnetosphere. Its boundary is called the magnetopause. Since the

Earth's magnetic field deflects all particles, a disturbance is created in the smooth

flow of the solar wind, just as a rock protruding out of a stream will disturb the

smooth flow of the water. The deflected particles, beginning at a distance of about

64 400 km (40 000 statute miles) on the sunward side of the Earth, form a kind of

half-sphere around the Earth on the sunward side. This is termed the bow shock or

shock front. It is thought that the Earth actually leaves a "wake" in the solar wind,streaming out into space away from the Sun. This, the so-called magnetic tail, is

believed by many to be quite long, possibly 1 astronomical unit (i.e., the distance

lrom the _,tc_ of the Sun to the oo,,ter of tho Enrth) The theoretical configllr_-

tion of the interaction of the solar wind with the Earth's magnetic field is illus-

trated in figure 7-2.

D IRE CT ION OF

EARTH'S TRAVEL

SHOCK FRONT MAGNETOPAUSE

FIGURE 7-2.--Effects of interaction of solar wind with Earth's magnetic field.

237



4_

,iP

MARINER-MARS 1964

Actually there is not much material in interplanetary space; densities of as

few as 20 protons and electrons/cm 3 (328/in?) seem to be the rule, and at this

value it would take a volume the size of the Earth to make up 3.6 kg (8 lb) ofmatter. But, despite the fact that this density of material is almost 1 million

times less than that in the best vacuums available in the laboratory, this exceed-

ingly tenuous matter and the electromagnetic fields and radiations associated with

it could have very serious consequences in future planetary and interplanetary

space travel. Therefore, instruments which could help in determining these possi-

ble consequences were considered essential for the Mariner-Mars 1964 payload.

The particles and fields in space and in the vicinity of Mars were thought to

be particularly significant for study, as was the cosmic dust distribution on the way

to Mars. The possibility of radiation belts near Mars had been hypothesized, and

the proximity of the planet to the asteroid belt made it seem possible that particu-

late matter might exist in greater quantity near the orbit of Mars than near the

orbit of Earth. Since the Sun is the dominating factor in the planetary and inter-

planetary environments, measurements of its effects throughout the mission were

considered especially important.

All six instruments chosen for the Mariner-Mars 1964 fields and particles

experiments--helium magnetometer, cosmic dust detector, ionization chamber,

cosmic ray telescope, trapped radiation detector, and solar plasma probe--were

capable of providing data on interplanetary space and on the near-Earth and

near-Mars environments as well. The instruments were carefully positioned on the

spacecraft to survey space and obtain the maximum amount of useful information

during the mission. Their locations and "look" angles are shown in figure 7-3. The

scientific investigators for the experiments are listed in table 7-I, and the measure-

ments made by the instruments during the Mariner IV mission are discussed in the

following paragraphs.

Helium Magnetometer

The helium magnetometer was included in the Mariner-Mars 1964 science

payload in order to measure the variations in magnitude and direction of the

planetary and interplanetary magnetic fields. It operated on the principle that the

presence of magnetic fields would modify the transparency of a plasma of meta-

stable helium ions to infrared radiation of 1.083-micron wavelength. The instru-

ment, weighing 3.07 kg (6.77 lb) and requiring 7.3 watts of power, basically con-

sisted of an infrared source, a helium cell, an infrared detector, and a set of coils

around the helium cell. In order to minimize the effects of spacecraft fields, it was

238



SCIENTIFIC RESULTSAND CONCLUSIONS

mounted in a remote position high on the low-gain antenna waveguide of the

spacecraft.

The coils produced a rotating magnetic field in the helium, resulting in a

synchronous modulation of the infrared intensity sensed by the detector. An

SOLAR PLASMA PROBE +x

C = DIRECTION TO CANOPUS

Z=DIRECTION TO SUN

M=DIRECTION OF MARINER 2_" TRAVEL

X,Y=AXES OF MARINER 2_ THROUGH SOLAR PANELS IONIZATION CHAMBER

FIGURE 7-3.--Locations and "look" angles of the Mariner IV interplanetary instruments.

239



MARINER-MARS 1964

Table 7-1.--Scientific investigators for Mariner-Mars 1964 science experiments

Experiment Investigators Affiliation

Television ......................

,p

Helium magnetometer . .. .. .. .. ..

Cosmic dust detector ............

R. B. Leighton a

B. C. Murray

R. P. Sharp

R. K. Sloan

J. D. Allen

E.J. Smith a

P. J. Coleman, Jr.

L. Davis, Jr.

D. E. Jones

W. M. Alexandera

O. E. Berg

C. W. McCracken



California Inst itute of Technology




University of California, Los Angeles


Brigham Young University and Jet Propul-

sion Laboratory

NASA Goddard Space Flight Center



Ionization chamber ..............

Cosmic ray telescope .............

Trapped radiation detector .......

Solar plasma probe ..............

Occultation ....................

L. Secretan

J. L. Bohn

O. P. Fuchs

H. V. Neher a

H. R. Anderson

J. A. Simpson a

J. O'Gallagber

J. A. Van Allena

L. A. Frank

S. M. Krimijis

H. L. Bridge a

A. Lazarus

C. W. Snyder

A.J. KlioreaD. L. Cain

G. S. Levy

V. R. Eshleman

G. Fjeldbo

F. Drake

NASA Goddard Space Flight CenterTemple University

Temple University



University of Chicago

University of Chicago

State University of Iowa



Massachusetts Institute of Technology

Massachusetts Institute of Technology


Jet Propulsion LaboratoryJet Propulsion Laboratory


Stanford University

Stanford University

Cornell University

aPrincipal investigator.

external field disturbed the normal detector output signal, producing an error

signal that was fed back to the coils as a current of sufficient magnitude to nullify

the external field. The measurement of this current provided the output signal of

the helium magnetometer. The dynamic range of the instrument was 4-360 gam-mas along each of the three spacecraft axes, with a 0.35-gamma resolution per

axis? The magnetometer made four vector measurements during every 12.6 see at

the high (331_-bits/sec) data transmission rate and every 50.4 see at the low (81/_ -

bits/see) data transmission rate.

t 1 gamma = 10 -5 gauss; the magnetic field of the Earth at the equator is about 0.5 gauss or 50 000 gammas.

240



t


For the first 162_ hours of flight, the Mariner IV spacecraft was programed to

roll at a rate of about 1 revolution every 30 rain. Since the magnetic field of the

Earth was known, it was possible to correct the magnetometer data for any bias

caused by the small remnant field of the spacecraft itself. The roll calibration

insured that later data taken from the instrument in the very small field of inter-

planetary space could be assigned absolute values.

As Mariner IV began its journey into space, it first passed through the mag-

netosphere of the Earth. The power spectra of the magnetic field variation in the

region of turbulence behind the shock front were determined for variations with

25-sec to 30-min periods and an average total energy density of about 5 gammas_/

cm _. Pronounced oscillations with periods near 3 rain were detected; they

appeared to have been caused by hydromagnetic waves associated with fluctua-

tions in the position of the magnetopause.

Through a fortunate circumstance, Mariner IV passed through the shock

front several times. After the spacecraft had passed through the front once, con-

ditions apparently changed, causing the front to move outward and overtake the

spacecraft from the rear. Later, Mariner IV caught up with the front at its new

location and passed through it again. There were approximately seven transversals

between 36.6 and 38.6 Earth radii (i.e., between approximately 233 300 and

246 200 km (145 000 and 153 000 statute miles), after which no further evidence of

Earth-related phenomena was obtained. A portion of the helium magnetometer

data taken during the passages of the spacecraft through the shock front is shown

in figure 7-4.

The average magnitude and direction of the interplanetary field were con-

sistent with values taken during the previous missions of Mariner II in 1962 and

the first interplanetary monitoring platform IMP I (Explorer XVIII) in 1964.

Throughout the mission in interplanetary space, the magnetometer data showed a

pattern of alternating disturbed and quiet intervals related to daily changes in

solar activity. Apparently because the Sun in late 1964 and in the first half of 1965

was slightly "quieter" (i.e., less active) than in 1962, the fluctuations in the inter-planetary magnetic field during the Mariner IV mission were not quite as large as

the fluctuations observed during the Mariner II mission. Several times, however,

Mariner IV recorded abrupt changes in the magnetic field coincident with the

arrival of plasma thrown out by solar flares. Large irregular fields were observed

2 days after the solar flare of February 5. On one occasion, the magnetic-field

value reached 35 gammas, or more than 7 times the typical average daily value.

241



MARINER-MARS 1964

SHOCK

FRONT

SHOCK SHOCK SHOCK SHOCK

FRONT FRONT FRONT FRONT

SHOCK

FRONT

w

p-

REGION OF

TURBULENCE

X-AXIS

INTERPLANETARY

SPACE

REGION OF

TURBULENCE

INTER -

PLANE-

TARY

SPACE

REGION OF

TURBU-

LENCE

INTERPLANETARY

SPACE

REGION

OF

TURBU-

LENCE

,% •.

• . :_::.::.::.•._... ":" " _ ... --7-._-'_.: .: .....:= %.:..• .w,-_, -.._..-.._._.. .: .... .: 4. _- a.." :-'-: • ' ":'..J"_.'..

-_"F --"_•:'"_ " _ " --

, I l I

Y-AXIS

I .-:---.-- •

'"..-.'" . . ." - .. :.. '.:|.." ..... ,.-". "'-_" ..._.

• "....'....:,:':." 7..:... :.'.,." ." ..:

.... ". "':: ".T.: ":-- •', " •,. ",.

i

i I __A._

Z-AXIS

°

°+ °°

I I I I I I

;

"" i...". , :. _ -.. _

I " .'. ::'4: "•" .i "-'-"'-.

: = ..-.-; '..:!-..... : "-':" • i i :"

. J_ L---- -.L-- ............. -.--J

1! _ --.:i "

• - i". ::".: i "t -:

L• . ": . .: . i ...... "'_

•. • _ •. _._L ," ." ._'.. .- -L-- .: .".. '. .: '"'; -- '_" "._'

-...-._....:.:.. L+..:.. !-:.,....:. :':':'. I:'" " '

: "_:" : " i

• I i

J_.----__l l i, I : L l J--;.

06:30 06:40 06:50 07:0_

GMT, NOVEMBER 29, 1964 •

i I

37 _'_

GEOCENTRIC DISTANCE, Eorth rodli

: .-....._.:-- .:_; ..::..-...

•:,.... -.-.:.,_...- --'-

i ............... J-

06:10 06:2006:00

I36

FIGURE 7-4.--Helium magnetometer data taken during Mariner IV's passages through shock

front of Earth•

242




The Mariner IV helium magnetometer data corroborate the previous obser-

vations that the interplanetary magnetic field lies near the plane of the Sun's

equator but is strongly disordered. The average magnetic vector pointed away

from the Sun at an angle of about 45 ° to the direction of the Earth's orbital motion

and nearly parallel to the plane of the ecliptic. There was a tendency for the inter-

planetary field to exhibit a polarity pattern that rotated with the Sun. The specific

patterns during the Mariner IV observations were generally different from those

observed by Mariner II and IMP I. The Mariner IV data also revealed an evolu-

tion in the polarity pattern from one solar rotation to another; in fact, there were

long intervals where the pattern became difficult to recognize.

Because Mariner IV was moving outward from the Sun far beyond the orbit

of the Earth, it provided the first opportunity to investigate whether or not the

Earth had a long magnetic tail streaming out into space away from the Sun.

During late January and early February, Mariner IV passed through the pre-

sumed region of the ..... ,;o ,a;1 _f ,h o 12_rth (hehlnd the l¢.arth a_ viewed from

the Sun). At this time, the geocentric distance of the spacecraft was approximately

3300 Earth radii (approximately 21 million km (13 million statute miles)). During

7 days of flight within a geocentric angle between 1° and 5° from the presumed

centerline of the magnetic tail, the helium magnetometer data revealed no change

in the interplanetary magnetic field that would be associated with the Earth's

magnetic tail, if it existed. Although a negative result, it is considered particularly

significant as evidence against conjectures that the Earth's magnetic tail might bevery long. There is a possibility, of course, that such a tail exists and that it was

bent considerably above or below the route traversed by Mariner IV.

Theorists had predicted that the strength of the magnetic field of Mars was

probably quite small: no stronger than one-tenth that of the Earth. This prediction

was based on Mars' mass and its rate of rotation (virtually the same as the

Earth's). Predictions had been generated as to what distance from Mars the

helium magnetometer would detect or show the effect of having encountered a

magnetic field at Mars, assuming various values for the strength of the field. If the

field were strong, the instrument would have shown an effect several hours before

closest approach to Mars. The most sensitive point for detection was shortly after

closest approach and just before the spacecraft entered the Earth occultation

region at Mars (i.e., before the spacecraft passed behind Mars as viewed from the

Earth). No conclusive evidence for a magnetic field was noted on the output of the

instrument.

Figure 7-5 shows the calculated shock-front locations for Mars, assuming

243



MARINER-MARS 1964

magnetic moments of 10 -a (0.1 percent) and 10 -4 (0.01 percent) of the value for

the Earth. The trajectory of Mariner IV was such that it should have encountered

the shock front at these locations, if such a front existed. The data from the helium

magnetometer which were obtained as Mariner IV flew past Mars are given in

figure 7-6. The only effect which might be interpreted as representing a shock

front occurred at a distance of about 15 000 to 20 000 km (9300 to 12 400 statute

miles) from the center of Mars slightly after the closest approach of the spacecraft

to the planet. On the basis of these measurements, the upper limit of the magnetic

moment of Mars may be set at 0.03 percent of the value for Earth.

Cosmic Dust Detector

The cosmic dust detector was designed to measure the mass and flux distri-

bution of interplanetary dust particles in the vicinity of the Earth, between theorbits of Earth and Mars, and in the vicinity of Mars. The detector consisted of a

square (approximately 22.1 cm (8.7 in.) on each side) 0.08-cm-thick (0.03-in.)

MARS' MAGNETIC MOMENT

EARTH'S MAGNETIC MOMENT =

MARS

120° MARINER

TRAJECTORY

20 000

FIGURE 7-5.--Theoretical shock-front locations at Mars, assuming magnetic moments of 0.1

percent and 0.01 percent of the value for Earth.

244



SCIENTIFIC RESULTS AND CONCLUSIONS

15 v v i ! t I V 1 u v 1 _ v v v T V !

I0 FIELD INR DIRECTION (NORMAL TO ECLIPTIC PLANE) CLOSEST APPROACH:

OI:O0:SBGMT, 13 196 kmh..

05 .,,,.+.','" ' .................. , .................... "',.,.,,. ...... .,, ..................... , ............. ,.+.., . ....... ,,I ,'r},,.,, ,,. OCCULTATION IIW, t

i*t'11 jalOt ,l l I '

-5

-I0

-15 i _, £ : L _ , , , l _ A t l l i l I A J

15 l I 1 l ; I r i v v i r r y i i ! ! i

I0 FIELD IN T DIRECTION (PARALLEL TO ECLIPTIC PLANE)

5

EE _''h_''_''''1_'h''i_I_'_h_'_"'r''_m_'om''_I+''"'i'''_'¢''_'_'"i'_'P''_'h"h'_'*_ h̀_"i ......"'""h',Hl+l '",i°'"'OCCULTATION ,'0,.[h,+'"",',,,',h,o 0

o" -5.J

_uJ -tO

b._

_L_ -15 * _ * _ ' I i _ I I I I I I i I I I I

_5_ , t , , , t , t , , , , y , , , , ; ,

LO I0 t FIELD tN MARS' NORTH DIRECTION

o t5 ............... ' ......... ' ............ ....... " '..' "' " "' ................ ,............... It ,,,,l'" t,l,,, ',l ,','

-u0 ul,rht',".'"'" OCCULTATION

!-15 i , I I i I I I i A I i I i i i I i I J

15 l t l 1 t l r ! , _ v t 1 ! , t I ! r

IO ABSOLUTE FIELD "I' .,, 1

"' ,OCCULTAT ON l

5 ,",'I",'""..., III ........ ,'+'Flu '"..' ....... +, ............. ""'"'"""'""d, ....... It_, +++_IIO'""",,'I+,HI .,,,,I,....'H ' ql l l''+,.'l *++,''II''.'*'''_

0 _. l i I I I l I i I I I i I I i A l I ]

I00 021 B4 007 6B OIB 5`2 I01 36 424 '21 745 13 205 `21 314 35 926 51 590 67 `203

DISTANCE FROM CENTER OF MARS, km

I I

,9:00 20100 ' `2,1oo' `2,1oo' _,+':oo' 00100 ' o/:oo ' 02':00 ' 03':00' o,,Ioo ' '05:00

JULY 14 _,,JULY 15 _-',

GMT AT SPACECRAFT

FIGURE 7-6.--Helium magnetometer data taken during encounter of Mariner IV with Mars.

aluminum sensor plate mounted perpendicular to the velocity vector of the space-

craft. To this plate was attached a sensitive microphone (threshold sensitivity= 6.04-0.7 X 10-5 dyne-sec).

The sensor plate of the cosmic dust detector was coated on both sides with a

thin film of insulating material, over which was deposited a thin film of metal. A

voltage was applied between the outer conducting film and the aluminum plate,

thus making it a capacitor. When a particle of solid matter from space penetrated

the metallic outer film and the insulator, it discharged the capacitor and gave rise

245



b

t

MARINER-MARS1964

to a pulse. The microphone produced a signal related to the mechanical acoustic

pulse of the sensor plate. The electronic instrumentation contained a storage sys-

tem which recorded all dust-particle impacts and provided eight levels of pulse-

height analysis to each signal from the microphone. The dynamic range of the

pulse-height analyzer extended to 1.96×10 -_ dyne-sec. The storage system also

provided information concerning the occurrence of coincidences between signals

from the microphone and the capacitor sensor. A calibration device contained in

the instrumentation performed a calibration approximately three times a day

when the telemetry data transmission rate was high (331_-bits/sec) and approxi-

mately once each day when the data transmission rate was low (8a/_ bits/see). The

cosmic dust detector weighed 0.95 kg (2.1 lb) and required 0.23 watt of power.

On its trip to Venus in 1962, the Mariner II spacecraft reported that the

number of dust particles in free space was very small: only about 1 for every

10 000 in the vicinity of the Earth. However, Mariner IV was expected to en-

counter larger numbers of particles for two reasons:

1. It would pass close to several known meteoroid streams, two in December

and' one in March; these streams are presumably composed of particles that once

tG

to

t.2 1.3 1.4

HELIOCENTRIC RANGE, astronomical units

DISCRETE

INTERVAL COUNT

R APHELION

ENCOUNTER

I15 1.6

FZGURE7-7.--Number of cosmic dust impacts during mission of Mariner IV as recorded by cosmic

dust detector.

246




belonged to the nuclei of comets that gradually dispersed in space, as small comets

are known to do.

2. Just beyond the orbit of Mars is the so-called asteroid belt, an area in

which many large asteroids and many thousands of small asteroids orbit the Sun;

collisions between the smaller asteroids would produce dust that would then grad-

ually approach the Sun.

Thus, Mariner IV flew toward a suspected source of cosmic dust.

Approximately 235 impacts were recorded during the mission. The number of

impacts for each 5-day interval is plotted as a function of heliocentric range in

figure 7-7. A comparison of the cosmic dust recordings of Mariner IV with those

of Mariner II gave the following results: Out to about 1.2 AU (i.e., 1.2 times the

distance from the center of the Sun to the center of the Earth), the flux of dust

recorded by Mariner IV was almost identical to that recorded by Mariner II in

traveling to 0.72 AU. Beyond 1.2 AU, the flux increased more or less steadily,

reaching a maximum at 1.38 AU, which corresponds to the perihelion, or near

point, of Mars' orbit to the Sun. The decline in flux beyond that distance suggests

that some of the dust once present has been swept up by the planet. No statistically

significant evidence of any well-defined streams of dust particles was found.

Ioniza)ion Chamber

A duplicate of an instrument flown to Venus by Mariner II, the Mariner IV

ionization chamber was designed to measure the average total ionization of fast-

moving particles that could penetrate a stainless-steel shell 0.025 cm (0.01 in.)

thick. To penetrate the shell, electrons had to have energies above 0.5 MeV, pro-

tons above 10 MeV, and alpha particles above 40 MeV. The chamber itself con-

sisted of a sphere, 12.7 cm in diameter (5 in.) filled with argon gas at a pressure of

4 atm. Particles that entered the sphere and passed through the gas knocked elec-

trons out of the argon atoms. The loss of one or several electrons meant that the

atom which was normally electrically neutral became ionized. The time required

to collect a preset amount of electrical charge resulting from this ionization wastelemetered to Earth, giving a measure of the ionization capability of space

radiation. A GM tube mounted nearby counted the number of ionization-causing

particles. By combining the time measurement with the particle count, much

could be deduced concerning the ionization capability of the average charged

particle that passed through the instrument. The weight of the instrument was 1.3

kg (2.9 lb), and 0.46 watt of power was required.

247



m

MARINER-MARS1964

The intensity of radiation within the Earth's magnetosphere and near the

equatorial plane was found to vary steadily with distance out to 8.2 Earth radii

(approximately 52 300 km (32 500 statute miles)); from 8.2 to 15.3 Earth radii

(approximately 97 500 km (60 000 statute miles)), the intensity varied complexly;

and, at greater distances from Earth, only interplanetary radiation levels were

observed. After a solar flare on February 5, the GM tube measured an 80-fold rise

in the number of interplanetary particles, and the ionization chamber readings

were 200 times larger than the normal value. The general variation of intensity

with time after the flare can be described by diffusion of particles in the inter-

planetary medium. In addition, there were later fluctuations which were ascribed

to the subsequent ejection of particles from the Sun or to modulation of particles in

the interplanetary medium.

The solar flare radiation of February 5 was apparently the cause of damage to

the GM tube; after February 10, this damage caused data from the tube to become

unusable. This failure, in turn, affected the power supply of the instrument which

was common to both the GM tube and the ionization chamber. After February 5,

the counting rate failed to return to the normal background level. Then, after

approximately 10 days of operation at a level 30 percent higher than normal, the

counting rate rose sharply from 40 counts/sec to 18 000 counts/sec and fluctuated

between 18 000 and 6000 counts/sec until March 17. On that date, the counting

rate dropped to zero and remained there permanently.

Since the same detectors were flown on both the Mariner II and Mariner IV

missions, it was possible to compare the data collected near Venus in 1962 with

those collected in 1964-1965. The omnidirectional flux and ionization rate

recorded by Mariner IV were about 40 percent higher than those recorded by

Mariner II. The variation of cosmic ray flux with time, as measured by Mariner

IV, correlates very well with similar observations made with high-altitude bal-

loons during the same 2-year period, as shown in figure 7-8. The somewhat lower

increase (24 percent) in flux given by the balloon measurements may be explained

by assuming the increase to be largely composed of low-energy particles which

were able to penetrate and thus were registered by the Mariner IV instrument,

but which were unable to penetrate the residual atmosphere above the balloon-borne instruments.

It also appears that cosmic ray intensity was greater during 1954 than during

1964-1965, both periods being of low solar activity. This is presumably because of

lower solar activity during the 1954 period, with the consequent production of

more low-energy particles.

248




575

550

525

5OO

o

o

E 475

E

'_w 450

.__

Z0

7- 425

Z

_o

400

375

350

I1_ ESTIMATED FROM MARINER £_" DATA OF FEB 16, 1965

=E--------ESTIMATED FROM MAR/NER.Z_ DATA OF DEC 3, i964

_9ALLOON DATA

THULE, GREENLAND

f AUG I0_ 1954

APR 16, 1965

FEB 16, 1965

DEC 3, 1964

JULY 28, 1963

B, 1962

325 I

0 50 I00 150

ATMOSPHERIC PRESSURE, g/era 2

FmURE 7-8.--Comparison of ionized particle counts taken by Mariner IV ionization chamber

experiment and those taken during high-altitude balloon flights.

249



II

t,

MARINER-MARS 1964

Cosmic Ray Telescope

Similar to instruments previously flown on the IMP satellites, the cosmic ray

telescope measured charged particles in the interplanetary environment and in

that of Mars. It could detect the cosmic rays which were also detected by the

ionization chamber and, in addition, particles which were not energetic enough to

cause ionization in the chamber. Weighing 1.17 kg (2.58 lb) and using 0.598 watt

of power, the instrument was composed of three solid-state detectors and absorbers

placed one above the other (like lenses in a telescope) to accept incoming particles

within an acceptance cone of 40 °. The depth of penetration of a particle and its

energy loss in the detectors determined the energy of the particle and identified

the particle. The measurement of these quantities was made by a 128-channel

analyzer.

Looking in the direction away from the Sun, the instrument was on the

shadowed side of the spacecraft during the cruise phase of the mission. It detected

protons in three ranges from 0.8 to 190 MeV and alpha particles in three ranges

from 2 to more than 320 MeV. The average observed cosmic ray intensity (protons

and alpha particles greater than 1 MeV) was 3 counts/min. The instrument re-

ported a complete absence of trapped electrons and protons at Mars and sub-

stantiated the magnetometer and trapped radiation detector data in setting an

upper limit on the Mars magnetic moment of 0.1 percent of the value for Earth.

Since the energy ranges for this experiment were the same as those of previous

experiments flown on IMP satellites, an excellent space-time correlation of events

was possible. During the Mariner IV mission, one moderately large solar-proton

event and several much smaller events associated with solar activity (i.e., solar

flares or 27-day recurring regions on the Sun) were observed by the cosmic ray

telescope. Almost all these events were observed simultaneously by IMP satellites.

The analysis of the time and space relationships between these simultaneous obser-

vations made possible several conclusions that would not have been evident from

measurements at a single point in interplanetary space. Among these conclusions

were the following:

1. Solar particles are strongly guided along, rather than across, the spiraling

interplanetary magnetic-field lines of force, as was shown in figure 7-1.

2. These particles are sometimes stored and carried around the Sun within

magnetic field structures rooted in the Sun.

New information on the spatial variation of galactic cosmic ray intensities

was also provided. This new information showed a much larger radial gradient

during solar minimum than had previously been anticipated. Therefore, very

250




significant changes were made in our understanding of the relative importance of

intergalactic magnetic fields and cosmic ray particulate energy density.

Trapped Radiation Detector

The trapped radiation detector on Mariner IV was designed to measure the

electron and proton radiation belts (Van Allen belts) around the Earth, similar

formations around Mars if they existed, and particle phenomena in interplanetary

space. The radiation belts consist of protons and electrons that are "trapped" by

the Earth's magnetic field. The trapped radiation detector instrument, weighing

0.98 kg (2.17 lb) and using 0.44 watt of power, consisted of four detectors and

associated shielding and electronics circuitry. The detectors comprise two GM

tubes (detectors) and one silicon solid-state surface-barrier proton detector whichpointed 70 ° away from the Sun-spacecraft line, and one GM tube which faced

135 ° away from the Sun-spacecraft line. All the detectors had acceptance-cone or

"look" angles of 60 ° . Since the instrument was very sensitive, electrons and pro-

tons of the energies shown in table 7-II could be detected. Each of the five record-

ing channels reported in rotation the total counts in a 45-see period.

Data taken near the Earth with the GM tubes clearly show effects during

traversal of the Earth's magnetosphere. These data, displayed in figure 7-9, pro-

vided a calibration of the capabilities of the system and a basis for the interpre-

tation of observations obtained later during the encounter of the spacecraft with

Mars. At a radial distance of 10.5 Earth radii (approximately 66 930 km (41 600

statute miles)), protons in the range of 0.5 to 11.0 MeV were no longer detectable.

Electrons with energy greater than 40 keV were detected continuously out to 23

Earth radii (approximately 146 400 km (91 000 statute miles); the fringe of the

magnetosphere), with an outlying intensity "spike" at 25.7 Earth radii (approxi-

mately 164 100 km (102 000 statute miles)).

During 7 days of flight in the presumed region of the Earth's magnetic tail

(at a geocentric distance of approximately 3300 Earth radii), the trapped radia-

tion detector failed to show any increase in electron density.

The average count rate during the Mariner IV mission for electrons withenergies greater than 40 keV and protons with energies greater than 550 keV was

0.7 count/see. For protons with energies between 0.5 and 11.0 MeV, the rate was

0.1 count/see. During the February 5 solar flare, the maximum count rate oc-

curred: 60 counts/see for electrons with energies greater than or equal to 40 keV

and for protons with energies greater than 550 keV; and 9 counts/see for protons

with energies between 0.5 and 11.0 MeV.

251



MARINER-MARS 1964

Table 7-11.mTrapped radiation detector energy range

Detector

GM tube A (1 recording channel) ...............

GM tube B (1 recording channel) ...............GM tube C (1 recording channel) ...............

Surface-barrier proton detector D (2 recordingchannels: D1 and D2)

Electrons

> 45

> 40

> 150

None

Energy range, keV

Protons

670 _30

> 550 ±20

3100

500 to 11 000; 880 to 4000

The data shown in figure 7-10, obtained during the approach and encounter

with Mars, contrast sharply with those obtained in the vicinity of the Earth just

after launch. The same detectors which were able to detect electrons of energy

greater than 40 keV out to a radial distance of 23 Earth radii failed to detect any

such electrons during the close approach to Mars at a radial distance of 13 270 km

(8250 statute miles) from the planet.

A new type of energetic particle event was reported; namely, the impulsive

emission of _40 keV electrons from the Sun. These electrons were actually

observed on three occasions. They had a steeply falling energy spectrum, isotopic

angular distribution, and (as seen from Earth) they seemed to be associated with

radio bursts and X-rays from the Sun. Because of the low relative mass of the

electrons compared with that of protons, this may become a new tool for investi-

gating the interplanetary medium since the electron gyroradius is only about10 -_ AU.

Figure 7-11 is a special polar projection showing the encounter geometry of

Mariner IV in relation to the hypothetical magnetopause and shock front and

assuming that Mars had a magnetic moment 0.1 percent that of the Earth. The

crossed circles on the diagram are the points in space where electron intensities

should have been the same as those at 23 and 25.7 Earth radii from Earth, if the

magnetic moment of Mars were 0.1 percent that of the Earth (provided it is

assumed that the same physical processes leading to acceleration and trapping of

electrons in the Earth's magnetic field would be found in Mars' magnetic field).

Since no electrons of that intensity were detected, it may be concluded that the

magnetic moment of Mars is surely less than 0.1 percent that of the Earth and

probably less than 0.05 percent. Corresponding upper limits on the equatorial

magnetic field at the surface of Mars are 200 and 100 gammas, respectively, as

contrasted with a field at the Earth's surface of about 50 000 gammas. These find-

ings indicate the possibility that the solar wind interacts directly with the Martian

252




e,,

i.-z

§

io5

10 4

10 3

io2

ioo

io s

io4

io3

102

I0 I

I0 o

" I I" A,'_RENTO_N_rS_

IOa DETECTOR C 1 I

103104 _l _ COUNTSI0 z

II0 "1 I I I I I I I I I I I 1 I I I 1 I I I I 1 I

O IO 15 20 15 30

GEOCENTRIC DISTANCE, Eorth rodii

FIGURE 7-9.--Trapped radiation detector data taken during passage of

Mariner IV through magnetosphere of Earth.

253



MARINER-MARS 1964

2:50 200 170

1.0 D ET EC TO R =A • I •

. • Oo,}_ .;•,'.-°"0,5 % "° ":* " _I""

DISTANCE FROM CENTER OF MARS, km x 105

140 I10 80 50 20 13 20 50 80 I I0 140 170

i i, • i= " = = CLOSEST' APPROACH OC_CULTATI(31_I = • •

-' .:._:: N-i ' :' '...o .... . _• .. .o°° o? £,_ •° °•_ t "•,- .o •.. I,Zo•.O° o ." °o°° °°• ..

1.0 I

DETECTOR C • ••. . ,. ,..

• • * '• ° °_ .... i .

" -._° • .=

o_ ""'" ""_"""........ " "" ":""......:-"

@

u_

"E

(2:

I-

.'.... : .

I0° - DETECTOR O i .,, .. • .. ....... %

-°

i0-o _.._.'.V. ,,, ,; ...._

,_ ° • °

=. ,°

10-2

I00 :.- DE TECTOR D 2

I .-'-.--..:-_.: ,.r0 -t ._.:..'. . . . . ".. .... - - :- ":," ".." • . ..;.'. " . ..., .'-- ...'. -]

...... ,i.°°_°o ° .

I0 -2 Jl I | I l I l l I t I I I I

I0:00 13:00 16:00 19:00 22:00 01:00 04:00 07:00 I0:00 13:00

JULY 14 _ JULY IS

GMT AT EARTH

FmuRE 7-10.--Trapped radiation detector data taken during encounter of Mariner IV with Mars.

atmosphere and that the production of secondary particles by high-energy cosmic

rays occurs below the surface of the planet.

Solar Plasma Probe

The solar plasma probe was included in the science payload to measure the

density, velocity, and direction of the charged particles making up the solar wind.

In the instrument, the energies of protons and alpha particles were assigned to any

of 32 energy bands (logarithmically related), ranging from 30 to 10 000 eV.

Weighing 2.91 kg (6.41 lb) and using 2.65 W of power, the instrument was

_D4




_km X IO3

_UN- _o ,4/ ," _ ,o _o _o ,o _o _o _o- , \\,,, J , , , , , , ,

TOPAUSE,o--<.._MARINER J_'SHOCKRONT

3o __4O

POINTS WHERE ELECTRON

INTENSITIES SHOULDOUAL THOSE AT 23

AND 25.7 Eorth rodil

50 FROM THE EARTH, ASSUMING

MAGNETIC MOMENT OF MARS IS0.1% THAT OF THE EARTH

FIGURE 7-11.--Relationship between Mariner IV trajectory and hypothetical

magnetopause and shock front at Mars, assuming magnetic moment of Mars

to be 0.1 percent that of the Earth.

mounted on the spacecraft facing 10 ° from the Sun-spacecraft line with a conical

field of view of 60 ° . Three equal pie-shaped sectors were designed to provide direc-

tional information on the flow of the solar wind.

During Mariner IV's first day in space, the solar plasma probe recorded the

passages of the spacecraft through the Earth's shock front; good correlation with

the magnetometer data was obtained. From November 29 to December 4, the

solar wind density ranged from about 2 to 12 parficles/cm3; the velocity of the

solar wind during this period ranged from approximately 275 to 425 km/sec (171

to 264 statute miles/sec).

Unfortunately, because of failure of a high-voltage resistor in the power

supply 8 days after launch (which caused the solar plasma probe to sweep the

255



MARINER-MARS 1964

energy spectrum in an unknown way), the data obtained at the high data trans-

mission rate (33x/_ bits/sec) during most of December were impossible to interpret.

Analysis of the failure allowed some recalibration of the instrument, and, when the

low data transmission rate (81_ bits/sec) began in January, it was again possible

to obtain meaningful measurements.

Solar winds observed during the first 3 months of 1965 varied in velocity from

a little less than 300 km/sec (186 statute miles/sec) to a maximum of about 550

kin/see (342 statute miles/sec). The 3-hour averages of proton flux, density, and

bulk velocity in the data sampled from January 18 to 26 are shown in figure 7-12.

This sample exhibits a general feature noticed throughout the flight: the tendency

for density increases to precede velocity increases. The same effect was noted

using data from the solar plasma probe carried on the Mariner II spacecraft to

Venus. Longer periods of the data show a repetitive structure in plasma flux,density, and velocity.

The solar plasma probe data obtained during the Mariner IV encounter with

Mars agree with the results obtained by the helium magnetometer and the trapped

radiation detector: there was no clear indication that any magnetic fields from

IO

a. " o I

z).'E

_

Zo__0_ E

a.

26$ iS 20 21 22 23 24 25 26 27

DAY OF JANUARY 1965

FIOURE 7-12.--Solar plasma probe data from January 18 to 26, 1965.




Mars were influencing the plasma flow. At encounter, the velocity of the solar

wind was 330 km/sec (205 statute miles/see), with a density of 0.8 particle/cm 3.

TELEVISION AND OCCULTATION EXPERIMENTS

Introduction

Certain physical properties of the Earth and Mars are listed for reference in

table 7-IIi. Mars approaches close to the Earth every 25 months at opposition

(when the Earth and Mars are in line with the Sun in the same direction from it).

Even at opposition, Mars may be as far as 99 million km (62 million statute miles)

away from the Earth. It is never closer than about 56 million km (35 million

statute miles), at which time the diameter of Mars appears to be about one-

seventieth of the diameter of the Moon as seen from the Earth.

On Mars, it is possible to perceive permanent surface markings. Bright areas

are concentrated in the northern hemisphere; dark areas, in the southern hemi-

sphere; and white areas near the poles. The material covering the bright areas is

thought to be fine dust, since dust storms of the same color are frequently observed.

Concerning the white areas, photographs show clearly that something re-

sembling an ice cap first forms on one pole and then on the other as the inclination

of the axis of the planet to the plane of its orbit around the Sun produces winter

and summer seasons (scaled to the 687-Earth-day Martian year). The polar caps

Table 7-111.--Physical characteristics of Earth and Mars

Characteristic Earth Mars

Mass:

kg ................................................

Ib .................................................

Diameter:

Km ................................................

Statute miles ........................................

Average density:

g/cm s ..............................................

lb/in, s .............................................

Surface gravitat ional acceleration:cm/sec 2............................................

in./sec 2............................................

Mean distance from Sun:

Km ...............................................

Statute miles .......................................

Angle between rotational axes and orbital plane, deg .........

Length of year, Earth days ... ... .. ... .. ... .. .. .. .. ... .. ..

Length of day, hr .......................................

12 870

8000

5.54

0.20

980

386

151X106

93.75;<106

23.5

365

24

0.64X1024

0.14X10_8

6840

4250

3.88

0.14

371

146

229 XIO n

142.5X10_

24.5

687

24.6

257



MARINER-MARS1964

slowly disappear with the coming of spring on Mars. The atmosphere of the planet

is exceedingly rarefied (as verified by the Mariner IV occultation experiment

discussed later in this section), and scientists do not believe that the planet can

contain enough water vapor to give rise to polar caps of this nature. These polar

caps may consist of frozen carbon dioxide.

The dark areas are probably of most interest, since their appearance and

behavior have prompted many to propose that there is life on Mars. They consist

of three distinct regions: maria (seas) ; canals; and oases (where several canals may

intersect). (These designations are still in popular use, even though they are not

applied in the same sense to Mars as they are to Earth.) The most prominent

surface feature is a mare called Syrtis Major, which in shape and location some-

what resembles the terrestrial subcontinent of India. With the changing seasons,there are also apparent changes in the coloration of the dark areas such as Syrtis

Major. The coloration of these areas has been debated at considerable length.

Temperatures in the dark areas are believed to be somewhat lower than those in

the bright areas.

The canals, appearing as straight-line markings on observational maps, are

undoubtedly the most widely discussed features of Mars. Two and one-half cen-

turies after Galileo had been barely able to distinguish the disk of Mars in the first

astronomical telescope, Schiaparelli, working at Milan Observatory in 1877,

noted these surface features and called them "canali" ("canals" or "channels")

since they were dark and seemed to reach across the "lands" from "sea" to "sea."

Figure 7-13 shows Schiaparelli's map of the canals he saw on Mars. However, not

everyone saw the canals to which Schiaparelli and later an American, Percival

Lowell, referred. Lowell was so convinced of the implications in the word "canal"

that he founded an observatory where initial activity centered on these "inland

waterways" of Mars. Lowell's map of Mars showed what de Vaucouleurs de-

scribed as a "veritable cobweb" of canals. Books published by Lowell contained

speculations about irrigation from the polar caps and led to considerable reaction

both by the scientific community and by the public as well. Although recent

photographs have helped resolve some of the questions by apparently showingsome large-scale canallike features, these photographs are subject to interpretation

and discordant views persist.

Important properties of the atmosphere of Mars include pressure, composi-

tion, particulate content, and the presence of clouds. Early estimates of the

surface pressure were drawn principally from photometric and polarization obser-

vations after making certain assumptions about the composition of the atmosphere

258



b

L SC/ENT/F/C RESULTS AND CONCLUSIONS

259



MARINER-MARS1964

and the reflecting properties of the planet's surface. After evaluating various

estimates, de Vaucouleurs gave a value for the surface pressure of 854-4 millibars

(as compared with the Earth's sea-level pressure of about 1013 millibars). This

value was in general use until 1963, at which time a new technique, based on

infrared spectroscopy, was applied to the problem. According to the theory used

in this technique, the spectral absorption lines of the carbon dioxide on Mars are

broadened or smeared as a function of gas pressure. Therefore, comparison of the

spectra obtained on Mars with carbon dioxide spectra obtained at various pres-

sures in the laboratory should give, after much analysis, the pressure value for

Mars. This technique produced a value of 254-15 millibars. Several programs to

confirm and refine this low value produced other values, but all were of low

magnitude when compared with de Vaucouleurs' value.

The atmospheric pressure at the surface is a function of the total amount andkinds of gases in the atmosphere (given Mars' gravitational field, which is 38

percent that of Earth's). The first molecule to be identified in the atmosphere of

Mars was carbon dioxide. The only other molecule detected in the gas phase was

water. For the other possible constituents (e.g., argon and nitrogen), only upper

limits are available.

The existence of a somewhat permanent load of particles in the atmosphere of

Mars was suspected from the study of photographs. The most likely particles were

ice or carbon dioxide crystals and dust particles with diameters in the submicron

range. As the estimated amount of carbon dioxide on Mars has increased over the

years, the probability that the particles are composed of carbon dioxide has also

increased.

With Mars' small gravitational field, the decrease of atmospheric pressure

with height is less rapid than that for the Earth. Three cloud colors have been

observed: blue, white, and yellow. Yellow clouds, almost certainly due to dust

storms, are frequently observed at altitudes from 4.8 to 9.6 km (3 to 6 statute miles)

and occasionally much higher. These clouds have been observed to move as fast as

137 km/hr (85 statute miles/hr), though more often in the 32- to 48-km/hr (20- to

30-statute-mile/hr) range. Wind velocities necessary to pick up the dust may

range as high as 201 km/hr (125 statute miles/hr) at a 91-meter (300-ft) altitude.There seems to be little doubt that the blue and white clouds are composed of

crystalline water or carbon dioxide. Recently it was suggested that the white

clouds occur most frequently over bright areas and adjacent to dark areas. It was

proposed that these clouds are analogous to clouds produced on the Earth when

moisture-laden winds blow at right angles over a mountain range. Downwind of

260




the range are regions where the air is at a low pressure, and adiabatic expansion

results in sufficient cooling to permit condensation. If this process is occurring on

Mars, it implies that at least some of the dark areas are elevated and that the winds

blow from the dark to the bright areas during the times when clouds are observed.

From this brief discussion on Mars, it is obvious that, although Mars is prob-

ably the best observed of the planets, the questions concerning it that remain are

many and are not easily answered because of the limitations of Earth-based obser-

vational techniques. Certainly a closeup view of the planet would do much in

providing a more accurate picture of its surface characteristics. Therefore, when

the Mariner-Mars 1964 scientific payload was selected, a television experiment

was included to help resolve some of the controversies surrounding the nature of

the planet's surface. The picture-taking sequence was designed to obtain a view ofa wide variety of Martian features. The best telescopic resolution from Earth of the

planet's surface, combining visual and photographic observations, generally is

believed to permit distinction of features no smaller than 80 to 97 km (50 to 60

statute miles) across. A resolution of 2.4 km (1.5 statute miles) was expected with

the television instrument included in the Mariner-Mars 1964 scientific payload.

In order to obtain more information concerning the atmosphere of Mars, an

occultation experiment was included. This experiment, requiring no added equip-

ment on the spacecraft, was based on the principle that, if the spacecraft passed

behind the planet as viewed from the Earth, its radio signal would pass through

the atmosphere of Mars. Any amplitude and phase changes detected in the signal

would enable scientists to draw conclusions regarding the atmospheric charac-

teristics of the planet.

The scientific investigators for the television and occultation experiments

were given previously in table 7-I.

Television

Description of experiment

The television experiment on the Mariner IV spacecraft had three basic

objectives: (1) to make preliminary topographic reconnaissance of the surface ofMars; (2) to attain, through photographs, an improved knowledge of areas of

possible living matter; and (3) to obtain additional information concerning the

planet's surface reflectivities to aid in the design of future, more detailed photo-

graphic experiments. The instrument selected to achieve these objectives con-

sisted of one camera and a television system that recorded 40 000 (200 by 200)

picture elements per frame. The camera had a focal length of 30.48 cm (12 in.)

261



MARINER-MARS 1964

9

8

and could resolve surface features of 3 km (1.88 statute miies) at the Mariner IV

passage distance. The shutter speed had to be 0.2 sec or less to limit blurring of the

image caused by the motion of the spacecraft with respect to Mars. The light

sensitivity of the television picture tube established the focal ratio off/8. For the

optical system, a reflecting telescope of the Cassegrain type was selected, with an

aperture of 3.81 cm (1.5 in .) . The television system is shown in figure 7-14.Experience had demonstrated that the best way to send a weak radio signal

through space in the presence of background noise is to use a signaling method

known as pulse-code modulation. With this method, the output of an electronic

device such as a television camera is coded into a sequence of binary digits (bits) of

zeros and ones that represent a particular level of intensity. The output of the

Mariner IV television camera was translated into a 6-bit code that identified thebrightness of each picture element on a scale that had 64 steps from full black to

full white. The 64 steps of the sequence ran from 0 to 63. A sequence of six ones

represented M i biack or no iight a t aii; a sequence of six zeros represented h i i

w-hite or maximum light. To encode the information contained in the 40 000 pic-

ture elements, 240 000 bits were required.

In order to obtain information about the surface coloration of Mars, the

television system was designed to take overlapping pair? of pictures, with one

member of each pair being taken through a green filter and the other through a

7CAMERA

SHUTTER AN D FILTERS VIDICON TUBEMIRRORS

FIGURE-1 4.-Mariner IV television system.

262




filter. A wheel carrying four filters, alternately red and green, was arranged to

ate 90 ° after each exposure, thus producing a sequence of pictures alternately

and green. One picture was exposed every 48 see. Had all these pictures been

orded, however, all the data-storage capacity would have been used up long

fore the television scan path had crossed the planet. In order to stretch out the

uence and yet have some pairs of overlapping colored pictures, therefore, every

rd picture was omitted from the stored sequence. The overlapping pairs of

tures thus followed a sequence of green-red, red-green, green-red, etc.Several hours before the actual recording sequence, the camera and tape

order electronics were turned on and began taking pictures. Since the tape

nsport was not yet activated, these first pictures were not recorded. During this

rmup period, the camera axis was positioned to trace the desired path across

planet. When the planet came into view, an internal signal activated the tape

nsport and recording began. The tape recorder was started and stopped at the

ginning and end of each picture to conserve tape. A total of 21 pictures and 22

es of the 22d picture was recorded.

Although the system was provided with automatic gain control to adjust foranges in the brightness of the surface of Mars, the gain adjustment could func-

n only after the first picture had been recorded on the face of the vidicon tube

d had been scanned electronically. Furthermore, in order to keep the gain

ntrol simple and not run the risk of a large error in correction between pictures,

e gain correction was made only on the basis of the green image and then was

ited to a gain change up or down of only one step. Pictures 1 to 18 were all at

e same gain level, which was selected as the one most likely to be correct before

ariner IV was launched. For the remaining pictures, the gain increased to the

xt highest level at each picture.

The television data were programed to be returned to Earth about l0 hours

ter the recording, with each picture requiring about 101_ hours for complete

nsmission. As the signals arrived, they were recorded on magnetic tape for sub-

quent analysis; this also provided a permanent record. After the 21 pictures and

e 22 lines of the 22d picture had been transmitted to Earth and recorded once, a

ocess that took slightly more than 8 days, Mariner IV was instructed to retrans-

it the entire set. It was of interest to see how closely a replay would duplicate the

itial values for the 40 000 picture elements in each of the 21 pictures plus the 22

es of the 22d picture. Any discrepancies between the two playbacks would

dicate the number of errors that had occurred in the transmission and would also

ll where they had occurred in each picture. The second transmission differed



MARINER-MARS1964

from the first in only about 20 elements of the 40 000 in each picture, making an

average of 10 errors per picture in each transmission.

Regionsof Mars photographed

Prior to launch of Mariner IV, it was known that a maximum of I percent of

Mars' entire surface could be photographed if everything on the spacecraft func-

tioned perfectly. The best scan paths were those that crossed the largest number of

bright and dark areas. Naturally, pictures of regions in which canallike markings

had been consistently reported were desired. Also, it was felt that some of the pic-

tures should include views of the side of the planet on which Syrtis Major is located.

In selecting the flight path, however, the needs of the television instrument

were not the only ones that had to be considered. For example, the spacecraftcould not be allowed to enter into the shadow of Mars or it would lose its fix on the

Sun. It could not go above Mars or it would lose its fix on Canopus. In addition,

the spacecraft had to travel behind Mars so that its radio signal would pass

through the atmosphere of the planet for the occultation experiment. But the final

requirement was the one that placed such a restriction on the flight path that,

during the Mariner IV mission, it proved impossible to have the camera pointing

anywhere near Syrtis Major. This requirement was that California be facing Mars

at the time of encounter so that the 100-kW transmitter at Venus DSS 13 in

California could send last-minute commands to the spacecraft if necessary.

The desired time of encounter was achieved by making appropriate adjust-

ments in the trajectory of the spacecraft during the December 5 midcourse

maneuver. This maneuver could have adjusted the encounter time to any desired

value over a period of several days; however, because the Earth and Mars rotate

on their axes at nearly the same rate, it was impossible to delay or accelerate the

encounter sufficiently for Syrtis Major to be facing Mariner IV's camera at the

same time that California was facing Mars.

The camera scan path foIIowed by Mariner IV at Mars started on the limb

(edge) of the planet at about 37 ° north latitude, swept southward across the

equator to about 52 ° south latitude, then curved northward again, and moved off

the planet at about 38 ° south latitude. The path crossed a region in which maps of

Mars show only a few canallike markings.

Table 7-IV and figure 7-15 show the coordinates for each of the 22 pictures.

It has been estimated that these locations are accurate to 1 o latitude and 3 °

longitude near the center of the set. The longitude uncertainty is greater for the

first few pictures and the uncertainty in both longitude and latitude increases to

264




Table 7-1V.--Martian coordinatesof Mariner IV photographs

Picture North latitude, deg East longitude, deg

1

2

3

4

5

6

7

8

9

10

11

12

1314

15

16

17

18

19

20

21

22

37.0

25.1

11.9

6.9

--1.8

--5.8

--13.5

--17.1

--23.7

--26.8

--32.7

--35.4

--40.1--42.2

--46.0

--47.6

--50.1

--51.0

--51.2

--50.2

--44.3

--35.8

173.0

175.0

177.8

179.1

181.5

182.9

185.8

187.4

190.9

192.9

197.3

200.0

205.6208.9

216.5

220.9

231.3

237.7

252.1

260.2

278.5

291.4

perhaps 4 o or 5 o for the last few pictures. These locations were derived by com-

bining information from the following sources: (1) trajectory; (2) orientation of

the spacecraft within the limits of the attitude-control subsystem; (3) position of

the scan axis with respect to the spacecraft; (4) position and orientation of the

limb of Mars in picture 1 ; and (5) motion observed by examination of the overlap

between adjacent pictures.

Data processing and reduction

The first and second playbacks were merged, and the bad lines were elimi-

nated. Discrepancies between individual pictures were repaired by hand rather

than by computer. Because the percentage of bit errors was very small and the

scene itself was one of low contrast, the errors that did exist were easily identified

and corrected in most instances. This was done by fitting the erroneous point to

the surrounding scene by correcting one of the 6 bits in the intensity word. A

similar averaging technique was used to remove the black fiducial marks.

The fiducial marks provided a means of rectifying the geometry in the pic-

tures. Before launch, the geometric distortions of the camera and optical systems

were recorded by photographing a grid pattern and noting its relationship to the

265



MARINER-MARS 7964

.

I "I

50"

N.

40'

30'

!OO

IO0

3 O

IO "

29"

3G"

40°

S.

5 d

FIGURE-1 5.-Locations on Mars photographed by Mariner IV

266




fiducial marks. The geometric distortions in the Mariner IV pictures were not

great, but were sufficient that in the application of calibration data it was neces-

sary first to remove the small amount of nonlinearity that did exist. Thus, by usingdigital computer operations, geometric fidelity was achieved. Calibration data

could then be incorporated.

The major correction to the data was compensation for shading on the

vidicon due to varying sensitivities over the photoconductive surface. Because the

pictures were of very low contrast, the shading corrections became extremely im-

portant, even though the vidicon shading was not great. The shading corrections

were made by using a large number of calibration frames. These frames were

obtained by exposing the camera to a uniform scene whose illuminance had been

accurately measured and was varied over the dynamic range of the systems. From

this set of calibration frames, a calibration matrix was constructed to give theintensity corresponding to a given data number and gain setting for each of the

40 000 positions of the picture elements. This extensive process increased confi-

dence in the stability and photometric accuracy of the vidicon system.

High-altitude haze which appeared in picture 1 raised the question of glare in

the optical system. Tests made in an attempt to create an optical defect that could

have caused the observed effects revealed no reasonable possibility that the optics

had degraded.

The early photographs were enhanced by a computer process which, in a

sense, stretched the darkest intensity value received to black and stretched the

brightest intensity value received in the other direction to nearly white. This

method of enhancement increased the degree of contrast between adjoining dots in

a photograph over that actually seen by the camera and made feature identifica-

tion easier.

One attempt to identify surface features was made by using an airbrush

technique to produce a relief rendition of picture 11. This rendition is presented in

figure 7-16. 2 The relief interpretations shown were b_sed on knowledge and

experience gained from interpreting and rendering lunar relief features. Picture 11

shows with reasonable clarity (even though contrast is enhanced by a factor of 4)

that the Martian surface is made up of the same kind of features as those found on

the lunar surface. Therefore, a maximum interpretation was made showing hexag-

onality on some craters, flat-floored craters, inverted cone-shaped craters, a dome

with caldera, and even linear (rill-like) or crater chain markings. A central peak

Prepared by Patricia Bridges of the ACIC (Aeronautical Chart and Information Center), Lowell Observatory,

Flagstaff, Ariz.

267



9

MARINER-MARS 1964

seems to show clearly in the crater a t the ==per left-hand cmner in figure ?-?6, yet

the very bright patch over the southwest wall of that crater is apparently obscuring

relief rather than being a surface feature with higher albedo (reflective power).That conclusion was reached when no logical form could be made of the bright

patch during attempts to interpret it as a surface feature.

In the most basic form, the Mariner IV photographic data are represented

m

268

FIGURE-1 6.-Relief rendition of picture 11




LINE 130

LINE 13[

LINE 13Z

LiNE [33

LINE 134

LINE 135

tl_E 136

LInE 137

LINE 13e

LINE 139

LINE 140

LINE [41

LiNE 14Z

LINE 143

LINE 144

tiNE 14S

tlNE 146

LI_E 147

UNE |48

LINE 149

LINE 150

LINE 152

LiNE 153

LINE 154

LINE ISS

LINE 156

LiNE 151

tlNE [5_

LINE L59

LINE 160

LINE 161

LINE LGZ

LINE [63

LiNE 164

LImE 165

LINE 166

LI_E 1_7

LINE 168

LINE 169

LiNE 170

LINE 171

LINE [72

LINE 173

LINE 174

tINE 175

LINE 176

LI_E 177

LINE 178

LINE 179

tlNE 180

LiNE [el

LINE 182

107 108 I09 llO III IIZ [13 114 115 116 117 lib Ii9 120 121 122 123 IZ4 125

71 71

71 71

7| 7Z

73 7|

72 72

73 72

73 73

73 69

72 72

7O 70

73 7O

73 69

77 ?0

73 7O

74 73

74 74

77 ?3

74 7_

TO 74

74 75

75

74

?8

77

77 75

7e 75

76 ?_

?5 75

76 16

76 16

77 76

77 77

77 t7

77 77

?8 78

7? ?8

78 81

7e 82

79 ?9

8o eo

79 79

76 _9

76 76

75 ?6

77 77

78 77

81 8O

81 8O

_Z el

78 ?8

78 ?e

;8 78

79 78

72 69 74 T3 77 75 75 75 7B 78 78

6q 73 73 74 7. 78 74 79 78 79 78

72 ?0 7O 74 ?_ 79 79 76 76 78 7e

68 69 ?4 ?_ 71 76 76 78 7g 79 7e

7z 69 71 75 75 7_ 79 8o 79 79 7e

72 7o 7o 75 76 77 79 17 t6 78 7g

?3 73 71 ?5 ?7 7e 7q 76 76 8o 79

73 73 71 7Z 75 75 76 T8 77 76 76

7O 7) 7_ 71 76 77 76 17 8O ?_ 79

73 73 74 7[ 77 73 71 16 8O 74 79

73 71 74 74 76 72 72 76 76 T9 79

?4 ?1 15 ?5 72 75 72 77 eO T9 7e

7o 71 t4 71 7Z 72 76 7_ 8o 7o 7e

71 75 7O 7O 71 7Z 16 76 79 79 79

75 74 7Z 71 75 76 7? 8O r9 t6 76

75 74 74 71 75 75 76 8O 77 73 68

7. 74 74 ?! 75 75 eo _7 69 66 69

7_ 78 7_ 75 75 76 76 7z 66 65 69

74 75 79 75 ?4 ?8 7z 65 66 69 66

7_ 75 76 75 75 75 6q 66 66 66 66

74 t5 78 75 75 75 72 65 65 66 6_ 66

7_ 7_ 76 7_ T8 7z 68 66 6_ _ 6_ 65

74 75 76 76 76 69 68 61 62 o5 65 69

75 7_ 76 76 76 69 66 6[ 6_ 66 70 7_

75 75 76 74 69 67 62 67 68 70 7_

76 75 77 73 68 70 62 70 70 70 74

76 76 76 77 68 66 66 TO 69 73 ?e

76 ?6 76 73 72 67 67 73 73 11 ?7

75 73 77 76 1_ 66 7O _0 7O r4 77

76 73 77 76 74 7O 7l 7O 7o 73 77

77 ?7 74 77 ?e 71 7o 7O 7o 74 75

?6 76 77 77 7_ 67 71 I0 7o 69 ?3

73 ?8 17 74 71 61 66 71 70 7o 73

77 8o 78 ?s ?4 7o 68 67 7o 10 69

T6 76 78 ?5 7_ 74 7o _7 71 7o 7o

82 78 74 74 ?4 77 74 71 7O 7O 7O

7e 7e 79 71 78 75 74 14 74 14 7o

78 75 75 ?4 78 74 ?S 79 T8 7_ 74

79 ?e 71 74 7e 79 7_ ?_ 14 74 77

75 79 75 75 75 15 8O 75 78 78 77

75 7q 78 79 74 78 75 75 79 7_ 7_

t5 79 19 75 75 74 7S 15 79 75 74

78 79 75 75 79 75 7_ ?e 78 ?_ _5

77 77 78 79 79 75 74 7_ ?_ 74 7_

77 77 77 77 ?9 ?e 74 74 T4 T[ 74

77 ?7 77 77 79 79 1_ 74 _1 7_ 70

79 ?e 78 77 79 79 19 71 70 10 74

79 ?9 78 77 78 18 ?s T_ 70 TO 75

e0 79 _8 77 8O ?8 ?8 7o 7o 74 73

?g 78 77 17 78 7g 75 7o 7_ 69 73

78 78 78 77 79 78 75 74 7o ?0 t3

78 78 78 78 76 79 75 74 71 70 74

76 77 76 76 7? 77 ?S 7_ 71 71 72

FIGURE 7--17.--Unrectified

7B 7B 77 81 77 77

78 7B 7T 80 77 76

1g ?e 7B 7B 78 77

' 78 79 78 78 78 77

78 78 77 81 78 78

79 76 78 78 78 78

7Q 79 76 78 78 75

79 79 79 79 79 79

?q 79 75 75 79 76

19 78 7g 76 79 79

79 79 79 79 78 79

79 76 76 79 78 79

76 76 76 75 T5 79

76 76 7_ 7_ 75 ?9

72 72 ?Z 76 7Q 19

69 73 7_ 76 80 79

68 68 7Z 75 76 79

66 6q _8 ?Z 76 76

_6 6_ 65 6q 68 12

66 6_ _5 6_ _ 72

66 65 65 66 69 72

69 6_ _9 7_ 68 _

73 6_ 73 7_ _q 7_

73 70 7_ 72 7_ 73

73 7! 7_ 73 75 75

74 7_ 7_ 72 13 73

78 74 73 73 7) 74

7e 73 73 73 ?O 71

74 74 7_ 70 71 70

?3 7_ 74 70 71 71

74 7" 74 70 70 70

_3 7_ ?_ 70 ?0 74

73 74 78 74 ?1 71

73 78 ?o 78 75 71

71 78 _ 78 75 t[

7o 7_ 6_ 7_ 75 75

75 ?_ 71 70 74 _

?4 14 70 67 ti 75

77 ?_ 7_ 7Z _e 70

73 74 75 ?l 72 75

77 74 71 7Z ?s t_

77 ?4 7| 76 76 79

73 75 76 7_ 76 76

74 75 75 75 75 76

73 75 76 75 76 7_

70 71 71 75 76 76

74 75 75 75 7_ 7_

74 75 75 76 76 77

74 75 77 73 76 77

7_ 76 77 77 77 77

126 L27 IZ8 12Q l)O 131 132 133

77 78 78 78 79 7q ?Z 69

77 ?e 79 82 79 79 76 69

78 7_ 79 8_ 80 80 76 69

14 7_ ?g 83 79 8O 76 69

78 79 8O eo 79 83 77 7_

?e ?9 oo 8o 79 8O eL 73

79 79 8O 8O 8O 8O 78 74

16 ?_ 8O eo 8o 8O 81 78

?g 19 79 8O 77 77 81 ??

7_ 79 79 8O 77 71 77 ??

16 79 eo eo 77 77 77 77

76 77 76 8o 8o 77 77 T_

78 77 76 77 eo 77 77 7?

eo eo 8o eo 78 7e 7e ?e

8o eo oo eo el 78 ?8 7_

eo 79 eo 6l 81 eL 74 1_

8O 81 8O 8O 81 81 82 ?9

8O el _0 eL O_ 81 85 83

77 77 77 ?e e2 85 85 e6

76 ?4 77 78 82 82 e5 o_

?2 73 74 78 78 85 e6 o6

7_ 74 74 76 79 79 79 _?

7_ 7o 74 77 79 79 82 87

73 7o 72 7_ ?9 el o3 85

73 7o 10 74 79 83 e3 84

74 71 7_ 7e 7e e_ eo 83

71 70 75 14 eo oo 8o 8_

75 14 74 74 eo 8o eo ez

75 7_ 74 7_ 79 80 8o 83

72 75 75 ?e 79 8o lq 8_

72 ts ?s 74 18 8o e4 Bo

75 79 79 74 79 e3 8_ 8c

7_ 80 ?5 74 78 84 76 76

T[ _L 76 79 8_ 83 75 72

72 ?3 79 79 83 8O 73 72

71 _1 _0 83 8O 77 73 7_

76 t6 79 8O 76 73 ?3 t_

7_ 79 79 8O 76 76 77 73

76 8O eo 76 ?6 8O 77 74

75 8o 7_ 76 16 76 72 77

76 8o 76 76 76 76 7z 77

77 8O 60 77 77 77 77 76

76 uo ?6 72 76 76 76 ?7

8O 8o 8o ?6 76 76 77 76

eo 81 el 8o 76 76 77 71

78 60 eo 8o 81 77 71 77

77 el 81 8o 8o 7? 77 _7

81 81 ol 80 8o 76 77 77

eo 7o 81 eo 8o 76 73 77

77 76 81 _1 77 77 ?z 76

7Z 77 76 7t 77 78 8l 77 el B_ eL ?e 74 77

73 77 7? 77 78 79 78 77 82 8Z oz 7g 74 78

calibrated intensities for picture 11.

134 135 136 [37 138 139 140 14!

65 69 69 72 72 77 77 76

69 69 68 72 7l 16 77 77

69 69 72 73 7Z 12 76 76

69 73 7Z 72 7Z 72 76 ?6

69 73 ?2 76 76 13 76 77

69 73 71 75 72 76 ?3 77

69 72 73 72 76 ?Z 76 77

7) 73 69 7_ 73 73 76 77

77 7) _Z 69 73 ?3 73 74

78 78 73 73 ?3 74 74 10

?e 78 77 7_ 74 73 74 73

le ?e 17 73 74 74 75 ?4

79 78 17 74 74 74 7_ 7_

79 79 78 7e 77 74 14 74

79 19 79 79 ?_ T4 7_ ?7

79 78 77 78 79 74 74 73

79 ?e 78 78 79 78 74 78

79 78 81 81 ?9 78 75 74

eo ao 79 79 e3 Te 75 75

e7 eL 16 83 79 7_ 78 74

O5 8O _0 8O 7_ 7_ _ _S

_1 e4 79 e3 eo 7e 8o 8o

87 8_ oo 7_ 6o 8_ 1_ eo

89 ez T8 79 81 ez eo eo

91 8o 76 eo 84 8o 8O 8o

91 ao 77 81 8O oo eo eo

91 80 76 79 e3 eo 77 8o

87 76 _6 7q 8O eo 77 81

8_ 77 81 8O 76 ao 8o 81

81 81 77 8O 76 8O 77 77

8o 71 t6 84 al eo 77 77

8o 81 8[ 81 81 eL 77 77

77 81 71 77 78 82 ez 78

76 81 Te 77 77 82 e2 81

77 76 t? 77 78 78 e2 77

17 eo uo 76 77 ez 82 79

77 el eo 77 78 ?e 74 7e

78 81 e5 81 78 _8 7e 19

79 8z 8o e2 ?g ?e e2 7_

?8 ?7 81 81 78 78 79 7e

76 77 eL 8_ ?e 79 ?e 79

77 81 el 78 70 78 7_ ?9

76 _1 _2 7e 79 78 79 79

77 77 el 78 7_ ?9 79 75

73 1_ 6[ 78 75 79 79 79

1_ ?T 8_ 7e 7q 78 ?9 75

73 77 77 ?e 78 79 eo 75

77 77 7e 78 79 75 8o 8o

81 7_ 77 78 79 79 76 eo

el 7t 77 78 79 79 ?6 79

78 ?e 77 ?8 79 eo 71 ?6

78 77 7e 78 ?5 76 75 76

269



MARINER-MARS1964

by a numerical matrix for scientific analysis. An example of this presentation is

shown in figure 7-17. The crater area covered by this matrix is approximately

outlined by the rectangle drawn in figure 7-16. The numbers represent calibrated

intensities of the picture elements observed by the Mariner IV camera system.

Their values have been normalized such that 1000 is the observed intensity if the

Martian surface plane corresponding to each picture element consisted of a per-

fectly diffusing white surface oriented normal to the Sun. The absence of in-

tensity, or black, is represented by the number 0, and the intervening numbers

from 0 to 1000 indicate progressively lighter shades of gray.

During picture processing, compensation also had to be made for the reregis-

tration of the image format caused by operation in a region with no magnetic

field. All the camera ground testing had been performed in the Earth's magnetic

field, whereas the actual operation during Mars encounter occurred in essentiallya zero field. Because of this difference in environment, the image on the photocon-

ductive surface was translated and slightly rotated from the fixed reference of the

fiducial marks. Although the necessary reregistration was straightforward, a minor

problem developed because the last six picture elements in each television line

were in a portion of the photoconductive surface that had not been photometrically

calibrated. A reasonably good correction was made by extrapolation.

Descriptionof pictures

The entire set of pictures taken by Mariner IV is presented in figures 7-18 to

7-39 and is discussed in the following paragraphs. Figures 7-18 to 7-36 each shows

three representations of a particular Mariner IV photograph: (a) the "raw" data,

printed so that the picture-to-picture relative brightness levels as viewed by the

camera are shown; (b) the calibrated, reregistered data with errors and fiducial

marks removed; and (c) the same data as in (b), but with enhanced contrast.

Figures 7-37 to 7-39 show only the raw data. The "red" and "green" designations

in the descriptions which follow refer to the filter through which the picture was

taken. In all these figures, the north direction (toward Mars' north rotational

pole) is generally to the left, and east (astronautical convention) is toward the top

of the page.

It will be seen that, as the slant range over which the pictures were taken

decreased, the area covered by each picture grew smaller and the definition be-

came better. Toward the end of the picture-taking sequence, the light level and

contrast became very low. The illumination near the terminator (the border

between the daylight side and the night side) fell off more rapidly than expected,

and the camera could not adjust quickly enough to the loss of light.

270




Picture 1 (red).--This picture (see pp. 272 and 273), taken at a slant range of

about 17 000 km (10 600 statute miles), shows a bright area between Trivium

Charontis and Propontus II about 330 km (205 statute miles) along the limb of the

planet and about 1200 km (750 statute miles) from the limb to the bottom of thepicture. Phlegra, a bright area, is on the limb of the planet. The sun is 27 ° from

the zenith from the southeast. The main point of interest in this picture is the light

"smudge" area, which resembles a cloud above the horizon. This hazy patch

appears better in the raw version (a) than in the final version (c), in which the

contrast enhancement was designed to bring out the surface features. Initially it

was assumed that the area had been caused by an imperfection in the optical or

television system, but the smudge did not appear on calibration pictures taken

several weeks after encounter nor was duplication by the introduction of deliberate

defects in identical equipment possible. Thus, the experimenters have tentatively

concluded that what is seen is actually on Mars. Some evidence suggests that the

atmosphere of Mars may contain tiny crystals of frozen carbon dioxide at great

heights. The cloud shown in picture 1, if it is a cloud, extends to about 28 km

(18 statute miles) above the surface, while atmospheric haze appears to extend as

much as 150 kin (93 statute miles) above the surface.

271



#

MARINER-MARS 1964

272



. SCIENTIFIC RESULTS AND CONCLUSIONS

FIGURE-1 8.-Picture 1. (a) raw data ; (b) calibrated, reregistered data with errors and fiducial

marks removed; (c) same data as in (b), but with enhanced contrast.

273



MARINER-MARS 1964

FIGURE-13.-Picture 2. (a) raw da ta;

(b) calibrated, reregistered data

with errors and fiducial marks re-

moved; (c) same data as in (b),t- but with enhanced contrast.



I

8SClENTlFlC RESULTS AND CONCLUSIONS

Picture 2 (green).-The area seen in picture 2 is midway between the bright

areas of Elysium and Amazonis and is about 500km (310 statute miles) by 900 km(560 statute miles). The Sun is 20" from the zenith from the southeast (upper

right) and the slant range is 16 200 km (10 000 statute miles). This picture over-

laps the first by about 25 percent, as was planned for each pair of pictures. The

fact that the area which overlapped clearly showed some features on both pictures

proved that the camera system was returning valid data.

275



,

MARINER-MARS 1964 .

r 1

FIGURE-20.-Picture 3 . (a) raw data;



moved; (c ) same data as in (b),

but with enhanced contrast.

t- i



*SCIENTIFIC RESULTS AND CONCLUSlONS

Picture 3 (green).-Taken at a slant range of 15 300 km (9500 statute miles),

picture 3 shows a bright area 325 km (200 statute miles) by 514 km (320 statute

miles). This area is located southeast of Trivium Charontis on the western edge of

the Amazonis desert area. In this picture, the Sun is 13" from the zenith. Thc

3-km-across (2-statute-mile) resolution is a considerable improvement over that

of picture 1, where the resolution was 5 km (3 statute miles) across. A prominent

spot about 19 km (1 2 statute miles) across appearing in picture 3 is on the surface

of the planet.

277



MARINER-MARS 1964*

r

t-

7 -

I-

7

FIGURE-21.-Picture 4. (a) raw dat a;



moved; (c) same data as in (b),


e-

l

278



b SCIENTIFIC RESULTS AND CONCLUSIONS

Picture 4 red).-In this picture, taken a t a slant range of 14 900 km (9300

statute miles), the Sun is 14" from the zenith from the northeast (upper left). Abright area in Mesogaea measuring 304 km (190 statute miles) by 450 km (28C

statute miles) is shown. The few depressions which appear in picture 4 are called

craters simply because the later pictures showed them so clearly. Without the later

pictures, an analyst of this picture might conclude that it is a desert landscape

with a few roundish spots that might be dried-up lakes. The primary reason this

picture did not reveal very much is that it was almost noon on Mars when Mariner

IV passed this location, and, with the Sun overhead, there were no shadows.279



MARINER-MARS 1964b

r

FIGURE-22.-Picture 5. (a) raw data;



moved; (c) same da ta as in (b) ,


280



.a SCIENTIFIC RESULTS AND CONCLUSIONS

Picture 5 (red).-Picture 5 was taken at a slant range of 14 300 km (8900

statute miles), with the Sun 19" from the zenith from the north. The area is 281

km (175 statute miles) by 375 km (233 statute miles) and is located in a bright

area in eastern Zephyria. The surface detail became somewhat clearer in this

picture, but no positive conclusions regarding the roundish spots could yet be

made.

281



MARINER-MARS 1964

..

T





bu t with enhanced contrast.

IL J, J J

202



,SCIENTIFIC RESULTS AND CONCLUSIONS

Picture 6 (green) -Located in a bright area in eastern Zephyria as is picture

5, this picture covers a region that is 273 km (170 statute miles) by 349 km (217

statute miles). The Sun is 22" from the zenith from the north, and the slant range

is 14 100 km (8800 statute miles). As was true with picture 5, the surface detail

became slightly clearer in this picture.

283

A



,MARINER-MARS 1964






284



*

b

SCIENTIFIC RESULTS AN D CONCLUSlONS

Picture 7 (green).-Picture 7, taken at a slant range of 13 600 km (8400statute miles), shows an area 262 km (163 statute miles) by 310 km (193 statute

miles) in the bright region in southeastern Zephyria near Mare Sirenum. The

Sun is 29 O from the zenith from the north. This picture had a dramatic effect when

first received since it was the first to show that the roundish spots of the earlier

pictures were definitely craters. Ten or more of these craters can be seen.

285



MARINER-MARS 1964c

FIGURE-25.-Picture 8. (a) raw data ;





286



SClENTlFlC RESULTS AND CONCLUSIONS

Picture 8 (red).-This picture shows the border between Zephyria and Mare

Sirenum. The area covered is 255 km (158 statute miles) across by 296 km (184statute miles). The slant range is 13 400 km (8300 statute miles), and the Sun is

32" from the zenith. The center of picture 8 shows two craters measuring about 32

km (20 statute miles) in diameter side by side and a few smaller craters elsewhere.

The larger craters seem to have a broken appearance and generally flat-appearing

bottoms (when compared with some of the more recent lunar craters).

287



MARINER-MARS 1964

4

.

FIGURE-26.--.Picture 9. (a) raw dat a;





288




Picture 9 (red) .-Showing at least 20 craters of assorted sizes, this picture was

the first that strongly resembles an area of the Moon. I t was taken a t a slant rangeof 13 000 km (8100 statute miles) and reveals an area in Mare Sirenum, bordering

on Atlantis in the southwest (lower right) corner of the frame and measuring about

253 km (157 statute miles) by 225 km (140 statute miles). Th e Sun is 39O from the

zenith. One of the craters which may be seen in this picture has a central peak in it.

(A plot resulting from an analysis of this crater is given later in this discussion.)

There is also a slight indication of a straight edge in one part which is very dim.

289



MARINER-MARS 1964




moved; (c) same dat a as in (b),


290



L


Picture 10 (green).-The area shown in picture 10 is in Atlantis, borderingon Mare Sirenum in the northeast (upper left) corner of the frame. The dimensions

are 251 km (156 statute miles) by 26 7 km (166 statute miles). The slant range is

12 800 km (8000 statute miles), and the Sun is 42" from the zenith. .The edges of

two giant craters are seen jutting into the picture frame; in addition, about 12

rather small craters may be seen.

291



MARINER-MARS 1964


(b) calibrated, reregistered da ta


moved; (c) same da ta as in ( b) ,


292



.SCIENTIFIC RESULTS AND CONCLUSIONS


picture 11 shows an area in Atlantis between Mare Sirenum and Mare Cim-merium about 250 km (155 statute miles) across and 254 km (158 statute miles)

deep. The Sun is 47” from the zenith. This picture shows the largest crater with a

diameter of about 120 km (75 statute miles) and one of the smallest craters with a

diameter of about 6 km (3.8 statute miles). A straight edge may be seen that is

dark in some places and light in others. There are also some domed regions which

look like “inside-out” craters. The airbrush rendition in figure 7-16 was made from

the raw-data version of this picture.293



MARINER-MARS 1964

FIGURE-29. -Picture 1 2 . (a) raw data;

(b) calibrated, reregistered d ata






Picture 12 (red).-This picture, covering an area of 256 km (159 statutemiles) by 254 km (158 statute miles) in Mare Cimmerium, bordering on Atlantis

in the northeast corner (upper left) of the frame, is relatively featureless. Only a

few craters are evident. T he slant range for picture 12 is 12 400 km (7700 statute

miles), and the Sun is 52 O from the zenith.

295



MARINER-MARS 1964

FIGURE-30.-Picture 13. (a) raw da ta;





296



(c)

Picture 13 (red).-This picture was the first in which white areas began to

show on crater rims, as if covered with snow at their highest elevations. By thetime picture 13 was taken, Mariner IV was over fairly high southern latitudes

where it was midwinter. Another straivpht edge can be seen. There is also a sug-

gestion of a circular feature which would be interpreted as a crater except that it

seems to be rather light colored all around, rather than shadowed on one edge.

The area covered is 254 km (158 statute miles) by 242 km (150 statute miles) on

the border between Mare Cimmerium to the north and the bright region, Phae-

thontis. The slant range is 12 200 km (7600 statute miles), and the Sun is 58O from

the zenith from the north (left). 297



MARINER-MARS 1964

FIGIRE 7-31.-Picture 14 . (a) raw data;


with errors and fiducial rnarks re-

moved; (c) same da ta as in (b ),


298



.SClENTlFlC RESULTS AND CONCLUSIONS


picture 1 4 covers 257 km (160 statute miles) horizontally and 240 km (149 statute

miles) vertically in a bright area in northwestern Phaethontis. The Sun is 61O from

the zenith. Very marked light patches are visible, and the apparent craters are

ringed around their circumference like the crater seen in picture 13. The white

areas which ar e evident are thought to be frost on the surface of the planet since

this picture was taken in the subpolar regions, where it is generally believed that

frost exists.

299



MARINER-MARS 7964

4

JRE 7- -32. -Picture 15. raw data 7








Picture 15 (green) .-Beginning with this picture, the light level dropped

faster than the automatic gain control could adjust. Also, the atmosphere of theplanet could have obscured some features. N o very clear craters are evident, but

some light patches can be seen. However, in terms of the actual levels of intensity

and the angle of the Sun (67" from the zenith), this frame should have been the

most sensitive for detecting shading on the surface of the planet. The area shown is

266 km (165 statute miles) by 236 km (147 statute miles) in a bright area in

Phaethontis. The slant range is 1 2 000 km (7500 statute miles).

30



MARINER-MARS 1964






302



*SClENTlFlC RESULTS AND CONCLUSIONS

Picture 16 (red).-The Sun is 70" from the zenith from the north, and the

slant range is 12 000 km (7500 statute miles) for this picture. Covering 273 km(170 statute miles) east-west and 236 km (147 statute miles) north-south, the

picture shows a bright area in Phaethontis near Aonius Sinus. Except for some

white spots in a few places, virtually no features are evident. Version (c) shows

contouring," giving it a grainy, layered appearance, caused by the extreme con-

trast enhancement of a digital picture that was originally of very low contrast.

< <

303



*

MARINER-MARS 1964









Picture 1 7 (red).-A region 292 km (181 statute miles) by 238 km (148statute miles) in a dark area in Aonius Sinus is covered in picture 17. Its slant

range of 12 000 km (7500 statute miles), a little less than twice the diameter of

Mars, was the lowest during the picture-taking sequence. Th e Sun is 76" from the

zenith. The raw picture was very low contrast, so that even more contouring ap-

pears in the enhanced version than in that of picture 16.

305



MARINER-M ARS 1964




moved; (c) same data as i n (b),




.'r SClENTlNC RESULTS AND CONCLUSIONS

Picture 18 (green).-For this picture, also taken at a slant range of 12 000 km(7500 statute miles), the Sun is 80" from the zenith from the northwest. The area

shown is 308 km (191 statute miles) by 24 2 km (150 statute miles) in the same dark

area as picture 17.

307



.MARINER-MARS 1964

FIGURE-36.-Picture 1 9 . (a) raw data ;


with errors and fiducial marks rc-


brit with enhanced contrast.




Pictures 1 9 to 22.-Picture 19 was taken at a slant range of 12 100 km (7500

statute miles) and covers an area 360 km (224 statute miles) by 258 km (160statute miles) in the same dark area as pictures 17 and 18. The Sun is 88" from the

zenith from the northwest. The camera crossed the terminator to the dark side of

Mars in this picture; evidence for this can be seen in the eastern corner of the

frame. Picture 20 is almost entirely beyond the terminator, picture 21 is entirely

beyond it, and the 22 lines of picture 22 are entirely beyond the terminator and

may be partly beyond the dark limb of the planet.

309



MARINER-MARS 1964

FIGURE-37.-Picture 20 (raw data).

FIGURE-38.-Picture 21 (raw data) .

310



.SClENTlFlC RESULTS AND CONCLUSIONS

FIGURE -39.-Picture 22 (raw data).

311



MARINER-MARS1964

Conclusionsdrawn from the Mariner IV picturesof Mars.

With all systems having functioned perfectly, the total area photographed by

Mariner IV was about 1 554 000 sq km (600 000 sq statute miles), approximately

1 percent of the entire surface of Mars. The major surprise in the pictures was the

large number of craters: more than 70 of varying sizes are clearly distinguishable.

If what is seen on the Mariner IV photographs is typical of what would be seen

elsewhere on the planet, there must be more than 10 000 craters on the surface of

Mars. These craters were not expected by most scientists.

On the basis of the sample provided by Mariner IV, it may be said that the

number of large craters per unit area on the Martian surface and their size distri-

1000

E I00

eY

w

u

kuIi0

z

_HIGHLANDS)

MOON (MARIA) _ _

.%I IO IOO I000

CRATER DIAMETER, km

FICURE 7-40.--Number of craters on the Moon and on Mars (estimated) as a

function of crater diameter.

312




bution resemble closely the number and size distribution of craters on the Moon.

Figure 7-40 compares the number of craters estimated to exist on the surface of

Mars (as a function of crater diameter) with the number on the Moon. There are

fewer Martian craters with diameters below 10 km (6.25 statute miles) than would

be expected on the basis of our knowledge concerning the Moon.

The craters on Mars have rims that rise about 100 meters above the sur-

rounding surface and depths that extend several hundred meters below the

rims. The crater walls slope at angles up to about 10 ° . Figure 7-41 compares the

depths and diameters of Martian craters with those of the lunar craters. The

depths of the Martian craters were estimated by plotting the light values of succes-

sive picture elements on a line cutting across the diameter of a crater. Such a plot

is shown in figure 7-42. The light-intensity curve at the bottom of figure 7-42traces the light and dark values on the 64-step intensity scale assigned to each

picture element in one line across the middle of the picture.

There seems to be a tendency for the small craters to appear on the rims of

large craters. This trend suggests that there may be something special about the

E

k-

nWQ

1500• PICTURE 7

o PICTURE 8

& PICTURE 9

17 PICTURE I0

X PICTURE II

tooo

50o

/• !

1

NOT :WHENAME'RA R.PPEAREDNTWO,CTURESANO iTS SIZE WAS DIFFERENT

IN EACH, THE TWO VALUES

ARE CONNECTED

In

0 tO 20 30 40 50

DIAMETER, km

FIGURE 7-41.--Crater depth as a function of crater diameter for the Moon and for Mars.

313



MARINER-MARS 1964

42

43

44

45

46

47

40

\ /

FIGURE-42.-Plot used to compute the slope and depth of a crater. The plot refers to the

data numbers along a line running left to right through the middle of the crater shown.

(Crater shown is from picture 9.)

314




composition or texture of the crater rims that resists whatever forces tend to erode

small craters when they are formed elsewhere on the Martian surface.

Many of the craters are flattened along a portion of their circumference

instead of being circular. This phenomenon, also observed in lunar craters, is

believed to result from structural faults below the surface. The straight edge in

picture 11 intersects a crater and continues across the rim. This, too, might have

been caused by a fault.

The principal topographic features of Mars in the areas photographed by

Mariner IV have not been produced by stress and deformation originating within

the planet, in contrast to the features of Earth. Earth, of course, is internally

dynamic, giving rise to mountains, continents, and other such features; Mars, on

the other hand, has evidently long been inactive. The lack of internal activity isalso consistent with the absence of a significant magnetic field at Mars (as deter-

mined by the Mariner IV fields and particles experiments).

Although it may be difficult ever to arrive at an unambiguous identification

and interpretation of all the features seen on the photographs, it is felt that the

existence of a lunar-type cratered surface, even in only a 1-percent sample, has

profound implications about the origin and evolution of Mars and further en-

hances the uniqueness of the Earth within the solar system. By analogy with the

Moon, much of the heavily cratered surface of Mars must be very ancient: perhaps

2 to 5 billion years old. However, a definite statement concerning the age of the

Martian surface cannot be made until more of the surface has been photographed

and until more is known about the relative rates of impact of asteroid-sized bodies

on the Moon and Mars.

The remarkable state of preservation of this surface infers that no atmosphere

significantly denser than the present very thin one has characterized the planet

since its surface was formed. (The atmosphere of Mars is discussed later in this

section.) Similarly, it is difficult to believe that free water in quantities sufficient to

form streams or to fill oceans could have existed anywhere on Mars since that time.

The presence of such amounts of water (and consequently atmosphere) would

have caused severe erosion over the entire surface, as was true with the Earth.

Surface features on Earth are eroded and effaced in a few tens of millions of years.

Canals were looked for on the Mariner IV photographs, but nothing can be

seen that is obviously a canal. Although the trace of the camera view crossed

several of the canallike markings sketched from time to time on maps of Mars, no

such features could be identified with certainty. The apparent lack of these

features in the Mariner IV photographs could be due to several factors, including

315



t,

MARINER-MARS 1964

the following: (1) The season was unfavorable, because these canallike markings

do not show well during winter; or (2) the pictures may show whatever it is that

makes a canal in so much detail that it cannot be recognized. Therefore, nothing

positive concerning the existence or lack of canals can be concluded on the basis

of the Mariner IV photographs.

As anticipated, the Mariner IV photographs neither demonstrate nor pre-

clude the possible existence of life on Mars. Geologic experience gained on the

Earth suggests that the search for a fossil record on Mars appears less promising if

oceans never existed on the planet. On the other hand, if the surface of Mars is

truly as ancient as is now supposed, that surface may prove to be the best (and

perhaps the only) place in the solar system still preserving clues to primitive

organic development, traces of which have long since disappeared from the Earth.

Occulation

Obiectives oFexperiment

In the spring of 1964, after the Mariner-Mars 1964 spacecraft were built and

the mission objectives had been defined, an occultation experiment was included.

This experiment required no changes in the spacecraft; all that was needed was a

shift in the aiming point at Mars to insure that the spacecraft would pass behind

the planet Mars as viewed from the Earth. This stipulation meant that the 2300-

MHz radio signal from the spacecraft would pass through the atmosphere of Mars

as the spacecraft went behind Mars (as viewed from the Earth) and as it emerged

on the other side. The Earth occultation region, where the Earth would actually

be hidden from view of the spacecraft and the radio signal would disappear, is

illustrated in figure 7-43. As the radio signal passed through the Martian atmos-

phere, it would be refracted (i.e., deflected from a straight path when passing

obliquely from one medium to another in which its velocity is different), giving

rise to a change in the apparent motion of the spacecraft.

If all other factors producing apparent motion of the spacecraft were ac-

counted for (e.g., the actual motion of the spacecraft, the motion of the deep space

stations on the rotating Earth, the lengthening of the transit time of the signal, and

the refractivity of the Earth's lower atmosphere), the remaining unexplained

changes in the radio signal could be attributed to refraction by the atmosphere of

Mars. (For a successful experiment, it was necessary to account for the total

change in frequency or phase of the signal due to all causes other than refraction

316



b

SClENT/F/C RESULTS AND CONCLUSlONS

FIGURE-43.-Geometry of Mars and Mariner IV at the time of Earth occultation.

by the Martian atmosphere to an accuracy of at least one part in loll.) Since the

geometry obtained from the estimated trajectory is known, the measured changes

could be used to estimate the spatial characteristics of the index of refraction (or

refractivity) in the electrically neutral atmosphere and electrically charged ion-

osphere of Mars. Thus, by measuring and then analyzing the changes in the

characteristics (frequency, phase, and amplitude) of the radio signals from the

spacecraft, it was hoped to learn more about the composition, density, and scaleheight of the Martian atmosphere. Knowledge concerning these atmospheric

characteristics is essential to any attempt at defining the entry and landing re-

quirements for future spacecraft missions to Mars.

3The erm used t o designate the height of atmosphere needed to produce a given surface pressure if the density

of the atmosphere were constant from top to bottom. For the Earth’s atmosphere, the scale height is 7 km (4.4

statute miles).

317



MARINER-MARS 1964

Earth occultation during the Mariner IV mission

At 02:19:11 GMT on July 15, 1965 (approximately 1 _ hours after its closest

approach to Mars), the Mariner IV spacecraft entered the Earth occultation

region at Mars and the radio signal from the spacecraft ceased. Less than 1 hour

later (at 03:13:04), the spacecraft left the region and the radio signal was re-

acquired. During this phase of the mission, Pioneer DSS 11 and Echo DSS 12 took

standard Doppler (closed-loop) data 4 as well as open-loop records _ of the received

data, and Woomera DSS 41 and Tidbinbilla DSS 42 took Doppler data only.

When the signal first passed through the Martian atmosphere, it was after-

noon according to Mars local time, and the Sun was about 20 o above the horizon.

At its entrance into the Earth occultation region (as projected on Mars, be-

tween Electrus and Mare Chronium at 55 ° south latitude and 177 ° east longi-

tude), Mariner IV was 22 559 km (14 021 statute miles) from the limb of theplanet, traveling at a velocity of 2.09 km/sec (1.3 statute miles/sec) perpendicular

to the Earth-Mars line. All data were taken while the spacecraft's transmitter

frequency reference was provided by a frequency standard on Earth.

When Mariner IV emerged from behind the planet and the radio signal

again passed through the Martian atmosphere, the spacecraft flew tangent to a

point where, according to Mars local time, it was close to midnight. At its exit

from the Earth occultation region (as projected on Mars, above Mare Acidalum at

60 ° north latitude and 44 ° west longitude), the spacecraft was 36 119 km (22 448

statute miles) from the limb of the planet. A portion of the data taken at that time

was received while the spacecraft's transmitter frequency reference was provided

by a crystal oscillator on the spacecraft. In that mode, phase measurements were

significantly less precise.

Results and conclusions

The analysis of Doppler tracking data taken before and after the encounter of

the spacecraft with Mars yielded the Mariner IV trajectory at the time of occul-

tation with such precision that the range rate of the spacecraft was known at an

accuracy of 0.0015 m/sec (0.005 ft/sec). Thus, any significant deviation of the

received Doppler data from expected values based on trajectory analysis was

expected to have been caused by atmospheric and ionospheric phase-path effects.

4Taken from the ordinary tracking channel, which gave a cycle count of the spacecraft's frequency.

s Contained in so-called open-loop receivers, which were modified to record a frequency-translated version of

the actual received signal on tape in the audiofrequency range (about 2 to 3 kc). This system enabled much more

precise measurements of the received power of the signal as well as the frequency and phase relationships.

318




The changes in the phase of the Mariner IV radio signal, as recorded at

Pioneer DSS 11 prior to the entrance of the spacecraft into the Earth occultationregion, are plotted in figure 7-43. The times shown in figure 7-44 are those at

which the data were taken on Earth and are approximately 12 rain later than the

times at the spacecraft because of the period required for the signal to travel from

the spacecraft to the Earth. As can be seen, the signal phase changed about 10

cycles from the expected value (if no atmosphere or ionosphere were present) as

the signal passed through the electrically charged ionosphere. As the signal

reached the electrically neutral lower atmosphere, its phase began to return to the

expected value and then moved in the opposite direction until a value over 25

cycles from the expected value was reached. At approximately 02:31:11 GMT, the

radio signal from the spacecraft was no longer received on Earth, an indication

that the spacecraft had entered the Earth occultation region (at 02:19:11 at the

spacecraft). All data received had been taken in the brief period of about 100 sec.

The receipt of these data marked the first time coherent radio transmission was

used to probe the atmosphere and ionosphere of another planet.

Both the ionosphere and atmosphere of Mars were somewhat less dense than

expected. The data on the ionosphere obtained prior to Mariner IV's entrance

into the Earth occultation region show a distinct ionized layer with a peak electron

density of about 9.0±1.0×104 electrons/cm 3 (14.8+1.6×105 electrons/in. 3) at

an altitude of 121 to 126 km (75 to 78 statute miles). (The peak electron density of

the Earth's ionosphere is about 106 electrons/cm _ (16.4×106 electrons/in.3).)

Indications of a second less-dense ionized layer were noted. (No ionosphere was

detected as the spacecraft left the Earth occultation region at approximately mid-

night on Mars; however, the electron density in the atmosphere at night is at least

20 times lower than that during the day.) The electron scale height of the ion-

osphere above the electron-density peak is about 20 to 25 km (12.5 to 15.6 statute

miles). _Ihe low altitude of the peak electron density and the small scale height of

the ionosphere above it indicate that the temperature is considerably lower than

had been anticipated. The temperature at 121 to 201 km (75 to 125 statute miles)

is estimated to be less than 200 ° K (-99.4 ° F).Various theoretical models of atmospheres were proposed so that the data

obtained on the Mars atmosphere could be compared with those for the models

and hopefully a "fit" could be made. Such a fit is illustrated in figure 7-45, where

the solid curve represents the computed phase change for a theoretical model

atmosphere having a surface refractivity of 3.7 N-units and a scale height of 9 km

(5.6 statute miles). The actual phase change, as computed from data collected by

319



MARINER-MARS 1964

3O

8_, 20

,o

...1_ 0

n

-20

IONOSPHERE---_

ATMOSPHERE--\

02;28:30 02:29:I0 02:29:50 02:30:30 02:31:10

O2:28:50 02:29:30 02:30:10 02:30:50 O2:31:30

GMT AT EARTH, JULY 15_1965

FIGURE 7-44.--Signal phase changes prior to entrance of Mariner IV into Earth

occultation region at Mars.

(.9Z

IL.)

"I-Q.

Z

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5O

4O

I I I 1 [

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GMT AT EARTH, JULY 15, 1965

FIGURE 7-45.--Mariner IV data for electrically neutral atmosphere fit to theoretical

atmospheric model curve.

320




Pioneer DSS 11, Echo DSS 12, and Tidbinbilla DSS 42, is seen to fit quite well

with that model. (For the model, it was assumed that the atmospheric densitydecreases exponentially with altitude.) The best fit obtained yielded a value of

3.64-0.2 N-units for the surface refractivity and a value of 8 to 10 km (5 to 6.3

statute miles) for the scale height. (The refractivity of the electrically neutral part

of the Earth's atmosphere is 350 N-units.) The data indicate that there is no

obvious change of scale height with altitude up to at least 30 km (18.8 statute

miles). This finding, together with Earth-based observations of carbon dioxide in

the atmosphere of Mars, clearly indicates that the atmosphere consists primarily

of carbon dioxide and that the amount of nitrogen present is very small.

On the basis of the assumption of a Martian atmosphere consisting primarily

of carbon dioxide and on the basis of determined values of refractivity and scale

height, the mass density and surface pressure for certain models of the atmosphere

have been estimated. For a pure carbon dioxide atmosphere, the surface mass

density (or density at that point at which the refractivity is that given here) is

1.434-0.1×10 -5 g/cm a (5.184-0.36×10 -7 lb/in.a). The surface mass density

for an atmosphere of 80 percent carbon dioxide and 20 percent heavier gases

(argon and/or nitrogen) is 1.54-0.15×10 -5 g/cm a (5.43±0.54×10 -7 lb/in.a),

and that for an atmosphere composed of equal parts of carbon dioxide and argon is

1.754-0.10×10 -5 g/cm a (6.334-0.36×10 -7 lb/in.a). If these values are correlated

with the scale height, the temperature ranges for the three types of atmos-pheres can be established as follows: 1804-20 ° K (-135.4+36 ° F), 1754-25 ° K

(-144.44-45 ° F), and 1704-20 ° K (-153.44-36 ° F), respectively.

If the previously mentioned temperature ranges and the measured refractivity

range are assumed, the surface pressure can be estimated as 4.1 to 5.7 millibars for

a pure carbon dioxide atmosphere, 4.1 to 6.2 millibars for an atmosphere com-

posed of 80-percent carbon dioxide and 20-percent heavier gas, and 5.0 to 7.0

millibars for an atmosphere in which carbon dioxide and argon are present in

equal amounts. This lower-than-expected surface pressure leads us to believe that

it will be much more difficult to design capsules capable of landing on the surface

of Mars than was previously supposed. However, the scale height is less and the

mass density expected at the atmospheric peak is quite low, indicating that

density falls off quite rapidly in the upper atmosphere. Therefore, it should be

possible to orbit the planet at lower altitudes than had been thought feasible prior

to the Mariner IV mission.

In the data presented, there is a possibility of an error of about 10 percent

caused by the inability to determine exactly what feature on the Martian surface

321



MARINER-MARS1964

actually cut off the radio signal as the spacecraft passed behind the planet. If the

feature happened to be some large crater or a protuberance of another kind, the

values presented here are low. It can be estimated that the surface pressure would

change by about 1 percent for every 90 meters (300 ft) of elevation difference on

the surface. It is difficult to say whether or not Mars was observed all the way

down to its surface. The surface temperature is undoubtedly about 30 to 40

Kelvin degrees (54 to 72 Fahrenheit degrees) higher than that observed, and there

is bound to be some fluctuation of temperature with altitude at some point.

The near or complete absence of a static magnetic field at Mars has very

interesting implications with regard to understanding the ionosphere and atmos-

phere of both Mars and Earth. For Mars, it means that formation and loss

mechanisms can be better understood and related to the physical characteristics ofthe atmosphere and ionosphere, since there are no complicating effects of a mag-

netic field. For Earth, many ionospheric phenomena are still not well understood,

often because of the complicating effects of the magnetic field in controlling in-

coming charged particles, in affecting ionospheric motions, in storing high-energy

particles which may provide a heating and ionization source, in affecting and

controlling small- and large-scale ionospheric irregularities, and in providing

partial shielding from the solar wind. Results of studies of the Martian ionosphere

should therefore aid in separating and understanding various phenomena in that

region surrounding Earth.

3_





MARINER-MARS 1964

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APPENDIX A

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MARINER-MARS1964

Atlas D/Agena D launch vehicle. The development time schedule permitted, and

mission reliability demands made necessary, full use of both subsystem and system

validation testing concepts. Three complete, fully flight-qualified spacecraft wereto be provided, as well as spares of certain critical assemblies. In order to quafify

this hardware, the equivalent of three sets of checkout, test, and handling equip-

ment was to be furnished to support testing and launch operations. Spacecraft and

launch vehicles would be processed in parallel so that the second launch could

occur as soon as 2 days after the first launch.

JET PROPULSION LABORATORY ORGANIZATION AND

MANAGEMENT

Under contract to NASA and managed by the California Institute of Tech-

nology, JPL assumed project management responsibility and system management

responsibility for: (1) the design, fabrication, and testing of the spacecraft and the

required ground support equipment; (2) the space flight operations for the

missions from injection to the end of the missions; and (3) the tracking and data

acquisition activities. The system management organizations are shown in figures

A-3 to A-5.

Figure A-1 showed the system management function and responsibility

assignments defined by JPL in the Mariner-Mars 1964 project development plan.

The master schedule milestones and due dates were also included in that docu-

ment. At the beginning of the project, 131 milestones were selected and their

phasing determined across the 2-year period up to launch. The schedule for major

events is shown in figure A-6. The progress of events was reported regularly to

NASA Headquarters.

A project policy and requirements document was also prepared to establish

the operational procedures for the project within JPL. That document comple-

mented the project development plan by expanding the system management

concepts and presenting a compatible project master schedule.

In order to coordinate the efforts of the 1100 persons (peak) working on the

project at JPL, project representatives were chosen from each JPL technical

division to serve as central controllers for all project activities within the technicaldivision. Since each technical division was involved in activities other than just

those for the Mariner-Mars 1964 project, and since the project representatives had

control over only those resources committed to the project by the technical division

manager, monthly meetings of the division managers and Mariner-Mars 1964

project office personnel were held. The purposes were to inform the division

326



APPENDIX A

managers of the progress of the project, to make available reports of the contri-

butions of the technical divisions to the project, and to solicit the division man-

agers' ideas regarding the project. In addition, regular weekly meetings were heldwith the JPL system managers and the project representatives indicated in figure

A-7. The project representatives were required to submit progress reports at

regular intervals.

Besides the regularly scheduled reviews and meetings, numerous other

reviews were conducted as required or as deemed desirable; for example:

1. A detailed spacecraft subsystem review in spring 1963.

2. A series of preshipment acceptance reviews on each spacecraft in summer

1964.

3. A series of launch-readiness reviews at JPL and the AFETR (CapeKennedy, Fla.) to determine equipment flightworthiness.

4. Reviews concerning launch vehicle performance and preparation.

An ...... : "_,..,te._v,. quality assurance _.A ..,_;_hn;,,, v'"e, .....................

addition, a comprehensive documentation effort was implemented to provide each

person involved in this complex project all necessary information for executing his

assigned tasks.

LEWIS RESEARCH CENTER RESPONSIBILITY

The responsibility for the Atlas D/Agena D launch vehicle was assigned to

LeRC in January 1963. This assignment included administrative and technical

cognizance and control over the launch-vehicle system procurement, booster

launch and flight operations, and the delivery and analysis of flight performance

and tracking information up to the time of spacecraft injection. LeRC established,

as its principal agent, the Agena project office with the responsibility for insuring

proper vehicle support to several NASA projects using Agena and Atlas vehicles,

including the Mariner-Mars 1964 project. The Agena project organization is

illustrated in figure A-8.

Launch operations were conducted at the NASA John F. Kennedy Space

Center, Cocoa Beach, Fla. The actual conduct of launch operations was directed

by the Goddard Space Flight Center, under the technical cognizance of LeRC.

An additional responsibility of the Goddard Space Flight Center--the manage-

ment and operation of the Worldwide Communications Network during the space

flight testing and operations phases of the project--was under the supervision of

the NASA Tracking and Data Acquisition Office.

327



MARINER-MARS 1964 J

_ AC EC m_ T _ S_ _ AC -E R

D SCHNeq_

._ COGNIZANT _IEN_IST

OCCULW_ION

A _ LI _C

CO_ _IZAN, E_GIN_E_

CO_NIZANrENGq_ER

_NEVO_IE_

0 0. r,'C_RIS

COG_LZA_. t _C.I_W. PO_I_

0A_ *UTO_,ION _S_EM

¢O_NIZAN, E_GINEE_

¢OGNIZAN1 ENGI_k

PLAN_ _XN

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PROJECT_I_E_L. _"UTZ

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CO_NIZAN_ EN_INEE_0ATA E_ODER

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_ F D AT A S_S_E_

R. S,OF_

C_NIZAN_ E_GINEER __._ INSrRU_N_O_

S-_.N_ TR^NSeON_, S_'ERVlSO_

FIGURE A-3.--Spacecraft system management

328



APPENDIX A

ICOG NIZANT ENGI_ER 11 GU&DA_E AND CONTROl

o. R. rHO_S J' _" ACOR0

pROJEC T ENGINEER

CC_S

N. H. HAR_N

COGNIZANTU.,CC&SLINOONEGINEE_

I

"_t PROJECt ENGINEER

POWER SUSSYSIEM

K. M. DAWSON

COGNIZAN T EN01N[_R

pOWER SUbSYStEM CONTRACT

G, C. CLEVEN

COGNIZANT ENOIN[E__R PAN[kS

J. V. OOLDSM,TH

CO_ NITANI FNGINfflSAt_ERIES

W . = . _ ON O

C_NiZ ANT ENGINEER J

EL_Cr,lCAL co_E_re,s

R.L. SeENCER

1I I

J-t...............ROJECT ENGINEER GUIDANCE _D CON[ROLTTITU_ -CO NTROL _a_ STEM

W . W . BENJAMIN J.L. SAWNO

COGNIZANt ENGINEE. [ COGNIZANT ENGINEER COGNIZANt ENGINEER

ArTnuoE-CONIROL q POWE_ OSE raN-PULLERS AN0

_ACKAGING pIN._JLL£R _UI_S

,__ COGNiZaNT ENGINeeR

COGNIZANI ENGIN_ COGNIZANt _GINE_R SCI_NTIEICINStRUmeNtS

J_ VANE _ClUAIO_S ATTIIU_CONTROL OS_ pACKAG,NG

c . CA_ S,A_A J. _. P_TRAUA _ CA_FR

................ 1 [ I ! .................

................... I I .......

COGNIZANT ENGINIERC AN _U S _ N_

_. S, DAVIS

C_ NIZANI _NOI_I COG NIZAN_ _NGIN[[RSOLARVAN_ CONTROL TEL_CO_UNICAnONS

ELECTRONICS PACKAG,NG

r_J. 0ONLIN R. E_S"[_

.___ COGN,ZA_ ENG,NE_R

INERTIA_ S_NSO_S

p . J . H AN I>

COGNIZANT ENGINEER

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COGNIZANT _NGINEER

SUN SENSORS

k . F . _HM_T

_ COGNIZANT EN_IN_

AUmnLOT

organization for Mariner-Mars 1964 project.

I

IASSISTANT _OJECI

REPRE_ NTATIVE

J. D. SCHMUECKE_

ANTENNAS AND D_VISIGN

_II REPel_ NTAIIVE

W. E. LAYMAN

COGNIZANT _NGINEE l

_Y RONET_K S A NDOS_

COGN_ZANT_NGIN_R l I CO_N_Z_NTENOIN_R

SO_AR_ANE_ ___ GROUNO HANDU_

AC_UAIO_ E_U,_NI

co_.,_N,,,.....

_AN _TUA_OR _NGI_EE_

E. L. FLOYD C.W. _AGGIO, JR.

IE_ERAT_E CONTROL SOLAPPR_S_R_ VANEr_S_NG A N_ _ AS

_. B. GRAM J. C, RANDALL

_ RI_RY SIR_T _E SU_ERST_:TU_E ANOrH_R_L SHIELDS

329



MARINER-MARS 1964


SYSTEM MANAGER

T. S. BILBO

ISPACE FLIGHT OPERATIONS

DIRECTOR

D. W. DOUGLAS

I A ER /FLIGHT PATH ANALYSIS J

AND COMMAND

N. R. H AYNES t+F+II++'F+NALYSIS AND COMMAND AND COMMAND

A. G. CONRAD R. SLOAN

ISCIENCE OPERATORS J

(includes principal investigators

when acting in this capacity)

FIGURE A-4.--SFO system management organization for Mariner-Mars 1964 project.

I I I

I_'°...II_+1l ....1SS II DSS 12 DSS 13

S TA TI ON M AN AG ER STATTON MANAGER STATION MANAGER

J. IIUCKLEY H, OLiN J. BUCKLEY

I D SI F S YS TE M M AN AG ER. A. RENZEITI

DSIF OPERATIONS MANAGER

R. K. MALLIS

C . A . H OLR ITZ s A SS IS TA NT

I°++°1......ND NET CONTROL

A. T. BURKE

I PROJECT REPRESENT ATI VE IJ. R, HALL

I I i

I++II .........+°...ISS 41 DSS 42 DSS 51

STATION MANAGER STATION MANAGER STATION MANAGER

W. METI_EAR R. LESLIE D , H OG G

1

I++1S S 6 1

S TA TI ON M AN AG ER

P. TARDANi

FIGURE A-5.--DSIF system management organization for Mariner-Mars 1964 project.

330



APPENDIX A

EVENT

AUTHORIZE

STUDY

DESIGN

FABRICATE

ASSEMBLE

AND TEST

LAUNCH

FIGURE A-6.--Mariner-Mars 1964 project schedule.

The Air Force Space Systems Division (and its designees) acted as an agent

for LeRC in Atlas D and Agena D procurement, logistics, and management

support. The addition of special equipment for the Mariner-Mars ;1964 missions in

the Agena D and the adaptation of conventional Agena D military hardware were

performed by the Medium Space Vehicle Programs Office of the Lockheed

Missiles and Space Company (LMSC) at Sunnyvale, Calif., under the cognizance

of that company's Space Programs Division. This organization was under direct

contract to LeRC to execute all launch-vehicle responsibilities, except for the

procurement of the basic Agena D vehicle. The Medium Space Vehicle Programs

Office cooperated with General Dynamics/Convair (now General Dynamics/

Astronautics) at San Diego, Calif., producer of the Atlas D; and with Space Tech-

nology Laboratories (now TRW Systems) at Redondo Beach, Calif., the launch-

331



MARINER-MARS 1964

I

I

I

i

V EXPERIMENTERS

PROJECT SCIENTIST

R. SLOAN

NASA HEADQUARTERS

OFFICE OFSPACE

SCIENCE ANDAPPLICATIONS

G.REIFF

PROJECT

MANAGEMENT

J.N.JAMES

W.A.COLLIERT.H.PARKER

MAIN DEPENDENCY

-- -- -- ALTERNATE RELATIONS

X NOT APPLICABLE

;YSTEM

MANAGERS_

PRC IECTRI RESENTATIVES

PROJECT

ENGINEERING

J. CASANI

M. GOLDFINEJ. MACLAY

O. SHAWA. WILL IAMS

QUALITY

ASSURANCE

AND

RELIABILITY

R.WELNtCKF.WRIGHT

SYSTEMS

N. HAYNES

D. DOUGLAS

SPACE

SCIENCES

H. TROSTLE

TELECOMMUNI-

CATIONS

J. BRYDEN

GUIDANCE

AND

CONTROL

T. ACORD

ENGINEERING

MECHANICS

J, WILSON

PROPULSION

B. SCHMITE

ENGINEERING

FACILITIES

D. HESS

PROCUREMENT

L,WRIGHT

FI NANCE

MANAGEMENT

(FISCAL REPORTS

ONLY)

SPACE FLIGH1

OPERATIONS

T, BILBO

X

X

DSIF

N RENZETT I

X

X

X

X

X

X

LAUNCH

VEHICLE

s. c, HIMMEL

X

X

X

SPACECRAFT

D SCHNEIDERMAN

X X

X

X

X

X

FIOVRE A-7.--Mariner-Mars 1964 project matrix organization.

332



APPENDIX A

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333



MARINER-MARS 1964

vehicle trajectory and flight-performance analysis contractor. The Medium Space

Vehicle Programs Office also designed, qualified, and delivered the shroud system.

ADVISORY PANEL PARTICIPATION

Advisory panels formed during the Mariner-Venus 1962 project and the

Ranger project were also consulted by Mariner-Mars 1964 project office per-

sonnel and the spacecraft system manager. The panels, however, exercised no

technical direction over JPL, LeRC, or their contractors. The panels concerned

with performance control, trajectories, guidance and control, and flight dynamics

and with tracking, communications, inflight measurements, and telemetry con-

tinually monitored, compiled, evaluated, and coordinated data relating to their

respective areas as those areas interacted with the launch vehicle, shroud, and

spacecraft. The Launch Operations Working Group, acting as the prime coordi-

nator of flight preparations at the AFETR, participated in launch vehicle, space-

craft, support facilities, and range-readiness meetings. The Launch Vehicle

Integration Group, representing working levels of JPL, LeRC, and LMSC, was

concerned with resolving incompatibilities between the launch vehicle, shroud,

and spacecraft.

334



APPENDIX B

Abbreviations

AAC

ac

AFETR

AGE

APW

AU

BECO

CALIB

COS RAY

CPM

DAS

dc

DC

DN

DSIFDSN

DSS

ENC

EPM

EPW

FPAC

GCS

GM

GMT

GTS

IMP

JPL

LCE

LeRC

automatic aperture control

alternating current

Air Force Eastern Test Range

automatic ground equipment

analog-to-pulsewidth

astronomical unitbooster-engine cutoff

calibration, 1

central computer and sequencer

cosmic ray telescope

Canopus-probe-Mars (angle)

data automation system

direct current

direct command

data number

Deep Space Instrumentation Facility

Deep Space Network

deep space station

encounter

Earth-probe-Mars (angle)

Encounter Planning Working (Group)

Flight Path Analysis and Command (Team)

Ground Communication System

Geiger-Mueller

Greenwich Mean Time

ground telemetry subsystem

impact probability


launch complex equipment

Lewis Research Center

335



MARINER-MARS1964 ,J

LLA

LMSC

MAG

maxMDL

MT

N

NAMG

NASA

NASCOM

NRT

NRZ

OSE

OSSA

PAS

PCM

PDP

PL

PN

POS

QcRF

RT

RWV

SAA

SAF

SCM

SECO

SFO

SFOF

SIT

SMIT

SNSPAC

SPM

SSAC

TEMP

low-level amplifier

Lockheed Missiles & Space Co.

magnetometer

maximummaster data library

master timer

newtons

narrow-angle Mars gate

National Aeronautics and Space Administration

NASA Worldwide Communications Network

non-real time

nonreturn-to-zero

operational support equipment

Office of Space Science and Applications (NASA)

pyrotechnics arming switch

pulse-code-modulated

programed data processor

plasma

pseudonoise

position

quantitative command

radio frequency

real time

read, write, verify

S-band acquisition aid

Spacecraft Assembly Facility (JPL)

S-band monopulse feedhorn and bridge system

sustainer-engine cutoff

space flight operations

Space Flight Operations Facility (JPL)

separation-initiated timer

simulated midcourse interaction test

serial numberSpacecraft Performance Analysis and Command (Team)

Sun-probe-Mars (angle)

Space Science Analysis and Command (Team)

temperature

336



APPENDIX B

TPS

T/R

TTY

VCO

VECO

telemetry processing subsystem

transformer/rectitier

teletypevoltage-controlled oscillator

vernier-engine cutoff

337



I_!_CED_G PAGE BLANK NOT _"_ '"_

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46 3_-u.$, GOVEm_IMEN'r PRINTING OFFICE: 1_ O--271-409