Exploration of the Moon, The Planets, And Interplanetary Space 1959

8/6/2019 Exploration of the Moon, The Planets, And Interplanetary Space 1959

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, i n . ..-( N A S A CR O R TM X O R AD KiltdBERJ (CATEQORYJ

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REPORT NO. 30-1

EXPLORATION OF THE MOON, THE PLANETS,A N D INTERPLANETARY SPACE

EDITED BY ALBERT R. HIBBS

NASAFILE COPY

JET PROPULSION LABORATORYCALIFORNIA INSTITUTE OF TECHNOLOGY

PASADENA, CALIFORNIAAPRIL 30, 1959


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Copy No.&--

National Aeronautics and Space Administration

Contract No. NASw-6

REPORTNo. 30-1

EXPLORATION OF THE MOON, THE PLANETS,

AND INTERPLANETARY SPACE

Edited by Albert R. Hibbs

d & z d ? f ? / aAlbert R . Hibbs, C h i e fResearch Analysis Section

51

JET PROPULSION LABORATORYCalifornia Institute of Technology

Pasadena, CaliforniaApril 30,1959


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PREFACEPortions of the following report were originated under studies con-

ducted for the Department of Army Ordnance Corps under ContractNo. DA-04-495-0rd 18. Such studies are now conducted for the

National Aeronautics and Space Administration under Contract No.NASW-6.

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Page

1. Introduction .............................................. 1II. Basic Philosophy .......................................... 3

111. Technical Feasibility ....................................... 5A . Flight Mechanics ....................................... 58 . Vehicle Configuration ................................... 14C. Instrumentation ........................................ 18

1. Photography ........................................ 182 . Magnetometers ...................................... 193. Cosmic-ray instrumentation ............................. 194 . Meteor detectors .................................... 195 . Mass spectrographs .................................. 196 . Ion probe

..........................................20

7 . Spectrophotometers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208. Timers ............................................ 20

D. Payload Attitude Control ................................. 201. Methods of obtaining control torques ..................... 212. Low-energy digital control ............................. 22

E. Space .Naviga tion a nd Terminal Guidance . . . . . . . . . . . . . . . . . . . 25

1. Guidance systems .................................... 252. Post-injection radio guidance ........................... 264 . Terminal guidance to the moon .......................... 29

F. Communication Problems and Capabilities . . . . . . . . . . . . . . . . . . 32

E. F.Dobies

M. Gumpel

R.V.Morris

A.R.M . Noton

3 . Self-contained guidance systems ......................... 27

G .M . TetsukaM .6 . 6ain and 0 E.Hull

G . Data Processing ........................................ 34

1. Magnetic tape recording ............................... 342. Expected developments n the magnetic recording field ........ 353. Field data presentation . . . . . . . . . . . . . ..- . . . . . . . . . . . . . . . 354 . Telemetering data reduction ............................ 365 . Final data reduction .................................. 366 . On-site equipment and computer. ......................... 377 . Relay ............................................. 398 . Central computer .................................... 409 . The video problem ................................... 41

. .............................The Tracking of Space Probes 42M.EimerR.C. Hamil ton

1. Space Powe r . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 461. Solar power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 472. Electrochemical batteries ............................... 493. Nuclear power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 504 . Solid-state regulated ac and dc power .................... 53

CONTENTS


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CONTENTS(Contdl

Page...........................................V. Public Reaction 55

V . Scientific Considerations ................................... 59

A

.Background 591. The origin of li fe ..................................... 59

2 . The origin of the solar system ........................... 60B. TheMoon ............................................. 62C. Venus ................................................ 66

E. Other Planets .......................................... 741. Mercury ........................................... 742 . Jupiter ............................................ 7 4

R.W .Davies

R. .Newburn...........................................

N . H Horowitz. CIT

D Mars ................................................. 69

3 . Saturn . . . . . . . . . . . . .. . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . 774 . Uranus and Neptune .................................. 785 . Pluto .............................................. 78

F. Asteroids. Comets. and Meteo r Streams ..................... 781. Asteroids .......................................... 782 . Comets . . . . . . . . . . . . .. . . . . . . . . . . . . .. . . . . . . . . . . . .. . . . 7 93 . Meteor streams ...................................... 81The Interplanetary Mediu m ............................... 82

Charged particles .................................... 822 . Neutral interplanetary gas ............................. 85

Interplanetary dust ................................... 854 . Interplanetary electromagnetic fields ..................... 85

H The Sun .............................................. 86

Suggested Program ....................................... 93Flight Schedule ........................................ 93

B. Description of Typical Payloads 93

1. Lunar miss (Payload No 1. August 196 0) 95

G.M .M .Neugebaur

1.

3 .

. ....................................Supporting Research 88VI .

A............................

J .C. Porter. W . McDonald. M . G . Comuntzis. andM . Gumpel

. . . . . . . . . . . . . . . . . .. ..........

3 .4 .

Escape toward Mars (Payload No 2. October 1960)Escape tow ard Venus (Payload N o 3. January 1961)Lunar rough landin g (Payload No 4. June 196 1)

959898

. ......... ............. . . . . . . . . . . .. Lunar satellite (Payload No 5. September 1961) 98

6 . Venus satellite (Payload No 6. August 1962) 102

8. Mars satellite (Payload No 8. November 1962) 105

. . . . . . . . . . . . . . .. . ................Venus entry (Payload No 7. August 1962) 102

9 . 105. . . . . . . . . . . . .

. . . . . . . . . . . . .ars entry (Payload No 9. November 1962)


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CONTENTS (Contd)

C

Page

10 lunar orbit and return (Payload No . 10. February 1963) . . . . . . 0511 . Lunar soft landing (Payload No 11 . June 1963) ............ 0512 . Venus soft landing (Payload No . 12. March 1964) . . . . . . . . . . 10Development Schedule .................................. 114

Procurement of engineering design data .................. 114Ground test requirements ............................. 114Typical schedules ................................... 115

1.2.3 .

VI1. Conclusions .............................................. 120

1. Fundamental Characteristics of Proposed Vehicles . . . . . . . . . . . . . . . . . . 172 . Altitude-Control Accuracy Requirements ......................... 223 . Typical Propulsion Data for Trajectory Correction . . . . . . . . . . . . . . . . . . . 254 . Communication System Capability ............................... 345 . Suggested lu na r an d Planetary Flight Schedule .................... 946 . Weight Requirements of Payload Components ..................... 947. lunar M i s s (Payload No . 1. August 1960) ......................... 968 . Escape Toward Mars (Payload N o . 2. October 1960) . . . . . . . . . . . . . . . 979 . Escape Towa rd Venus (Pay load No . 3. January 19611 . . . . . . . . . . . . . . . 99

10 . lunar Rough landing (Payload No . 4. June 1961) . . . . . . . . . . . . . . . . . 001 1. lunar Satellite (Payload No . 5. September 1961) .................... 10112 . Venus Satellite (Payload No . 6. August 1962) ..................... 10313 . Venus Entry (Payload No I . August 1962) ....................... 10414 . Mars Satellite (Payl oad No . 8. November 1962) .................... 10615 . Mars Entry (Payload No . 9. November 1962) ..................... 10716 . unar Soft lan din g (Payload No . 11 . June 1963) .................... 10817 . Venus Soft l an di ng (Payload No . 12. March 1964) . . . . . . . . . . . . . . . . . .112

I

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. Report N o . 30-1.9 l e t Propulsion Laboratory

FIGURES

Page

6

7

7

4 . Heliocentric Geometry for Mars Flight ........................... 75 . Mars Arrival Date ........................................... 86 . Heliocentric Distance of Mars at Probe Arrival .................... 87. Heliocentric latitude of Mars at Probe Arrival .................... 88 . Heliocentric longitude of Mars at Probe Arrival . . . . . . . . . . . . . . . . . . . 99 . Heliocentric lon git ude of Earth at the Time of Probe Arrival a t Mars .... 9

10 . Geocentric Distance of Mars a t Probe Arr ival ...................... 911. Heliocentric Central Angle for Mars Trajectory . . . . . . . . . . . . . . . . . . . . 912 . Date o f Heliocentric Injection (Transfer from Geocentric

Conic to Heliocentric Conic) .................................... 1013 . Heliocentric Inclination of the Transfer Ellipse .................... 1014 . Geocentric Hyperbolic Excess Velocity .......................... 1015 . Angle Between Geocentric Velocity Vector an d

1. Approximate Firing Dates for Mars Flights. 1960 Through 1965 . . . . . . .2 . Approximate Firing Dates for Venus Flights. 1959 Through 1965 . . . . .3 . Diagram of Joined Conics .....................................

Ecliptic Plane at Heliocentric Injection ............................ 1116 . Angle Between Geocentric Velocity Vector an d Equator a t

Heliocentric Injection ......................................... 1117 . Geocentric Geometry of Initial Hyperbola ....................... 1 118 . Time Required f rom Geocentric Injection to Heliocentric Injection ..... 1219 . loc us o f Possible Geocentric Injection Points on the Surface of Earth . . . . 1220 . Geocentric Injection Speed ..................................... 1221 . Geocentric Injection Speed, Venus Trajectory ...................... 1322 . Geocentric Distance of Venus at Probe Arrival .................... 1323 . Speed Relative to Mar s of Probe at 2 Radii from Center 1424 . Speed Increment Required for Mars Capture at 2 Radii from Center 14

. . . . . . . . . . . . .. . . . .

25 . Speed Increment Required for Circular Mars Orbi t a t2 Radii fro m Center ........................................... 1415

15

. . . . . . . . . . . .6 . Speed Relative to Venus o f Probe at 2 Radii from Center27 . Speed Increment Required for Venus Capture a t 2 Radii from Center . . . .28 . Speed Increment Required for Circular Venus Or bit a t

2 Radii from Center ........................................... 15


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FIGURES ( Contd 1

Page

29 . Error Coefficient for Vari atio n in Geocentric Injectio n Speed.Mars Trajectory .............................................. 15

30 . Typical Space Vehicles of the Nati onal Program .................... 1631 . Attitude-Control System ....................................... 21

Maintenance Operation ....................................... 2333 . Attitude-Maintenance Operation Without Disturbances ............. 2334 . Attitude-Maintenance Operation With Disturbances . . . . . . . . . . . . . . . .

32 . Phase-Plane Diagram Describing Digit al Attit ude-

23

35 . A Possible Mechanization for Attitude-Maintenance Equipment . . . . . . 2436 . A Possible Optical Sensing Device for Attitu de-Main tenance System . . . . 2437 . A Possible Optical Arrangement for Pointing the

Vehicle at a Celestial Body .................................... 24

Angular Position of Planet Relative to Reference Star . . . . . . . . . . . . . . 28

40 . Guidance for Closest Approach to the Moon ...................... 3041 . Vernier Control of Braking for Soft Impact ....................... 3242 . Data-Handling System with Frequency Multiplier . . . . . . . . . . . . . . . . . 3643 . On-Site Data Preparation Equipment ............................. 3 744 . Spectrometer Scan Showing Possible Bandwidth Compression . . . . . . . . 3845 . Multistylus Recorder .......................................... 3946 . Kineplex Relay System ........................................ 4047 . Central Data-Reduction Facility ................................. 4048 . Trajectory Computation Program ................................ 4349 . Space Probe Tracking Program

51 . lunar-Or bit Sun-Tracking Gim bale d Solar-Cell Assembly . . . . . . . . . . . . . 4852 . Solar Power System ........................................... 4953 . Tapered Concentric Cylinder Thermionic Diode .................... 5154 . Preliminary Design of Thermionic Diode ......................... 52

38 . Measurement of Angular Coordinates (+. CY) to Define

39 . Relative Positions of the Planets ............................... 29

................................. 455 0. Solar-Cell Roll-Control Sun-Seeker ............................... 48

55 . Composite Photograph of the Full Mo on ....................... 6356 . Region of the lunar Crater Clavius (Photographed

Through the 200.in . Telescope) ................................. 64


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FIGURES (Contd 1

Page

(Photographed Through the 100-in. Telescope) . . . . . . . . . . . . . . . . 6557. lo ok in g North from Copernicus Across Mar e lmbrium

58. Crescent of Venus (Photographed i n Blue l ig h t Through the200-in. Telescope) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

59. Mars, Showing Atmosphere (Photographed in Blue lig ht )and Surface Features (Photographed in R ed light, Throughthe 200-in. Telescope) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

60. Jupiter, Showing Ganymede and Its Shadow (Photographedi n Red light Through the 200-in. Telescope) . . . . .. . . . . . . . . . . . . 75

61. Saturn

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .77

62. The Head of Halleys Comet (60411. Telescope, M a y 8, 1910) . . . . . . . . 8063. The Sun, Showi ng l a rg e Sunspots and Fine Structure of the Surface . . . . . 8764. Typical Development Schedule, Payloads No. 1, 2, and 3 . . . . . . . 11 665. Conservative Development Schedule, Payload No. 2 . . . . . . . . . . . 1766. Suggested Schedule of Ma jo r P ayload System Tests . . . . . . . . . . . . . 11 8


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ACKNOWLEDGMENTS

IIn assessing the technical feasibility of the space-

exploration program, the authors of this Report haverelied heavily on the broad background and experienceof the Jet Propulsion Laboratory. The major portion ofour concepts of launching vehicles, payload instrumenta-

tion, communications, guidance and control, structuraldesign, tracking, and da ta processing have come fromthis source. However, we have also made use of numer-ous ideas gained from the reports of other organizationsor from conversations with their personnel. These haveincluded the Army Ballistic Missile Agency, ConvairAstronautics, the Rand Corporation, North AmericanAviation, Inc., the Space Technology Laboratories, theGeneral Electric Company, the California Research Cor-poration, and numerous other organizations.

In assessing the scientific and technological merit ofvarious experiments, we have not only relied on theknowledge and skills of the scientists and t he engineersat the Jet PropuIsion Laboratory, but also have established

many rewarding contacts with scientists at universitiesand other organizations throughout th e country. Valuableassistance has been provided by the Space Science Boardof the National Academy of Sciences, particularly by theGeochemistry Committee led by Dr. Harold Urey and

the Westex Committee headed by Dr. J. Lederberg.

In addition, we have benefited from contact with theNational Acronautics and Space Administration and par-ticularly from participation in the NASA Working Groupfor Lunar Exploration, headed by Dr. Robert Jastrow.

In attempting to estimate the na ture of public reaction,we have established contact with various governmental

agencies which are concerned with this problem, such asthe Central Intelligence Agency and the U. S. Informa-tion Agency. We have also conducted numerous informalinterviews with friends and neighbors who have no directconnection with this program, aside from helping to payfor it.


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The Jet Propulsion Laboratorys Pioneer Iv Payload, Launched March 2, 1959-

The First Successful U. S. Space Probe


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ABSTRACT

The Jet Propulsion Laboratory has undertaken a survey of possible

objectives in a program of exploration of the moon, the planets, and

interplanetary space. This has been combined with a survey of t he

feasibility of engineering developments which would be required by

such an exploratioo program. The results of this study are presented

in this Report.

The Report describes the basis on which the study was conducted,

presents a review of current knowledge about the moon, the planets,

an d interplanetary space, gives a brief summary of t he results of the

Laboratory study on th e feasibility of a program for the exploration of

space, describes a program of lunar and interplanetary flights, and out-

lines the necessary development activities to support the exploration

program. The time scale covered extends from 1959 through 1964.

1. INTRODUCTION

In December of 1958, the Jet Propulsion Laboratorywas requested by the National Aeronautics and SpaceAdministration ro prepare a study of the space explora-tion program. In particular, the Laboratory was asked todesc ribe those portions of the program wherein it felt itmight make the greatest contribution. It was suggestedtha t th e period to be covered by the study extend through1964. Thus, this study was intended to be an outline fo r aLaboratory program over the next 5 years.

Work on this study was broken down into several dif-ferent areas; for example, vehicle development, guidanceand control, tracking an d communications, and SO forth.

One particular area so defined was the study of the scien-tific missions which might be undertaken in this program,in particular, those specific scientific objectives whichmight become th e primary objectives of t he Laboratoryprogram.

The Laboratory program will consist of both the design,

develop ment, and operation of some of t he rocket vehiclesto be used in the space program and th e design, develop-ment, and operation of some of t he payloads which willcarry the scientific measuring devices. In the area of pay-load development, it is the intention of the Laboratoryto concentrate on those payloads designed for lunar and

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planetary investigations, as contrasted to artificial earthsatellites. The area of study of the scientific missions inspace, like the other areas of the study program, wasgoverned by this sta tement of Laboratory intention-con-centration on the moon, the planets, and the spacebetween them.

This study of the scientific missions in space had thefollowing objectives: To tie together the important scien-

tific missions with feasible technical developments andproduce a realistic 5-year program for the scientificexplorat ion of space. The results of this study are beingused in the construction of a Laboratory program for thedevelopment of the necessary rocket vehicles and pay-loads to carry out this program. This Report presents theresults of this s tudy of scientific missions. It includes also

some portions of other areas in the over-all study programwhich were used to assess the feasibility of the proposedscientific program.

The results presented herein are not intended to behard and fast design decisions on vehicles, payloads, orscientific instruments. Furthermore, the schedules pre-sented herein are consistent with scientific potentialitiesand astronomical dates but are not to be interpreted as

program commitments. The results a re intended to be asrealistic as possible on the basis of present knowledge,but it must be kept in mind that further developmentswill undoubtedly change many details of the programin a very significant manner, Thus, these results repre-sent a typical program which can be used as a basis forprogram planning.


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II. BASIC PHILOSOPHY

Three criteria have been selected a s having the mostimportant effects on the program:

1. Technical feasibility

2 . Public reaction

3. Scientific and technical merit

These criteria have been listed, to some extent, in orderof rela tive importance. The question of technical feasi-bility has been the primary consideration throughout thisstudy. Before one can decide that a particular experimentshould be carried out because it is either worthwhile ordesirable or both, one must first make sure that it ispossible a t all.

It is more difficult to determine the relative importsnceof the last two criteria. Occasionally, they both lead oneto the same conclusion. For example, the search for lifeon another planet is of the greatest scientific importanceand, at the same time, is encouraged by a strong publicinterest. On the other hand, and in this same area, theproblem of decontamination of planetary probes may beapproached in quite a different manner by scientific and

nonscientific groups. Scientific groups recognize the needfor decontamination as primary for the success of futureexplorations for life forms. However, the public may ques-tion whether or not it is worthwhile to postpone a Marsshot, for example, for 2 years so that problems of decon-tamination may be fully solved.

In approaching this particular problem, we have takenthe rather optimistic point of view that (1) the problemsof decontamination can be successfully worked out intime to meet the proposed schedule, and ( 2 ) the publiccan be educated as to the importance of this problem SOth at they will neither begrudge the amount of moneyspent on its solution nor object to the limitation of experi-ments resulting from its possible lack of solution.

This example is characteristic of the manner in whichthe basic philosophy of this Report has been applied. Anattempt has been made to select the possible, worthwhile,and desirable scientific experiments in the program fo rthe exploration of the moon, the planets, and inter-planetary space. A representative flight program has beenconstructed in which these experiments will be under-

has been made to select the particular areas of the pro-gram which appear to need the most urgent attention.

In applying the criteria of technical feasibility, we haveassigned specific missions to the launching vehicles anddeveloped representative estimates of the payload weights

which these vehicles can carry to the various objectives.It must be recognized that these weights are far fromdefinite and are intended to be representative of the typeof vehicle availnble at a particular time during the pro-gram. Furthermore, we have broken down the payloadweights into the various major payload components andhave listed representative weights which might beascribed to each of these components. Here , again, these

weight estimates a re far from definite, but they are reason-able and serve to point out th e problems of weightlimitation.

We have also attempted t o estimate the degree o f com-plexity which can be assumed for each cf thcse com-ponents, including the scientific instrumentation. Thisestimate has been made with the clear realization thatthe need for extreme reliability limits the degree of nov-

elty which can be introduced into the various payloads.In investigating the scientific objectives of the program,

we have also considered the purely tcchnologicnl prob-lems which must be solved if the whole program is to besuccessful. The telemetered information sent back fromthe payload must contain the results of measurementsmade for purely engineering objectives. In order to makepossible the development of increasingly complex pay-

loads, we must develop a background of engineeringdesign data on the behavior of materials and componentsin the completely new environment of empty space andthe atmospheres of other plane ts.

The program developed from this basic philosophy andthe resulting investigations has been projected 5 yearsinto the future . The validity of estimates of technical

feasibility, public reaction, and scientific and technicalmerit is naturally degraded a s we proceed further andfurther into the future. It was felt that, on the basis ofpresent knowledge, a reasonable prediction could bemade for a period of no more than about 5 years.

taken over the next 5 years. On the basis of t he inves-tigations involved in setting up this program, an attempt

For this reason, present study does not include consid-eration of the man-in-space program. It is' not reasonable


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to assume that the n~an-i i i -spa~e lUgldlll would have anydirect bearing on the technical problems involved in theexploration of the moon, the planets, and interplsnetaryspace over t he next 5-year period. Although many of th escientific and technological experiments which will becarried out during this program will have a definite bear-ing on the design of the vehicles which carry men to theplanets, it is not felt that this objective is in any wayinconsistent with the already stated crite rion of scientificand technological merit.

Some time after th e close of this first 5-year period, theman-in-space program and the interplanetary space pro-gram will gradually merge.

Certainly, a manned laiidiiig UI I aiiother planet is oneof the most important objectives of a long-range program.Regardless of how clever we become with remote measur-ing devices, one hard-rock geologist landed on t he moon,for example, would be worth 'many tons of automaticequipment. The public interest in full-color ph*,tographstaken by a remote camera on the surface of Mars will belittle as compared to the wild reception which will greetthe first crew of astronauts which returns alive from thatplanet.

It is the basic philosophy of this study to develop pos-sible, sensible, and desirable beginning for a programwhich will eventually take man to the planets.

~

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II . TECHNEAL FEASlB ILlTY

A. Flight Mechanics

Th e feasibility of any particular planetary experimentdepends upon many factors. Among the most importantar e (1) the payload which can be carried to the targetplanet, (2) the accuracy with which a n intersection with

the target planet can be achieved, and (3) the capacityof the communication system to return to earth theinformation gathered.

The capacities of a given rocket and payload systemin any of these areas can be determined only with thehelp of an analysis of possible in terplanetary trajectories.This analysis must determine the speed with which thepayload must be launched from earth, the sensitivity of

its final position to errors made in launching position anddirection, the possibilities of correcting the course duringthe coasting period from earth to the target planet, andthe distance between earth and the target planet at thetime when experimental information is to be relayedback. In order to establish a flight schedule, it is neces-sary to know the precise times at which planetary experi-ments can be carried out; an attempt to launch a payload

at a particular planet is practicable for only a few daysin each synodic period.

An accurate analysis of interplanetary trajectoriesrequires the solution of the n-body problem, which inturn requires numerical integration on a high-speed elec-tronic digital computer. However, for many purposes ofa preliminary analysis, an approximate scheme is avail-able which permits analytical solutions. This scheme,

which might b e called the principle of joined conics,makes use of the fact that th e various bodies of interesthave their most important effect on the flight during dif-ferent periods of the flight time. Thus, immediately afterthe payload is launched from earth, the gravitationalfield of t he ea rth itself dominates the trajectory. After thepayload has coasted away from earth, the gravitationaleffect of t he sun takes over as the most important force

governing the shape of the trajectory. As the vehicleapproaches the ta rget planet, t he field of the planetbecomes a dominant effect.

Th e principle of joined conics gives an approximationto intcrplanetary trajectories in the following way: In thevicinity of t he earth, after the burnout of the last stageof the rocket launching system, the trajectory is approxi-

mated by a geocentric conic. For interplanetary trajec-tories, this conic has the shape of a hyperbola. F or regionsfar from earth, the trajectory is approximated by a helio-centric conic which is, in general, an ellipse, at least forthe flights considered in this study. In the vicinity of

the planets, the trajectory is again a planet-centeredhyperbola.

The points at which the various conics are joinedtogether can be chosen somewhat arbitrarily. For thisstudy, the point chosen was a t a distance of approximately2 X 1 0 km from earth, which is in the region wherethe effect of the suns gravity, as compensated by thecentrifugal acceleration of a coordinate system moving

with the earth , is approximately equa l to the effect of theearths gravity.

A preliminary analysis of interplanetary trajectories hasbeen carried out under the simplifying assumption thatall the planets move in circles in a single plane. On thisbasis, an estimate has been made of the times a t whichplanetary probes can be launched. The results of thisanalysis are shown in Figs. 1 and 2, wherein firing dateis plotted against time of flight for trajectories directedtoward Mars and Venus, respectively. The time periodcovered is from 1959 to 1965; this includes three synodicperiods for Mars and four for Venus.

A more exact investigation of trajectory requirementsfor a Mars probe in 1960 has been completed. This inves-tigation is instead intended to bridge the gap between a

simple heliocentric conic analysis restricted to the planeof the ecliptic and a comprehensive conic analysis to bemade using the IB M 704 computer, which will soon beavailable. The present analysis extends the in-the-planeanalysis to the more realistic three-dimensional case. Italso considers the geocentric portion of t he flight. Oneof the most serious limitations of this investigation is thatthe only heliocentric transfers considered are those in

which the ea rth is at the perihelion of the transfer ellipse,as shown in Fig. 3.

From the earlier (in-the-plane) analysis a Mars-arrival-date time interval was selected, as shown in Fig. 1. Onvarious days throughout the in terval, thc coordinates ofMars and the coordinates of ear th were obtained froman ephemeris. The geocentric distance of Mars was com-


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I965

24 JAN

22 JAN

20 JAN

18 JAN

16 JAN

14 JAN

12 JAN

10 JAN

8 JAN

6 JAN

DATE OF FIRINGI 9 6 2

6 DE C

4 DE C

2 DE C

30 NOV

28 NOV

26 NOV

2 4 NOV

22 NOV

20 NO V

18 NOV

I 9 6 0

17 OC T

15 OC T

13 OC T

I I OC T

9 OC T

7 OC T

5 OC T

3 OC T

I OC T

29 S E P

T I M E O F F L IG H T, days

Fig. 1. Approximate Firing Dates for Mars Flights, 1960 Through 1965

puted from the difference in these coordinates. This is thedistance over which information must be telemeteredfrom the probe to the earth at the time of arrival. Theheliocentric distance of Mars was also computed. Th eheliocentric geometry is shown in Fig. 4.

Elliptic trajectories were computed with the aid of anelectronic computer. It was assumed that the probeentered the heliocentric transfer ellipse at the perihelion

of th e ellipse. It was further assumed that the heliocentricdistance of perihelion was 1 A. U. (astronomical unit: themean distance from the earth to the sun). For variousvalues of perihelion velocity, time intervals were com-puted from the perihelion distance of t he ellipse to the

heliocentric radial distance of Mars at th e time of arrival.Also computed was the heliocentric central angle sub-tended by t he heliocentric transfer.

For each value of heliocentric Mars distance, a plotwas constructed of transfer time vs central angle. Eachvalue of Mars heliocentric radial distance corresponds toa particular Mars arrival date. A series of heliocentricinjection dates was selected, and the ephemeris was con-

sulted for th e heliocentric coordinates of ea rth on thesedays. The heliocentric angular distance between Mars atth e arrival date and each of the positions of th e earth atthe possible injection dates was computed. Also, the timeseparation was compute d. This curve of time difference

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1965

14 NOV

12 NOV

IO NO V

8 NOV

6 NO V

4 NOV

2 NO V

31 OC T

29 OC T

DATE OF FiRiNGI 9 6 4

9 AP R

7 APR

5 APR

3 AP R

I AP R

3 0 M AR

28 M AR

26 MA R

2 4 M A R

I 9 6 2

3 SE P

I SE P

30 AUG

28 AUG

26 AUG

24 AUG

2 2 AUG

20 AUG

18 BUG

1961

27 JA N

25 JA N

23 JA N

21 JA N

19 JA N

17 JA N

15 JA N

13 JA N

I I JA N

I 9 5 9

2 3 J U N

2 1 J UN

19 JU N

I 7 JUN

15 JU N

13 JU N

I I JU N

9 JU N

7 JU N

TIME OF FLIGHT, days

Fig. 2. Approximate Firing Dates for Venus Flights, 1959 Through 1965

HELIOCENTRICTRANSFER ELLIPSE

JUNCTION POINTAT PERIHELIONOF TRANSFER

GEOCENTRICHYPERBOLA

FARTH-

Fig. 3. Diagram of Joined Conics

HELIOCENTRIC CENTRAL ANGLE7

DATE

DATE

VERNALEQUINOX

LHELIOCENTRICONGITUDE OFEARTHON LAUNCH DATE

Fig. 4. Heliocentric Geometry for Mars Flight

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I .66

?a

ifIL[L

I-a 1.65,v)ILarLL0

wVz

a

a1.64

0[LI-zWu

-I

Wr

e

1.63

vs angular separation was plotted on the previously con-structed curve from the families of ellipses. The point ofintersection indicated the ellipse required and the helio-centric injection date necessary for arrival at Mars whenMars was at the selected heliocentric distance.

By selecting a particular day for the Mars arrival date,the following quantities can be determined: (a) Marsarrival date (Fig. 5 ) , (b) heliocentric distance of Mars at

arrival (Fig. 6) , (c) heliocentric latitude of Mars a t arrival(Fig. 7), (d ) heliocentric longitude of Mars at arrival(Fig. 8), (e) heliocentric longitude of earth at arrival (Fig.9), and (f ) geocentric distance of Mars at arrival (Fig. 10).

When the intersection point is determined on thepreviously described crossplot, the following quantitiescan be determined: (a) heliocentric central angle (Fig.l l ) , b) heliocentric transit time (used as an abscissa in

other plo ts), (c) date of heliocentric injection (Fig . 12),(d ) heliocentric inclination of the transfer ellipse (Fig .13), and (e) heliocentric velocity of the probe at helio-centric injection.

/

/

120 130 140 150 160 170 180 190

HELIOCENTRIC TRA NSIT TIME ,days

Fig. 5. Mars Arrival Date

HELIOCENTRIC TRANSITTIME, days

Fig. 7. Heliocentric latitude of Mars a t Probe Arrival

Heliocentric injection is easily analyzed in this investi-gation since it was assumed that the earth moves atconstant speed in the plane of t he ecliptic with no radialvelocity component. By further simplifying the helio-


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HELIOCENTRIC TRANSIT TIME,d a y s

Fig. 8. Heliocentric long itud e of Ma rs at Probe Arrival


Fig. 9 . Heliocentric longi tude of Earth at the Time ofProbe Arrival at Mars

HELIOCENTRIC TRANSIT TIME,days

Fig. 10. Geocentric Distance of Mars at Probe Arrival

ol

0

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a

a

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Fig. 1 1. Heliocentric Central Angle for Mars Trajectory

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HELIOCENTRIC TRANSIT T IME ,days

Fig. 12. Dat e of Heliocentric Injection (Transfer fromGeocentric Conic to Heliocentric Conic)

H E L I OC E N T R IC T R A N S I T T I M E , d a y s

Fig. 13. Heliocentric Inclination of the

Transfer Ellipse

centric injection, thy prol~t. w i t s usurncd to have no

heliocentric radial velocity component at injection. Toclarify the last statemcnt, it must be rerncmbcwd thatth e calculiitions of the time differences and angle differ-ences of transfer are based on the assumption that theprobe entered the transfer ellipse at its perihelion. Theintersection of time-angle curves required that the peri-helion be at zero degrees heliocentric latitude, or inth e plane of the ecliptic. In general, tlie inclinationof tlie transfer ellipse is not zero, and the geocentricdistance of heliocentric transfer is not zero (2 X 10'' km) .Therefore, the pro be is neither at tlie perihelion of t hetransfer ellipse nor in the plane of the ecliptic at the timeof heliocentric transfer. In a rigorous analysis, these fac-tors would be included; in this study they are unimportant.

Th e differencc lx tw ce n the hclioccntric velocity of th eearth a n d the reqiiircd heliocentric velocity of the probe

H E L IO C E N T R IC T R A N S I T T I M E , d a y s

Fig. 14. Geocentric Hyperbolic Excess Velocity

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1 1 1 1a a 30

120 130 140 150 160 170 180 190HELIOCENTRIC TRANSIT TIME,days

Fig. 15 . Angle Between Geocentric Velocity Vector andEcliptic Plane a t Heliocentric Injection

can be computed, both in magnitude and direction. Themagnitude is the speed residual of the probe afte r it hasescaped the ear th a nd is called the geocentric hyperbolic

excess velocity (Fig. 14). To find the angle that thevelocity vector makes with the plane of the equator atthe time of heliocentric injection (Fig. 15), the anglebetween the velocity vector (a t the time of transfer tothe heliocentric conic) and the plane of the ecliptic (Fig.16), the he liocentric longitude of t he earth, and the anglebetween the plane of the ecliptic and the plane of th eequator are used. Since the probe is traveling radiallygeocentrically at the time of heliocentric injection, thisangle is the sublat itude of the probe at heliocentric injec-tion. The geocentric geometry is shown in Fig. 17.

It is next assumed that the probe enters the geocentrichyperbola at the apex at an altitude of 350 km. It is no wpossible to compute the geocentric transit time; is., thetime required to go from geocentric injection to helio-centric injection (Fig. 18). It is also possible to computethe geocentric central angle from geocentric injection toheliocentric injection.

A circle is now visualized which is drawn on the sur-face of the ear th in the following way (see Fig. 19):Locate first the subtransfer point, the point on the earthgiven with t he intersection of the surface of a radius

HELIOCENTRIC TRANSIT TIME ,doys

Fig. 16. Angle Between Geocentric Velocity Vector and

Equator at Heliocentric Injection

SUBTRANSFER POINTAT HELIOC ENTRIC INJECTION

\GEOCENTRICCENTRAL ANGLE

I

Fig. 17. Geocentric Geometry of Initial Hyperbola

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Fig. 18. Time Required from Geocentric Injection to

Heliocentric Injection

7UBTRANSFER POINT

LLOCUS OF ANGLESUBINJECTIONPOINTS

drawn from the center of t he earth outward in tlie direc-tion of the geocentric velocity vector a t the t ime ofheliocentric injection (i.e., transfer to the heliocentricconic), Measure away from that point on the surfacethrough an angle equal to the geocentric central angle.The locus of all points on the surface separated from thesubtransfer point by this angle is a circle on the surface.This is the locus of all points over which geocentricinjection could be carried out.

In this visualization it is assumed that the earth is notrotating around its axis. Thus, the locus constructed inthis manner is actually a locus fixed in geocentric inertialspace rather than fixed on the surface. In order to achievethe necessary transfer orbit to Mars in accordance withthe geometry used in this study, burnout of t he last stageof the launching rocket must occur at a point above this

circular locus.

It is also possible to compute the speed which mustbe achieved at burnout of the last stage in order to obtainthe correct hyperbolic excess velocity. The geocentricinjection speed is shown in Fig. 20.

Fig. 19. locus of Possible Geocentric Injection Points

on the Surface of Earth

HELIOCENTRIC TRANSIT TIME, days

Fig. 20. Geocentric Injection Speed

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This analytical treatment of the approximation to inter-planetary orbits can be carried through for any planetarytarget. The results are used to estimate the performancerequirements of the rocket, th e time of flight, and thedistance between the earth and the target planet at thetime of arrival.

The results given here apply to a Mars mission in thefall of 1960. Similar results are being obtained for the

Venus mission early in 1961 and succeeding flightstowards Mars and Venus. The methods used and thetypes of results obta ined are similar to those reportedhere. A simplified two-dimensional analysis has beencarried through for Venus. Results have been obtainedgiving the geocentric injection speed necessary for aVenus trajectory (Fig. 21) and the distance between theear th and Venus at the time of arrival at Venus (Fig. 28).

13.(

0

cE

x

12.5

5:mw3n

k5a5nW

cnz

I-oW7

12.0

0

zI!aI-zw" 11.5WB

11.0.

H E L I O C E N T R I C T R A N S I T T I M E ,days

Fig. 21. Geocentric Injection Speed, Venus Trajectory

7 5

0

0

E 70

1-

cc 65a

X

x

5(I

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a-

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4 0

71

I I

90 I O 0 I O 120 130H E L I O C E N TR I C T R A N S I TTIME, days

Fig. 22 . Geocentric Distance of Venus at Probe Arrival

An additional analysis has been carried out whichshows the behavior of the probe in the vicinity of thetarget planet and, in particular, the following:

(a) the speed in a Mars-centered system at the timethe prob e arrives at a distance 2 radii out from thecenter of Mars (Fi g. 23) ;

(b ) th e speed increment necessary t o achieve captureby th e gravitational field of Mars when the speedincrement is applied at the 2 Mars radii distan ce(Fig. 24); and

(c) the speed increment necessary to achieve a circu-

lar orbit a t this distance (Fi g. 25) .

Similar results (Figs. 26-28) have been obtained for thebehavior of the payload in the vicinity of Venus. Theseresults are necessary in order to compute the size of theretro-rockets necessary to establish an artificial satellitearound the target planet.

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13

12

V

\

1

w I I

E

LOrd IOL

5

E 9

Wm

aLL0

0WW

b e

7

6


Fig. 23. Speed Relative to Mars of Probe at2 Radii from Center

It is also possible to estimate the sensitivity of missdistance at the target planet to errors made during thelaunching phase. These numbers depend critically uponthe shape of the trajectory; for the Mars trajectory ingeneral, they have t he following values: An error of 1 degin th e direction of the geocentric velocity vector will

contribute between a few hundred thousand and 2 X 10,km to the miss distance. An error of 1 meter/sec in thespeed of geocentric injection will contribute between10,000 and 50,000 km to the miss distance at Mars (seeFig. 29).

6. Vehicle Configuration

The most important factors in determining the tech-nical feasibility of a space mission are the performanceand reliability of the rocketry system carrying the pay-load. The mission schedule presented herein is designedto fit into the National Space Vehicle Program (Fig. 30).


Fig. 24. Speed Increment Required for Mars Capture at2 Radii from Center

I

V

m

1E

Prd

L

t-akm

aa

E

e

5

[L0

3u

uII:

lzW

aV

z0WW

ula

HELIOCENTRIC TRA NSIT TIME,days

Fig. 25. Speed Increment Required for Circular MarsOrbit at 2 Radii from Center

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Fig. 26. Speed Relative to Venus of Probe at

2 Radii from Center


Fig. 27 . Speed Increment Required for Venus Captureat 2 Radii from Center

HELIOCENTRIC TR ANS IT T IME,days

Fig. 28. Speed Increment Required for Circular VenusOrbit at 2 Radii from Center

120 130 140 150 160 170 180 190

HELIOCENTRIC TRANSITTIME, days

Fig. 29. Error Coefficient for Variation in GeocentricInjection Speed, Mars Trajectory

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ADVANCED VEGA SATURNGA

Fig. 30. Typical Space Vehicles o f the National Program


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The dates assigned fo r use of t he vehicles do not, ingeneral, begin with the date at which the particularvehicle is scheduled for its first flight test. Only theVega will be employed as a space vehicle on its first test.It is assumed that the application of the AcZvnnced Vegaand the Saturn to the space exploration program willfollow some months after the first flight tests of thesevehicles.

C. Instrumentation

1. Photography. Of the instrumentation proposed forthe missions in the NASA study, photographic equipmentis probably the most important. From an astronomicalstandpoint quite a number of questions can be resolvedby adequat e photographic data, particularly if it becomespossible to take pictures in very selected portions of thespectrum, e.g., visible light, ultraviolet and infrared.

Aside from weight constraints, the most difficult prob-lems in utilizing photography in deep-space explorationare the following:

1. Because of the presence of the Van Allcn belts,taking film away from the vicinity of t he ear th withoutappreciable radiation fogging is difficult.

2. Because of the limited communication bandwidthavailable at p lanetary distances, it will be somewhat of aproblem to transmit the pictures. It must be realized that,since even crude pictures contain the order of 10" bits ofinformation which must be telemetered back with anavailable information rate sometimes as low as 10 bits persecond, it will take a little more than 20 hours to do. Nat-

urally the total number of bits in a picture is tied to thedefinition or raster required as well as the number ofshades of gray, and a degradation of these requirementswill make the problem somewhat easier.

3 . Th e accuracy with which th e attitude of a vehiclecan be controlled enters into the problem of pointing thecamera. If th e film employed is somewhat insensitive to

radiation, it will be comparatively slow. It t hen becomesnecessary to maintain this atti tude with great precisionover an appreciable time.

The present state of the art makes it possible to takepictures ou t in space employing a raster of 200 X 200 linesper in., with a line resolution along the line of 512 linesper in. The capability of the ground system is 1200 X 1200

lines per in. This resolution is quite far from that whichcan be expected in the years to come. It is anticipatedthat , by continuing the development of the present system,it can be improved by a factor of 2 within a year. Thepresent camera is able to take 6 pictures, develop themand telemeter them back at a rate of 1 picture per hour.The actual effect of radiation in the belts upon the per-formance of this system is yet to be found, and the bestprocedure will probably be to expose some film to a

known test pattern, transport it through the radiationbelt, and then develop and telemeter it back.

There are a number of ways to deal with these prob-lems, and various ways are under consideration to cir-cumvent them. Regarding the radiation-belt problem, itshould be possible to develop a film with an increaseddynamic range to limit the influence of radiation. An

alternative is to use an electronic film based on solid-statephenomena; and research should be accelerated in thisdirection. This would simplify the procedure of develop-ing t he film, which with normal film is not easy to accom-plish on an automated basis out in space.

Th e alternative to picture-taking is television; a numberof systems have been suggested , and some of them havebeen carried through partial development. One is a vidi-con system which allows a picture to be taken and thenstored electronically. The scanning rate of this picture isthen regulated to conform to the available informationrate. The resolving power and weight a re still problems tobe solved. At the present time, it is possible to obtainvidicon tubes with a sensitivity in the red region of thespectrum; development has been started to extend thesensitive region of these tubes to the far infrared. Devel-

opment of high-resolution television has been started i;industry. These systems have rasters of 2500 X 2500 linesper in., with spectral sensitivities that can be chosen inranges from the ultraviolet to the far infrared.

Th e communications problem may be solved by p ropercoding of the telemetering . It is not necessary to telemeterthe information on an absolute basis; it is possible totelemeter th e changes in information from the previouslysent information and thereby make it possible to sendmore information with a given bandwidth. All in all, anumber of promising possibilities exist in the field ofspace photography and/or television, but considerabledevelopment will be necessary before high-grade photog-raphy can be accomplished.

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2 . Magnetometers.Fluxgate or nuclear-precession mag-netometers for space exploration ar e already available andmore refined instruments with greater sensitivity andreliability will be ready for space flight within 6 monthsfrom now. Among the types that appear most promisingare the alkaline-vapor magnetic-resonance type, with asensitivity of lo-' gauss, and the nuclear precession-typerecently improved by Russian scientists. Magnetometersof an extremely rugged construction capable of with-stand ing landing shocks in the vicinity of 1500 g are underdevelopment at the present time. These magnetometerswill permit placing stationary instruments on the moonprior to the soft-landing phase of the program.

3 . Cosm ic-ray instrumentation. The number of experi-ments proposed in this particular field is very great and

of considerable interest to both cosmologists and astro-physicists. The instrumentation will require considerabledevelopment of individual components and transducersand further system development is needed.

In the first category, a few items slated for develop-ment, or rather redevelopment, should be mentioned:(1) cintillation detectors with adequate sensitivity and suf-ficiently low noise, (2) sensitive photomultipliers ruggedenough for the boost phase, (3) transistorized pulse-heightanalyzers with low power requirements, good reliability,and very good discrimination, and (4) ruggedized ioniza-tion chambers and counters for rough-landing vehicles.

The second category (system development) includessuch things as system engineering, packaging, and opti-mizing the da ta received in comparison to the power andweight penalties implied by the instrumentation. A greatnumber of these experiments have already been done onearth and must now be performed out in space. Develop-ment of the instrumentation or these experiments seemsto be progressing reasonably, but a certain amount ofre-engineering will have to be done, particularly in thefield of automation and self-calibration, n order to rendera maximum of da ta for the weight and power expended.

4 . Meteor detectors. The experiments flown to datehave shown tha t the concentration of micrometeorites isvery nearly that which was originally anticipated. Thegauges that have been employed in this field have beenvery primitive, giving data only on a go/no-go basis. Fo ra comple te study of interplanetary material it is necessary

to measure not only the abundance but also the totalmomentum of micrometeorites and also, if possible, theirdirection. This will ncccssitatc dcvclopmcnt of trans-ducers of a more complicated natur e. Gauges based onthe established leak ra te of spheres of appreciable sizewould be able to indicate the number of puncturingimpacts per unit area and time, and gauges based onsecondary emission would give the total momentum. Thiswill take considerable development, but it will take stillmore engineering to construct simulation devices to testthese gauges before they are sent aloft. The problem ofmeteor detection, which will be important in flights tocome, must be solved now.

5 . Mass spectrographs. A number of experiments usingmass spectrographs have already been made in the field

of upper-air research. These instruments have been carriedby rockets and have obtained some data; however, con-siderable development work remains to be done if theseinstruments are to be able to analyze the composition ofgases in the neighborhood or on the surface of a planet.Conventional magnetic-deflection instruments are com-petitive in almost every way with the best rf instrumentsbeing used in rockets. They do have the advantage of

having been long tested in precise gas analysis so thatthey give more reliable results, and research and develop-ment along these lines should be stimulated nnd sup-ported in addition to continued experimentation with rfinstruments. With available instruments of both types,complications arise in measuring the neutral matter inspace because of the very low pressures tha t exist there.The lowest pressure at which a magnetic-deflectioninstrument will give an acceptable current lies aroundlo-" mm Hg and the corresponding figure for an instru-ment of the rf type is l O - ' O mm Hg. These limitingpressures are far too high; it becomes necessary to usea matter accumulator to collect material for a long periodof time and then release it into a mass spectrometer foranalysis. The use of titanium as a matter-collection devicehas been suggested, but considerable work must be doneto understand the basic mechanism of this metal as well asto investigate the reversibility of the reaction. It mightvery well be that titanium will not be used, but prelimi-nary study shows it has possibilities.

For the electronics used in conjunction with theseinstruments, some development will be necessary, mainlybecause of the power and weight constraints associated

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with the payload, as well as the problems of outgassingthe payload to prevent gasses from that source frominterfering wi th t he measurements.

When these instruments are used to analyze the com-position of an unknown atmosphere, a fur ther problemarises in regulating the pressure. The output figure iscritically dependent on the maintenance of pressure at aknown constant level. While all these problems are by nomeans insurmountable, much development will be neededbefore reliable figures can be derived from this type ofinstrumentation.

6. Zon probe. Determination of the properties of thecorpuscular radiation from the sun is a scientific mission

of great interest. Instrumentation to measure the char-acter and intensity of this radiation must be developed.A modified Langmuir probe or an improved model of theprobe used in Sputnik I l l could be used. In all cases, thistype of instrumcntation will need elcctrorneter tubesrugged enough to sustain the boost phasc and a t the sametime much less sensitive to emission variations derivedfrom variation in the filament supply employed. In the

same category are eqiiipments to measlire the direction,energy spectrum, polarity, and flux of these particles.These measurements are important not only from an astro-physical standpoint but also because of the expected con-tribution of these particles to the solar pressure on objectsof reasonable area, such as antennas or solar panels. Thispressure will have to be taken into account in theguidance and attitude control of space vehicles.

7. Spectrophotometers. These instruments have beeninvaluable tools of astronomers on the earth and with theadvent of deep-space exploration more elaborate require-ments for this type of instrumentation will emerge. Farabove our atmosphere, with an undistorted view of theobjects in question, these instruments will be called towork in the extreme ends of the spectral range. A broad

development program will be necessary to adapt theseinstruments to space flight. Some preliminary work alongthese lines has been started, bu t a more more general pro-gram is required. Consider the problem of assigning aspectrophotometer to work in the infrared portion of thespectrum. Prism materials that will be able to stand thevibration an d acceleration of the boost phase d o not now

exist. A broad investigation of the materials question willbe requir ed, as well as development of techniques ofautomation and self-alignment for these instruments.Employment of such instruments for space explorationwill also, in most cases, put severe requirements on theatti tude control of t he payload, since for the normal casethe cone angle to the target will be very small. If thisquestion can be solved and the angle maintained forreasonably long periods of time, requir ements on theinformation rate needed to telemeter the results shouldnot be too severe. Development of sensors for aimingand tracking these instruments will be necessary, sincethe information from the spectrophotometer will, by itsvery nature, be difficult to use as the criterion for correctalignment to the target.

8. Timers. In addition to the timing requirements inthe boost phase, the payload itself will also have certaintiming requirements. The timing functions required ofthese timers include cycling the transmitter, switchingtelemeter channels when the number of experiments inthe payload exceeds the number of channels available,acting as general logic circuitry in the payload, etc. Thetimers employed must, by the nature of the mission, havea long life. Associated with logic circuitry, they shouldoperate in the payload with their own independent powersupply. A reasonable amount of development work willhave to be performed for this timing problem in all itsdetail.

The projected timer will have to be a solid-state typebecause of the power constraint, and should be able tobe programmed easily, both as a one-shot timer and as arecurrent one. The over-all timing cycle will depend onthe mission; the timer should be able to function for atleast 6 months. The over-all power drain should be nomore than 200 mw. The accuracy should run better than1% and the incremental accuracy better than 34% in thetotal environment experienced.

D. Puyloud Attitude ControlControl of the angular orientation, or attitude, of a

space vehicle may be required for (1) guidance observa-tions, (2 ) scientific measurements, (3 ) radio communica-

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b

tion: (4) aiming solar cells at the sun, and ( 5 )maneuvering.Attitude-control subsystems for use with comparativelyhigh-thrust chemical rocket stages are referred to as auto-pilots and are not discussed here. The present discussionrefers only to subsystems intended to operate duringperiods of low thrust and/or coasting flight (when nopropulsive or maneuvering thrust is applied).

MOMENTUM-CHANGINGDEVICETTITUDE SENSING b CONTROL COMPUTER b

The function of an attitude-control system is to main-

tain prescribed relationships between reference directionsin the controlled body and external reference directions.An attitude-control system has four major subdivisions,as shown in Fig. 31. The attitude-sensing devices mayinclude radiation detectors, such as photocells, and iner-tial elements, such as gyroscopes. The la tter typ e ofdevice is limited by its inability to maintain a referencedirection over a long period of time. The radiation detec-

tor can maintain a reference direction over an indefiniteperiod, as can a sun-seeking device. Furthermore, accu-rate determination of direction is possible with opticaldevices, provided they are not hampered by the size ofan extended source. However, the problem of identifyingthe radiation source to be used as a reference does exist.

The control computer relates the changes in attitude

information to the appropriate corrective actions. Thecomputer may contain both linear and nonlinear opera-tions, the specific arrangement being determined by themission of the vehicle.

The momentum-changing device adds or removesangular momentum from the hull as changes in its angularvelocity are commanded by the control computei. Exam-ples of such a device are sets of f lywheels or of gas jets.The energy supply is necessary for the operation of thesensing equipment and control computer, as well as forproducing the control torques that change the angular

momentum of the hull. This energy might be supplied bysolar cells or other electrical sources, supplemented bytanks of gas or monopropellant.

1 . Methods of obtaining control torques .Three methodsof obtaining control torques have been investiga ted: (1)gravitational field effects, (2) removal of angular momen-tum from the vehicle, and (3) transfer of angular momen-tum between the hull and attached devices. The lattertwo types of system will be referred to in t he followingmaterial as removal and transfer systems, respectively.

Because the gravitational field of a celestial body isinversely proportional to the square of th e radial distancefrom the body, the attractive forces on particles of equalmass at different radial distances are not the same. Hence,for a body with a nonuniform mass distribution, there

exists a spring-like restoring torque which tends to keepthe body aligned in a particular direction relative to thelocal vertical. For a body of revolution, this torque isproportional to twice th e angle between the axis of revolu-tion and the local vertical. The restoring torque is alsoinversely proportional to th e cube of the radial distancefrom the body; hence, it is only useful at relatively smallradial distances for such tasks as satellite attitude con-

trol. A t the surface of the ear th, for example, this torqueis capable of causing angular accelerations of up to1.5 X rad/sec2.

A distinction can be drawn between attitude-controlsystems intended only to maintain the angular attitudeof t he vehicle in a fixed direction with respect to a givenreference and systems that are intended to vary the atti-tude of t he vehicle in accordance with commands.Systems utilizing spin stabilization or differential gravityeffects are of the former class, and ar e useful only inrather specialized cases. Command systems, at least about

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certain axes of the vchiclc, will be required for most

space missions. Some operations that utilize a command-type attitude-control system are listed in Table 2 .

For a command system, a controllable energy sourcenot dependen t on the attitude of t he vehicle is necessary.Such sources include a nuclear electric power supply;tanks of pressurized gas or propellant; a nd , in manyinstances, a solar-electric power supply. The requiredtorques may be obtained from the expulsion of mass, orby accelerating or precessing flywheels. Of the severalfactors affecting the choice of energy supply, weight isone of the most important.

Studies of vehicles in the 300- to 1000-lb class haveled to the following tentative conclusions regarding thechoice of energy supply: The studies are based on the useof pressurized gas for momentum removal and of accel-

erated flywheels for momentum transfer. Torque levels of0.2 ft/lb or less are sufficient for the attitude control ofa 1000-lb vehicle during coasting flight if no rates inexcess of 0.05 rad/sec about any axis are required. Theselevels are too low to justify th e use of a hot-gas system.To date, no comparison has been made between accel-erated and precessed flywheels.

In most cases, the attitude-control system will contain

a gas-jet system. For voyages of more than 60 to 180days during which the attitude is held constant orchanges at rates less than 6 deg/hr, it is economical(weightwise) to use a momentum-transfer system forattitude maintenance, reserving the removal portion ofthe system to make more rapid maneuvers and to cancelthe effects of initial conditions and disturbances. Atransfer system alone is feasible from the weight stand-

point only when the initial rates and rates built up as aresult of external disturbances are extremely small. For

thc removal of an initial rate, a momeniiiii-i-remov~l

system is more economical (weightwise) than a transfersystem because, in the latter type of system, conserv a ' onof angular momentum necessitates a permanent changein flywheel speed or orientation in accelerated and pre-cessed systems, respectively. For the accelerated system,this results in added weight to take account of the speedbias, and for a precessed system, the effective torqueaxis is displaced.

For attitude maintenance, however, a transfer systemis preferable because the gas-jet expulsion system isessentially an on-off device subject to limit cycles.Although one type of control computer (discussed ina succeeding paragraph) has been studied that is quiteeconomical with respect to gas and electrical power, themomentum-transfer method of attitude maintenancebecomes competi tive for voyages of 60 days or morc.

A preliminary study has shown that gaseous nitrogenstored at a temperature of 40F and an initial pressureof 300 psia has many desirable properties for themomentum-removal portion of an attitude-control systemfor coasting flight.

2. Low-energy digital control. Th e weight of an

attitude-control system increases as the average magni-tude of th e angular rate of the vehicle becomes largerbecause greater changes in the angular momentum of th ehull must be effected. For this reason, it is desirable tolimit angular rates to low values whenever possible. Thenecessity of limiting bo th th e power level and energy con-sumption and the difficulty of measuring small rates ad dto th e problem of attitude control during th e major

portion of the voyage, when t he primary task of t he systemis to keep the vehicle pointing at a specified reference.

Table 2. Attitude-Control Accuracv Requirements

Operation

Mid-course maneuver

Satellite measurementsSatellite observatory

MeteorologicalDeep-space measurements

Photographs of planetsNavigation sightings"

Corn mu nica tion20-db-gain parabolic antenna36-db-gain parabolic antenna

Solar power

Position Accuracydesi

C 0 . 2 5 to 2

C 0.5

k 1 . 0

20.1 to 1.00.005 lo 1.0

& 8C 1

* l o to 2 0

Rate Accuracydeg/sec

kO.001 to 10.0

UA d d hi h hi l i d f d l i l i i

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A method of attitude maintenance that diminishes thepower level and weight of the system has been studied.The system requires neither rate information nor 3-bitdigital position information to maintain a slow, stablelimit cycle or to follow slow changes in the at titude-reference direction. The operation of this digital attitude-maintenance system is described below, with the aid ofFigs. 32 through 35. Figure 32 is a phase-plane diagramof angular position of the vehicle

about a single axis, with the origin at zero rate and zerodisplacement from attitude-reference direction.

vs angular rate

I

Fig. 32. Phase-Plane Diagram Describing Digi tal

Attitude-Maintenance Operation

The operation of the system may be described as a slowlimit cycle between operating limits ++,, Fig. 32) ofangular displacement, with an occasional excursion to alarger safety limit t 8 as a result of disturbances, systemunbalance, or change in the reference direction. Theperiod of th e limit cycle is kept long by making calibratedvelocity changes of magnitude ,8 at t he operating limits,or of magnitude @ at the safety limits, where 0 < E < 1.These changes are only large enough to provide a reversalof motion and a small drift velocity in the oppositedirection. This method of operation requires that the

initial conditions be reduced to an angular displacementfrom the reference smaller than the safety limit, and arate error less than ,8 in magnitude.

The details of operation when no disturbances arepresent ar e described with the help of the phase diagramin Fig 33 Assume that the point (the space probe SP)

-4s -40 40 4s

i-I----t--t-PIFig. 33. Attitude-Maintenance Operation

Without Disturbances

representing the state of the vehicle is initially at A .Then, if no disturbing torques are present, the velocity

remains constant and the SP moves to B . Here, a changein velocity of -,8 is introduced, moving the SP to C.Because the velocity at A is less than ,8. the velocity atC is of the opposite sign. It again remains constant, andthe SP moves to D. Here, a change in velocity of ,8 isforced, moving the SP to E . The velocity is now the sameas it was originally, and the SP moves through A, th ecycle repeating itself.

If disturbing torques a re present, the safety limits maybe reached. The operation is illustrated in Fig. 34. Assumethat the SP is initially at point E . If a constant disturbingtorque is present, the SP will move to F . Here, a changein velocity of -,8 is forced, moving th e SP to G. Becausethe initial velocity was greater than ,8 in magnitude, noreversal of sign in velocity occurs, and the SP moves pastthe operating limit toward H at the safety limit. Provided

I

Fig. 34. Attitude-Maintenance OperationWith Disturbances

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that a certain relationship betwceii +",+,y, and the disturb-ing torque is satisfied, the velocity at H is less than KP,where 0 < K < 1. At the safety limit, a velocity changeof - P is forced, moving the SP to J, where K < r< .Hence , the velocity changes sign, and the SP moves toward-4" . Again, provided that a certain relationship issatisfied, the velocity remains negative, so that the SPmoves to K where a change of P is forced, and the SPmoves to L For sufficiently small disturbances, the magni-

tude of the velocity does not increase beyond ,L?before +is reached, and at M a velocity change of - P moves theSP to N . The same events which occurred after the SPreached J are then repeated, but it can be seen that theentire phase-plane path has been displaced in the positivevelocity direction. Eventually, a condition similar to theprocess from E to F occurs, and the velocity at +,, isgreater than ,8. Hence, the safety-limit forcing has the

effect of resetting the normal operating cycle.

If the disturbing torques are greater than a certainmagnitude, it is possible to establish a limit cycle operat-ing from - p / 2 to p /2 , wherein the SP never reaches-+,. Other more complicated behavior is also possible,in which all of the forced velocity changes are of thesame sign. The saturating effects of this behavior on amomentum-transfer type of system must be consideredif such a system is contemplated , and provision made forthe removal of such stored angular momentum with amass-expulsion device.

Figure 35 is useful in discussing a possible mechaniza-tion for the attitude-maintenance equipment. It isassumed that an optical sensing device is used, and that

the celestial body toward which the body must be pointedis illuminated and sufficiently distant so that its angulardiameter is smaller than the distance between either +"or +8 - +",whichever is smaller.

I

Fig. 35. A Possible Mechanization for

Attitude-Maintenance Equipment

The discussion in the preceding paragraphs indicates

that a corrective action must be taken whenever themagnitude of the angular displacement becomes greaterthan or c+~. Also, negative displacements require dif-ferent action than positive displacements. Hence, theattitude sensor need only be a device that changes statewhen any of the following events occurs: (1) displace-ment is larger in magnitude than +",(2 ) displacement islarger in magnitude than +s, and (3) displacement is

positive.

The various possible states may be represented by a%bit word, as illustrated in Fig. 35, where conditionsC, , nd A are represented by t he left, middle, and rightdigits, each of which is zero when the condition listedabove occurs. Figure 36 illustrates a possible sensingdevice. The rays of light from the celestial body arefocused into a narrow elliptical spot, with the major axisparallel to the axis about which control is being main-tained. Strips of photosensitive material are applied tothe plate in the manner shown, and each strip is used todrive a relay or other switching device. Figure 37 shows

YPICAL LIGHT-BEAM SHAPE

Fig. 36. A Possible Optical Sensing Device forAttitude-Maintenance System

+TO SOURCE HA LF - SILVERED

L

L

L E N S

Y- A X I S P L AT E

Fig. 37. A Possible Optical Arrangement for Pointing theVehicle at a Celestial Body

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Correcting

Velocityft/sec

420

210

32

270

a3

a possible optical arrangement for pointing the vehicle

at a celestial body, a task requiring two-axis control.

Rocket Propellantan d Chamber

% pay load(Isp 265)

7.2

3.6

0.55

4.5

1.4

For proper system operation, once the displacementhas increased in magnitude beyond +"or +s, the velocitychange commanded by such a state change should takeplace only once during the interval between the occur-rence of the state change and the time when the displace-ment is again less than +o in magnitude. This may be

accomplished by using additional relays and contacts onrelay A to form lockout circuits.

Typical TargetSpread Without

CorrectionMissionmiles

E. Space Navigation and Terminal Guidance

Distance fromTarget when

Correction Appli edmiles

For futu re long-range space missions, the injection con-ditions cannot be controlled to sufficient accuracy. Subse-quent correction of the coasting trajectories, referred toas mid-course or terminal guidance, will be necessary.Terminal guidance is also taken to include brakingmaneuvers to achieve, for example, a soft landing on themoon.

The operation of most systems of trajectory correctionis based upon linear perturbation theory, i.e., the assump-tion that an actual trajectory differs only slightly fromsome previously computed standard trajectory. From thattheory it can be shown that, to correct an actual trajectoryto hit a moving point in space at a given time, the threecomponents of t he correcting velocity applied to thevehicle are calculable in terms of linear combinations ofthe perturbations (from the standard trajectory) in posi-tion and velocity. However, the desired end condition isusually to hit the moving point irrespective of time. In

that case, there is a degree of freedom in the correctingvelocity vector; e.g., the direction of the vector can bechosen to minimize its magnitude. Apart from economiz-

ing on rocket propellant, this also makes any miss (at

the destination) a second-order function of angular errorsin orienting the thrust vector.

It is implied that the actual trajectory differs from thestandard only because of injection errors; in other wordsthere is no analog of mid-course disturbances, such aswind, as in th e case of short-range surface-to-surfacemissiles.

1. Guidance systems. In considering schemes of guid-ance, closed-loop systems (such as those used for theguidance of homing missiles) have been ruled out. Con-tinuous correction of the trajectory would be too expen-sive in terms of rocket propellant for space missions. Thethree systems seriously considered would correct thetrajectory by one or two discrete-impulse, open-loop cor-

rections. The schemes are characterized b y the methodof measurement: (1 ) radio measurements from the earth,(2) optical sightings from the vehicle, and ( 3 ) angular-rate measurements of th e target body relative to thevehicle. The power requirements for active radio-echoguidance are prohibitive, and the flight times are too longfor all-inertial guidance. Methods (1) and (2) would prob-ably constitute mid-course guidance; in order that the

angular rates be measurable, method (3) would, in prac-tice, be used for terminal guidance.

It is desirable that any trajectory be corrected as earlyin flight as possible; the effect of the weight of rocketpropellant is thus minimized, but , unfortunately, the earlymeasurements tend to be more exacting. Table 3 showstypical da ta for the correction of various trajectories.

For a ground-based radio tracking and commandsystem, the critical measurements would be the azimuth

Table 3. Typical Propulsion Data for Trajectory Correction

-r-1,Ooo 10 x loJ20 x lo312 x 10'

k 400,000 30 X lo"

Venus t 50,000 3 0 X lo"

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and elevation angles. To keep the final dispersion at thc

moon down to 2 5 to 50 miles would require angularaccuracies of t .1 mil.

To determine position by optical measurements fromthe vehicle (space navigation), sightings would berequired on the moon, the earth, and two stars, one ofwhich might be the sun. For an interplanetary mission,two planets would be sighted relative to their star back-ground. For a final dispersion at the moon of _ t 25 to 50miles, mid-course sightings should be accurate to about0.07 mil; a dispersion at Mars of 20,000 to 40,000 mileswould correspond to sighting errors of about 0.1 mil at30 million miles from Mars. Of course, subsequent correc-tions may be made much nearer Mars.

Although space navigation by optical sightings seemspromising for some interplanetary missions, it does notappear feasible for lunar missions since the sightingswould be on bodies of non-uniform brightness (or onlypartially illuminated) subtending large angles. Forexample, at 120,000 miles the earth subtends 67 mils;bearing in mind th at it may be only partially illuminated,it would hardly be practicable to locate the center to 0.1%,the required accuracy. For this reason, terminal guidanceusing angular-rate measurements has been studied.

Such a system woiild have an optical sensor fixed inthe vehicle, and the vehicle, or perhaps just the guidancepackage, would be rotated to keep the optical axis point-ing a t the center of the target. The correcting-velocityvector is calculable in terms of t he range (from theangular diameter) and from the measured angular ratesof the vehicle.

2. P ost-injection radio guidance.A system of early mid-course guidance based on radio measurements and com-mand from the ground is attractive because vehicleequipment would be simple with most of th e complica-tions on the g round. However, the key factor in assessingthe feasibility of such a system is whether t he availableaccuracies of measurement meet the required accuracies.

In order to make a trajectory correction the six coordi-

nates of position and velocity must be known, but theyneed not be measured as such. On a calculable ballistictrajectory with no unpredictable disturbances i t is merelynecessary to compute the appropriate number of bound-ary conditions for the second-order differential equationwhich determines th e motion of the vehicle in space.Thus, in considering the radio guidance of a vehicle some

three days after injcction into an interplanetary orbit, it

would be feasible to measure only range rate, elevationand azimuth. The measurement of range may require awide-bandwidth radio link (or a complicated procedurefor integrating range rate) and t he angular rates wouldbe inconveniently small for measurement, e.g., 0.05 micro-rad/sec. On the other hand, range rate is convenientlymeasured by radio-doppler techniques with a transponderin the vehicle. Some numerical analysis was therefore

carried out to determine how accurately it would benecessary to measure range rate, elevation, and azimuthfor guidance with only these quantities.

Taking the example of a 165-day Mars trajectory, witha trajectory correction 1 million miles from the earth, themiss at Mars due to measurement errors was computed:

Range rate ............................................ 4,000 miles/ft/secElevation angle .................................... 20,000 miles/millirad

Azimuth angle .................................... 30,000 miles/millirad

In measuring range rate by a doppler system, the limit-ing accuracy would probably be a steady drif t of th ereference frequency source on the ground during the10-sec interval th at t he signal takes to cover twicethe 1-million-mile distance. Even so, a frequency drift of1 in lo5in 10 sec-achievable in practice-would be suf-ficient to make insignificant t he effect of range-ra te errors.

The accuracy of the JPL Goldstone 85-ft-diameterantenna is currently reported to be about +0.3 millirad;ultimately it will be about 2 0 . 1 millirad by Faking advan-tage of the world tracking net.

The real-time tracking procedure, described in Section111-H of this Repor t, is capable of fitting the trackingdata to give a smoothed estimate of either angle with astandard deviation of less than 0.1 times the uncertaintyin any single-position measure. However, uncertainties i nthe systematic errors (boresight errors) may make thisdeviation somewhat greater .

Using an optimistic figure of +-0.1 millirad for thesmoothed angular errors, the miss at Mars would be

4000 miles. This figure is considerably less than the error(+50,000 miles) due to the uncertain ty in the astronom-ical unit, to which earth-bound guidance systems aresusceptible. The accuracy corresponding to 0.1 milliradin the angles could not, therefore, be fully realized untilthe astronomical unit were known more accurately, pre-sumably from a previous space-probe experiment or fro

Documents

Exploration of the Moon, The Planets, And Interplanetary Space 1959