The Interplanetary Pioneers. Volume 2 System Design and Development

8/7/2019 The Interplanetary Pioneers. Volume 2 System Design and Development

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SPACECRAFT SUBSYSTEMS

in figure 4-16 reflects the complexity added by items (3), (4), and (5); thepower-conditioning, power-distribution, and protective equipment.

The Electric Power Subsystem Interfaces

103



The Electric Power Subsystem Interfaces

The preceding sections have dealt almost exclusively with electromag-netic and information interfaces-those associated with the brain andnervous system of the Pioneer man-machine system.In a biological analogy,

the power subsystem must represent the heart and blood vessels of thespacecraft. Subsystems are completely dependent upon electrical power todo things, even the pyrotechnic stored-energy devices that effect boomdeployment are detonated electrically. To illustrate the close relationshipbetween action and power on Pioneer, one has only to examine the STLfeasibility study and proposal. Early in the program, the CDU was con-

sidered part of the power subsystem rather than the command subsystem,because of the relationships between commands, energy, and physicalaction. The CDU was later consigned to the command subsystem becausein reality it is a complicated control valve that permits pulses of power toflow to command-selected spacecraft equipment. The pulses, in turn,operate switches and fire ordnance. Thus, pulses of power animate thes?acecraft while the steady bus power keeps the vital functions going.

Other interfaces are more straightforward. Because the power subsystemmust sustain the spacecraft electrical load continuously and cannot dependupon the battery fo r anything but short bursts of power, the solar arraymust be kept directed toward the Sun as accurately as possible. Thus, thepower subsystem imposes on the orientation subsystem the requirementthat the spin axis be perpendicular to the Sun line within 20. The shadowingor solid-angle interface with the spacecraft booms is the cause for the solar-cell-viewing band or bellyband around the girth of the cylindrical portion

THE INTERPLANETARY PIONEERS

time. Glass covers are applied to reduce this effect. Figure 4-17 illustratestwo of these considerations:

(1) If the spacecraft ventures closer than 0.8 AU to the Sun, the solararray becomes "voltage-limited." Increases in solar power are more thanff t b lt l d h i O d f 0 8 AU h

104



offset by voltage losses due to overheating. Outward from 0.8 AU, thepower subsystem is power-limited by the dwindling solar flux.

(2) The predicted useful power generated drops about 10 W between6 months and 3 years due to radiation damage of the solar cells.

Pioneer 9, an inward Pioneer, could not operate at 1.2 AU due to increasedloads over the nominal Pioneer. The increases in electrical load came pri-marily from the instruments and the convolutional coder.

These interface forces obviously play a leading role in power subsystemdesign.

The Design Approach

Flexibility and reliability were two critical design goals. Flexibilityapplied not only to the spacecraft's capacity to handle various scientificpayloads, but also the ability to operate between 0.8 and 1.2 AU withoutthe basic spacecraft design's being altered. The problem of the scientificinstruments' differing from spacecraft to spacecraft was handled by theprovision of a convenient bus voltage and placement of the burden ofmaking further modifications upon experiment power converters.

By design, the bus voltage varied with distance from the Sun (fig. 4-17).The entire Pioneer power subsystem "floated" at a voltage determinedby the solar-cell temperature. The spacecraft was overpowered intentionallyon inward missions. The battery was provided with taps at lower voltagesfor use during the inward missions. Enough solar cells were added to thenominal spacecraft that it could operate at 1 2 AU; thus there were too

SPACECRAFT SUBSYSTEMS 105

Voltage limited -- Power limited



Array capability: 6-month mis

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these factors in mind, STL computed the reliability for the entire powersubsystem to be a very high 0.9963for a 6-month lifetime (fig. 3-1).

Pioneer Power Budgets

106



Power requirements changed slightly from mission to mission. Thelargest change took place between the Pioneer 8 and 9 missions, when theconvolutional coder was added and the Goddard magnetometer was

replaced by one from Ames. These changes are summarized in table 4-15;these are, of course, average power levels, and the switching among themany spacecraft and experiment modes always created a varying powerprofile (fig. 4-15).

The Solar Array

The Pioneer solar cell is a high efficiency, solderless, n-on-p type, with1 to 3 ohm-cm base resistivity. Each cell is 1X2 cm and is covered by a0.15-mm glass slide for radiation protection. Early in the program, theaverage cell efficiencytarget was 12 percent; this was never achieved andthe cells on the spacecraft averaged about 10.5 percent. Both suppliers,

RCA and Texas Instruments, had considerable difficulty manufacturingcells to the demanding Pioneer specifications.

The individual cells were fabricated into two types of modules. In thefirst type, 12 cells were interconnected so that 3 were in series and 4 inparallel; in the second, there were 6 in series and 4 in parallel (figs. 4-18and 4-19). A close look at figure 4-19 seems to show the cells "shingled"

together along the long edges according to conventional practice. Actuallyeach cell is soldered to metal connectors; this makes the modules bothself-supporting and flexible. It was this flexibility that allowed the modules


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FIGURE4-18.-Pioneer solar array output characteristics.

was thus eliminated. The self-supporting property did away with the usuald l b t t ( d i the STL proposal) d i g l ll

108 T H E I N T E R P L A N E TA RY P I O N E E R S



FIGURE 4-19.—One of the Pioneer solar panels, showing both 12- and 24-cell modulesmounted on a curved substrate. (Courtesy of TRW Systems.)

spacec raf t equ ipm en t an d expe r im en t s . T h e bus vo l t age " f loa t s " a t t heso la r-ce l l a r ra y vo l tage . N A SA spec i f ica t ions res t r ic ted the vo l ta ge sw ingof this bus to 28 ±4 vol ts for an y loa d b etw ee n 15.3 a n d 5 5.6 W in int er

l t b t 0 8 d 1 2 AU th i l l i f t i f 6


TABLE 4-16.-Solar Array Capability

Type of cell---------------------------------- n/pEfficiencyofbare cell-- 10.7 percent min.Cell internal resistance------------------------ 0.46 ohm/cellArrayconfiguration 48 strings in parallel (each

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Arrayconfiguration-------------------------- 48 strings in parallel (eachstring: 4 parallel, 54 seriescells;total cells/array: 10 368)

Assemblylosses-- 1---------------------------percentGlasslosses--------------------------------- 5.4 percentLossesdue to proton damage - ------------------ 0Diodelosses-------------------------------- 1.0 V max.Temperatures ------------------------------- + 116 ° F at 0.8 AU

+52 ° F at 1.0 AU+4 ° F at 1.2 AU

RangeCalc min.net power 0.8 AU 1.0 AU 1.2 AU

Bus volts---- 24 25 26 24 26 27 29 27 28 29 30Watts------ 103 89.5 63 79.6 81.4 79.7 62.3 59.7 60.5 60.6 57.3

Typical peak loads include instrument-power peaks, fault clearing, coaxial-switch operation, and pneumatic-valve operation. The battery is eventuallydisconnected when its age begins to make it a poor risk, sometimes as lateas 18 months after launch.

Originally, a non-rechargable battery was proposed for the spacecraft,but a study of the MIT plasma probe power requirements showed theneed for a rechargeable battery that could meet the experiments' peakdemands.

The battery finally chosen for Pioneer was of the sealed silver zinc type

THE INTERPLANETARY PIONEERS.

The Converters

Three classesof power converters transform bus power into usable formfor: (1) the TWTs, (2) the scientific instruments, and (3 ) the rest of thespacecraft equipment.

110



Each TWT has its individual converter. The TWT converters are similarto those used for the spacecraft equipment except for special output voltages,including 1000 V for the TWTs. The 1000-V line also must be regulated

ratherprecisely: -0.5

percent over the 28+ 5

voltage swing of the bus bar.The TWT converters may be commanded on and off separately; but theswitching logic is such that the TWTs cannot operate simultaneously.Furthermore, if the bus voltage falls below 23.5 V for more than 0.4 sec,an undervoltage command automatically shuts the TWTs off to precludedefocusing them or burning them out. Removal of the TWT load of ap-

proximately 30 W causes an immediate rise in bus voltage under normaloperating conditions. Minimum efficiency specified for the TWT con-verters was 80 percent. Each converter weighs 2.35 lb and has a volume of64.12 in.3 .

As mentioned earlier, each scientific instrument possessesits own con-verter tied directly to the bus. All experiments are turned off simultaneously

via a ground command or an undervoltage condition, but they may beturned on one at a time. Simultaneous turn-off permits quick diversionof power to spacecraft equipment in the event of an emergency. Theinstruments are also turned off by the same automatic undervoltage com-mand that shuts down the TWTs.

The two equipment converters are packaged together, weigh 3.2 lb, andoccupy a volume of 111.3 in.3 . The units are identical, but their outputsare partially cross-strapped. Converter no. 1 supplies receiver no. 1 anddecoder no. 1 , while converter no. 2 provides power for receiver no 2 and


in the CDU (fig. 4-14). The power-distribution portion of the CDU isillustrated in figure 4-20. Tables 4-13 and 4-14 show a large number ofon-off commands that are really power-on/power-off commands thatconnect or remove components from the power source. With command107, the mission controller can override the automatic undervoltage

111



switch that disconnects the TWTs and experiments. Command 000,labelled "undervoltage simulate," is used if the TWTs and experimentsmust be disconnected all at once. Command 000 is then OR-gated withthe undervoltage signal (which has not yet disabled the TWTs and ex-periments because the voltage is still within bounds), and the TWTs andexperiments are turned off. Command 107 may also be employed to lockout command 000.

The undervoltage control senses the bus voltage from a voltage dividerwith a half-volt resolution. The trip point is adjustable and is usually set

for 23.5 V. To prevent the inadvertent shutdown of the TWTs andexperiments due to transients, the undervoltage control has a half-secondtime constant. Like all electronic circuits, the undervoltage control isfallible and might fail in a way that would shut down the spacecraft. Theundervoltage override command was introduced primarily to prevent suchan occurrence.

A number of current and voltage monitors report the operational con-dition of the electric-power subsystem to the mission controller back onEarth. These are listed in tables 4-9 and 4-10 with the other housekeepingtelemetry words.

THE ORIENTATION SUBSYSTEM

Only a small, spin-stabilized spacecraft could meet the cost, reliability,and launch-vehicle constraints of the Pioneer. For maximum utility, the

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adjust spacecraft orientation if the axis drifted out of the 9004-2° attitudewith respect to the plane of the ecliptic.

The most important components needed in such an orientation maneuverare: (1) a device to torque the angular momentum vector of the space-craft, (2) sensors to control the direction of axis motion, (3 ) sensors to

113



signal the status and, hopefully, the success of the orientation maneuver,and (4) a nutation wobble damper to dissipate nutation energy inducedduring the orientation. The small solar sail added at the tip of the high-gain

antenna mast to offsetany residual torque due to solar pressure' wasnot part of the original design.The components will becovered in more detail later; first the orientation

concept will be sketched out completely. After the spacecraft is injectedinto the plane of the ecliptic, two pairs of Sun sensors determine the atti-tude of the spacecraft with respect to a line joining Sun and spacecraft.

The Type-I orientation maneuver commences automatically. The Sunsensorscause the nitrogen gas jet to fire and torque the spacecraft spinaxis through the smallest angle until it is perpendicular (within 40.5percent) to the spacecraft-Sun line.'3 At this point, thermal control ispossible and the solar array generates full power. The Type-II orientationis commanded from the ground and is controlled by monitoring the

strength of the spacecraft transmitter's signal strength. When it is maxi-mized, the Pioneer spin axis is also perpendicular to the spacecraft-Earthline; the desired accuracy is -1.0 percent. If the spacecraft is perpendicu-lar to both the spacecraft-Sun and spacecraft-Earth lines, it is also ap-proximately perpendicular to the plane of the ecliptic. Orientation is nowcomplete. Spin-axis orientation is maintained through spin stabilization atroughly 60 rpm (ref. 4).

Pioneer Specification A-6669 stipulated the performance of the orienta-tion subsystem more precisely:


The Sun Sensors

The sensitive elements of the Sun sensorswere quad-redundant, photo-sensitive silicon-controlled rectifier (PSCR) chips, manufactured by SolidState Products, Inc. The chips were developed especially for the Pioneer

114



Program. They delivered a signal to the orientation-control circuitrywhenever the Sun was in view. However, the view of each Sun sensor wasrestricted by aluminum shades (fig. 4-21). On Pioneers 6 and 7 the light-

sensitive chips were protected against space radiation damage by 20-milquartz covers. Several months after launch, however, it was discoveredthat the Sun-sensor thresholds had changed. Laboratory testing impliedthat radiation damage was the primary cause, and the quartz covers onPioneer 8 were made 100 mils thick. The trouble persisted. The realproblem was discovered inadvertently at TRW Systems when the sensors

were tested under ultraviolet light to see if it degraded the adhesives usedin sensor construction. During these tests, it was discovered that the sensorswere ultraviolet-sensitive. In space, the ultraviolet light from the Sun hadcaused the change in the sensor thresholds. Simple ultraviolet filters wereadded to 60-mil quartz covers; this cured the situation on Pioneers 9 and E.

The five Pioneer Sun sensors are mounted on the spacecraft with the

fields of view specified in figure 4-22. Sensors A and C, located on thespacecraft bellyband, looking up and down respectively, help position thespacecraft during the Type-I orientation. As long as the spin axis does notpoint within 1 0 0 of the Sun, except for a small overlap of the field of view,sensorsA and C will see the Sun once each revolution as the spacecraftspins. The Type-I orientation proceeds as sensor A or C, whichever one isilluminated, stimulates a succession of gas pulses from the jet on the endof the orientation boom. Each pulse lasts for 450 of spacecraft rotation andtorques the spin axis around about 0 150 in the direction of the smallest

S PA C E C R A F T S U B SY ST EM S 11 5




Spacecraftrotation

116



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Sensor E establishes the reference position of the spacecraft with respectto the Sun and sends signals to the scientific experiments. Sensor E is also

mounted on the viewing band of the spacecraft. It possessesonly a 20 fieldof view that provides short, sharp pulses,as it seesthe Sun roughly once eachsecond. Because the field of view is only 40 ° ' in the other direction (fig.

l i f

117



4-22), Sun pulses appear only when the spin axis is within 200 of beingperpendicular to the spacecraft-Sun line. The appearance of Sun pulsesalso indicates that the Type-I orientation is proceeding successfully and

near its end.The Pneumatics Assembly

Short bursts of cold nitrogen gas from the pneumatics assembly changethe spacecraft angular-momentum vector. Gyroscopes, hot-gas jets, smallpyrotechnic devices, and miniature rockets have all been used on Earth

satellites for purposes of attitude control. The cold-gas system chosen forPioneer is simple and extremely reliable. It had already been well provenon other space missionswhen Pioneer was being designed.

The pneumatic assembly is a titanium alloy pressure vessel containingabout 0.9 lb of nitrogen at 3250 psi (fig. 4-23), a pressure regulator, asolenoid valve, a pressure switch, and a nozzle. The nitrogen had to be verydry to preclude the valve's icing at low temperatures. An electrical signalopens the solenoid valve fo r a moment, releasing a burst of gas at about50 psi which provides the desired impulse. The solenoid valve and nozzleare located on the end of a 62-in. boom to increase the angular impulseand isolate the iron core in the valve solenoid from the magnetometer(fig. 4-24). Originally the nozzle was on the tip of the high-gain-antennamast, but it was displaced by the magnetometer during the evolution ofthe spacecraft. Finally, both were placed on booms.

THE INTERPLANETARY PIONEERS118



FIGURE4-23.-Components and block diagram of Pioneer pneumatic equipment.

SPACECRAFT StIBSYSTEMS 119



FIGURE 4—24.—The Pioneer or ientat ion boom, shown on Pioneer C during thermal-va cu um test . (Spacecraf t is upside d ow n.)

t e r r e s t r i a l command s t a r t s t he cha in o f even t s l e ad ing up to t he gas pu l se .A c o m m a n d f ro m E a r t h is n e e d e d fo r e a c h Ty p e - I I g a s p u l s e ; t h e r e m a yb e h u n d r e d s d u r i n g a c o m p l e t e m a n e u v e r.

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pointing error will equal the peak amplitude of the wobble. Even with asevere wobble, this degree of orientation achieved will probably be sufficientfor thermal control and nearly full power production. The spacecraft willbe self-sustaining and not dependent upon the battery. Thus there will betime for a wobble damper to suppress the residual wobble. After this occurs,the orientation electronics can be turned on again, automatically initiating

121



another Type-I orientation to trim the spacecraft attitude more finely. Infact, the accuracy of the orientation can be checked by noting whether thegas jet fires upon the initiation of a Type-I maneuver. If nothing happensor if only one pulse is detected, the spacecraft is oriented precisely enoughfor sensors A and C both to see the Sun.

Most wobble dampers in use on satellites and other spacecraft removewobble energy by dissipating it as friction-generated heat. On the Pioneerspacecraft, the energy of nutation was dissipated by beryllium-copper ballsrolling inside and impacting at the ends of a pair of tubes located at the endof the 62-in. boom. Rolling friction and inelastic collisions at the ends ofthe tubes extracted the energy of nutation, converting it to heat. Originally,the tubes were to be filled with gas to provide hydrodynamic friction, butit was found that the gas was unnecessary. The damper was built by STL.

Weight, Reliability, and Power Drain

The entire orientation subsystem weighs only 6.5 lb, including about0.9 lb of nitrogen. This figure includes almost completely redundant parts(with voting circuits) in the electronics and Sun sensor assemblies. Thepneumatic equipment is not redundant, although this was seriously con-

sidered early in the program. Even so, the reliability of the entire subsystemwas calculated as 0.980 for a 6-month lifetime in space. When the orienta-tion subsystem is in a standby mode (as it is most of the time), it consumes


or situations that occurred before the spacecraft broke into full sunlightfollowing launch:

(1) The launch-pad environment-the spacecraft-determined air-conditioning requirements had to be examined.

(2) Aerodynamic heating of the shroud during launch and the con-sequent transfer of heat to the spacecraft-this was controlled by adding

122



thermal insulation (chargeable to spacecraft payload) to the shroud inquantities dependent upon the specifictrajectory selected (ch. 7).

(3 ) Aerodynamic heating of the spacecraft at very high altitudes aftershroud ejection-analysis showed that no problem existed here.

(4) Radiant heating of the bottom of the spacecraft by the third-stagerocket plume-the switch to the X-258 third stage, which used aluminumoxide additives in the rocket grain, stimulated concern over excess radi-ation; a special STL study determined that a thermal shield was needed toblock the solar

array's view of the plume.'4

(5) Cooling during eclipse of the Sun by the Earth during ascent-thisperiod, which would last at the most 30 min, would not be long enough toallow the spacecraft to cool excessively. (Actually, the dark side of theEarth contributes considerable thermal radiation to the spacecraft duringeclipse.)

In summary, analysis of the transient events from launch pad to solar orbitresulted only in the addition of a radiation shield fo r the thermal louveractuators and varying amounts of thermal insulation to the shroud. Thelong cruise around the Sun controlled the major aspects of spacecraftthermal design.

So far, only the thermal control of the spacecraft interior has beenmentioned. The solar cells, Sun sensors,antenna mast, and booms must bemaintained within operating limits, too. Because they are spacecraftextremities the thermal control techniques applying to the interior of


internal heat load by 50 percent at these extreme points in the missionspectrum. Active thermal control was the best solution, even though theaddition of moving parts would detract from overall spacecraft reliability.

The whole Pioneer mission concept depended upon the concept of aspin-stabilized spacecraft with a spin axis normal to the plane of theecliptic. The curved sides of the cylinder receive essentially all solar radi-

123



ation, while the ends point toward cold space. This situation is ideal for athermally insulated spacecraft with active thermal control. Insulationaround the sidesof the structure allows only a small portion of the solar heatload to reach the inside of the spacecraft. Insulation on the top leaves thebottom as the only possible exit for heat (fig. 4-26). This heat leakage,which varies depending on the distance from the Sun, can be radiated outthe spacecraft bottom along with the variable internally generated heatload. The variability is handled by changing the effective radiating area

of the bottom of the spacecraft. Mechanization of the concept consistsof a set of Venetian-blind-like louvers that varies the effective radiatingarea, increasing it as the internal temperature rises and reducing it whenthe inside of the spacecraft becomes too cool (fig. 4-27). The setting of thelouvers is controlled by bimetallic thermal actuators sensitive to theinternal temperature. When the Pioneer program began, STL was also

applying this basic concept to the OGO, which, though much larger thanPioneer, was fully stabilized in orbit and had many of the same thermalproblems. The louvers used on Pioneer came directly from OGO tech-nology. (STL was also the OGO spacecraft prime contractor.) Other spaceprobes and stabilized Earth satellites have also used the same approachas Pioneer.

The thermal insulation covering the spacecraft sides and top thermallyisolate the antenna mast, the booms, the Sun sensors,and the solar arrayfrom the volume that is temperature-controlled by the louvers. As we shall



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FIGURE 4-2 7.— Bo ttom view of the Pioneer spacecraf t with the third-s tage mo tor inp lace . The thermal louvers , cover ing two- th i rds o f the equipment p la t form, a re shownin an open posi t ion.

t he bas i c spacec ra f t t hus p ro t e c t ed cou ld ven tu re be tw een 0 .55 and 2 .0 A Uw i t h o u t t h e r m a l c o n t r o l m o d i f i c a t i o n s , a l t h o u g h c o n t r a c t u a l l y t h e y w e r eco m m i t t ed t o on ly 0 .8 an d 1 .2 A U .

TH E INTERPLANETARYPIONEERS

TABLE 4-17.-Thermal-Sensor Locations

Thermal-sensored equipment

Receivers I and 2 -----------------TWTs I and 2...........TWT

Thermal-sensor location

On receivers, near voltage-controlled oscillatorAt juncture of mounting screw and platformE i f

126



TWT converter.....................Transmitter driver ------------Digital telemetry unit - ---

Data storage unitEquipment converters 1 and 2Battery - - - - - - - - - - ----.-.-.-.-.-.-Upper solar panel and

lower solar panel

Platform 2

Antenna mounting bracketLouver actuator housing -----------Sun sensor A --------------------Platform 1Nitrogen bottleSun sensor C ----Platform 3 -- - - - - - - - - - - - - - - -

Magnetometer sensor (Ames) 4 ------

Magnetometer, electronics (Ames) 4Plasma, electronics (Ames) 5 ------Cosmic ray (SCAS) 6 --------------Cosmic ray (Minn.) I

Exterior of converter top coverOn platform close to driverAt juncture of mounting screw and platform

At base of data storage unitAt base of equipment converterInternal to batteryApproximately 300 to right (top view) of

orientation boom. Between substrate andinsulation (insulated from compartment).

On platform position 2

Midway on bracket, within compartmentBetween insulation and housingIn head of sensorOn platform position IEpoxied directly to bottleIn head of sensorOn platform position 3

Internal to boom-mounted magnetometersensor

Internal to instrumentInternal to instrumentInternal to instrumentInternal to instrument

Because complex geometry and the manifold heat paths make spacecraftthermal analyses so difficult, it is customary to build a thermal mockup or


high heat impedances with fiberglass mountings to minimize heat leaks intothe interior. On the other hand, heat paths from internal components tothe bottom of the equipment platform are designed with high conductivityin mind. The spacecraft's instrument platform and the boxes mounted onit were made thermally "black" to encourage temperature equalization.Other "inside" surfaces, such as the top cover and spacecraft sides, wereeither aluminum or l i i d M l 'Th i t l tf i

127



either aluminum or aluminized Mylar. 'The equipment platform is analuminum honeycomb panel constructed with the "starved" bondingtechnique to insure good thermal conductance through it to the radiatingsurface on the bottom.

The louver system (fig. 4-27) consists of 20 individual louvers, eachactuated independently by a spiral-wound, bimetallic spring. Springs areinsulated so that they are responsive only to local temperatures. The openradiating area was approximately 3 ft2 . One-third of the platform area, theportion

directly under the magnetometer electronics, does not requirethermal louvers. Instead, it is covered with aluminized-Mylar insulation.The louver blades themselves were made highly reflective and specular toinfrared radiation to minimize radiation from them back to the equipmentplatform when they were in the full open position. They are also goodthermal insulators, so that when closed they help retain heat within the

spacecraft. The bottom of the equipment platform is the emitting surfacefor all waste heat.

Protection of the spacecraft from thermal-plume heating during injectionconsisted of applying aluminum foil around the top of the Delta third stageand aluminized-Mylar insulation around the interstage ring. The plumeheating, however, was not so severe as expected. 15

Controlling Extremity Temperatures


Thermal Control Subsystem Reliability

The thermal coatings, thermal insulation, and thermal conduction

paths in the Pioneer spacecraft present no reliability problems. The onlymoving parts are the individually actuated thermal louvers. Catastrophicfailure of several louvers in the neighborhood of a large source of thermalenergy is highly unlikely. In fact, the use of individual actuators for the

128



energy is highly unlikely. In fact, the use of individual actuators for thelouvers makes the probability of acceptable operation over a 6-monthlifetime very high, roughly 0.999.

THE STRUCTURE SUBSYSTEM

The structure subsystem, like the thermal control subsystem, is a largelypassive, but critical, subsystem. Spacecraft have rather complex structure

subsystems which must be analyzed as painstakingly as the communicationssubsystem or any other subsystem. The Pioneer structure (figs. 4-28 through4-30) consists of the following major sections:

(1) The interstage ring and cylinder(2) The equipment platform and struts(3) High-gain antenna mast supports

(4) Solar-array substrate and supports(5) Boom dampers(6) The booms and associated deployment and locking equipment(7) The Stanford experiment antenna

Overall Configuration

Spacecraft structure is highly variable. In orbit about the Earth areli d h b t t h d d th l h d S b


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(2) The booms provide magnetic isolation for the magnetometer andexile the orientation nozzle solenoid and the wobble damper, both of whichpose magnetic cleanliness problems.

(3) The effectivenesses of the orientation nozzle and wobble damper areincreased by placing them on the ends of booms.

The decision to add deployable booms to the spacecraft was most criticalf h d i B h b

132



from the structures standpoint. Booms are moving parts that must bestowed in a launch configuration and then unfolded and locked in position

after the launch vehicle fairing has been ettisoned. Other scientific satellites,such as OGO 1 , have been compromised by boom failures. In the case ofPioneer, the advantages of using booms far outweigh their potentialliability.

Externally, the Pioneers are cylinders 37.3 in. in diameter and 35.14 in .long, with three booms 1200 apart extending 82.44 in.from the spin axis

(fig. 4-28). The Stanford experiment antenna projects downward whendeployed, and in appearance and complexity is a fourth boom. The high-gain antenna mast projects roughly 53 in. above the top edge of the cylinder.Pioneer, therefore, presents appendages in all directions, in contrast to therelatively clean configuration first suggested by STL (ref. 5).

Internally, the major requirements were support fo r scientific instru-

mentation and spacecraft subsystems and, once again, spin-axis symmetry.Symmetry must be taken here to mean the judicious placement of massaround the spin axis to preclude the spacecraft's wobbling. The farthercomponents were located from the spin axis, the greater the spin stability;that is , the better the spinning spacecraft could resist destabilizing in-fluences. The internal configuration (fig. 4-29) follows general spacecraftpractice-the major structural element is a strong equipment platform. Thisplatform supports all internal components, the three radial booms, and the


from the aluminum interstage ring to another ring attached to the bottomof the equipment platform. The nitrogen pressure vessel supplying theorientation subsystem nests within this cylinder. Six struts link the platform

with the bottom of the thrust cylinder, providing additional rigidity. Theequipment platform supports the rest of the spacecraft (fig. 4-30). Threealuminum-mast struts absorb side loads transmitted by the antenna mast.The top cover, which is made of an aluminized-Mylar blanket, bears no

133



loads. The original cover was an aluminum sheet, but it was discarded tosave weight. The bottom of the "can" is really the equipment platform,although the solar-array skirt continues downward for another 20 in.

The equipment platform is an aluminum honeycomb sandwich, 0.75-in.thick with 0.016-in.aluminum face sheets.The material weighs 3.1 lb/ft 3 .It carries the loads transferred from interstage cylinder and struts to thebooms, antenna mast, solar array, etc. The thermal louvers are mountedon two-thirds of the lower surface of the platform. An insulation blanketcoversthe remaining one-third of the surface. On the top surface, the spacesbetween mounted equipment are covered with black paint.

Aluminum-honeycomb sandwich material was also used fo r the solar-array substrates. The inner and outer face sheetsare "prepreg" fiberglasssheets. The substrate panels are attached to the equipment platform by

fiberglass brackets around the lower ring of solar panels and through theSun sensor brackets around the upper ring. Support rings at the top andbottom ends complete this part of the structure. Upper panels are inter-changeable among themselves, as are those in the lower ring. The 6.75-in.band between upper and lower rings was left bare because the boomshadows would have rendered solar cells useless in that area anyway.

This band is closed thermally with an aluminized Mylar blanket. Thebooms are hinged in this "bellyband."

The 62-in. radial booms, which isolate the orientation nozzle, wobble


Other Structures Tests and Analysis

STL performed the stress analysis of the Pioneer spacecraft. The studiesinvolved static analysis and examination of such dynamic factors as balance,moments of inertia, rigidities, limitations of vibratory response, spacecraftspinup and separation, and attitude stability and damping. The dynamicsof appendage deployment were of particular concern because of pastdifficulties with booms. This concern led to a special test apparatus which

134



difficulties with booms. This concern led to a special test apparatus whichwas built to check deployment under close-to-actual conditions (see ch. 7).

Spin tests, vibration tests, and the other related tests described in chapter 7were extensive. They required the construction of a special "structuralmodel" of the spacecraft, wherein the major structures either duplicatedthose in the intended flight model or, in the case of electronic equipment,simulated them in weight.

Structural reliability analysis is not as advanced as it is for'electronics

equipment; nevertheless, some estimates can be made. STL calculated thatthe overall structure reliability would be 0.998 fo r launch, boost, injection,and free flight. This estimate was based upon tests performed upon cablecutters and deployable booms built for OGO and other space programs.Of course, such estimates based on moving parts assume that no failures ofstatic structural members occur. The useof factors of safety during the stress

analysis gives this assumption some foundation. Pioneer structural studiesassumed a yield factor of 1.35 and an ultimate safety factor of 1.50.

The possible effects of the space environment upon the spacecraft werealso analyzed carefully. The analysis showed:

(1) Solar heat flux was controlled by the louvers and thermal coatingsdiscussed previously under thermal control.

(2) Solar particulate radiation, which has a potential for degradingmaterial structures, was several orders of magnitude below damageh h ld

SPACECRAFT SUBSYSTEMS 135

TABLE 4-18.-Block-I and Block-II Spacecraft Weight Breakdowns

Equipment Block I I Block II b

Communication subsystem -------------------------- 14.35 lb 14.33 lbReceivers (2 ) ----------------------- 6.14 6.14Transmitter driver ------------------- ---- .- .- .-- 1.31 1.29TWTs (2 )------------------- 1.89 1.89Attenuators and supports ------------------------- 0.12 0.12



Branch line coupler ------------------------------ 0.24 0.25Diplexers (2 ) -----------------------------. 1.39 1.38

Bandpass filter ------------- -----.-.-.-.-.--.-.-- 0.25 0.25Coaxial switches (5 ).-..- --- 1.00 1.00High-gain antenna ---------- 2.01 2.01

Data handling subsystem ------------------ 10.65 10.78

Digital telemetry unit ---------------------------- 8.57 8.64Data storage unit ------------------------ 1.73 1.75Signal conditioner .-- - 0.35 0.39

Command subsystem ------------------------------ 10.72 11.47

Command decoder ------------------------------ 5.60 6.15Command distribution unit .. 5.12 5.32

Electric power subsystem ------------------ 36.54 38.03

Solar cells, substrate, glass, etc. . . . . . . . . . . . . . . . . . . . . . 13.98 13.41Support rings and brackets ----------------- 2.96 2.96

Battery -------------------------------- 2.19 3.16Equipment converter .---------- 3.02 2.99TWT converters (2 ) --------------- 4.52 4.49Cabling and connectors ------------------ 9.87 11.02

Orientation subsystem ------------------- 6.68 6.95

Nitrogen bottles and supports --------------------- 1.54 1.75Nitrogen gas --------------------------------- 0.87 0.93

Solenoid valve ---------------------------------- 0.44 0.40Regulator ------------------------------ 0.99 1.00Nozzle ---------------- _ ..... ..01 0.01

136 THE INTERPLANETARY PIONEERS

TABLE 4-18.-Block-I and Block-II Spacecraft Weight Breakdowns (Continued)

Equipment Block I Block II b

Magnetometer boom flange ----------------------- 0.05 0.07Hinge fitting structure --------------------------- 1.50 1.50Boom dampers (3 )-------------- 0.57 0.57Boom tie-down ------------------- - 0 79 0 79



Boom tie down 0.79 0.79Solar sail --------------------------------------- 0.06 0.06Hardware ------------------- 1.18 1.25Inertia weights (on boom) -------------------- 0.12 0.55Platform struts ---------------------------------- 1.08 1.08

Total spacecraft weight without experiments ----------- 103.20 106. 66

Experiments ---------------------------- 34.74 40.54

Magnetometer (Goddard/Ames) ------------ 5.81 7.74Cosmic ray detector (Chicago) -------------------- 4.71

Cosmic ray detector (GRCSW) -------------------- 4.39 5.55Plasma probe (MIT) --------------- 6.13Plasma probe (Ames) ---------------------------- 6.33 5.92Stanford radio propagation experiment I ----------- 7.37 7.01Cosmic dust (Goddard) -------------------------- 4.29Cosmic ray (Minnesota) ------------------------------ 7.91Electric field (TRW) ------------------ ---------- ---------- 0.87

Convolutional coder --------------------- 1.25Total spacecraft weight with experiments ....- ... 137.94 147.20 d

a Reported spacecraft weights vary slightly for each spacecraft depending upon thedata source. The weights shown in this column are for Pioneer 6 ; taken from "MonthlyInformal Technical Progress Report, Pioneer Spacecraft Program." Period 1 Decemberto 31 December 1965, TRW Systems Report 8400.3-247, January 10, 1966.

b For Pioneer 9. Taken from "Monthly Informal Technical Progress Report, PioneerSpacecraft Project." Period 1 November to 30 November 1968, TRW Systems Report,December 9, 1968.

Includes balance compensation weights of 1 33 and 063 lb respectively

CHAPTER 5

Scientific Instruments



SCIENTIFIC O B J E C T I V E S

H E PIONEERS are multidisciplinary spacecraft. From the scientificXstandpoint they are very closerelatives of the Interplanetary Monitoring

Platforms (IMPs) orbited around the Earth and Moon from 1963 on. Infact, many IMP experimenters are also Pioneer experimenters, and their

instruments are similar on both series of spacecraft. This is not surprising,as both types of spacecraft were designed to measure the same importantfacets of the interplanetary medium: the plasma, cosmic rays, magneticfields, cosmic dust, electric fields, and space propagation properties. TheIMPs, however, center on the Earth-Moon system,while the Pioneers areSun-centered.

The scientific objectives of the Pioneer spacecraft are to measure theabove-named facets of the interplanetary field. In 1962 virtually nothingwas known of what transpired in interplanetary space. In particular,Earth-bound scientists had little feel for how plasma, cosmic rays, etc.,varied spatially and in the time and energy dimensions. The Pioneerscientific objectives were sharpened in three ways:

(1) The spacecraft were launched at intervals that permitted the solarcycle to be covered from minimum to maximum. (The long lifetimes of

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APPLICATIONS O F PIONEER D ATA

In 1962 and 1963, the Pioneers would hardly have been called "appli-

cations" spacecraft, so firmly were they directed toward satisfying scien-tific curiosity. Solar events, however, have wide repercussions, jigglingmagnetometers on Earth, disrupting long-distance communication, and thelike. Pioneers 8 and 9, cruising well behind the Earth in its path aroundthe Sun, can radio warnings to the Earth of solar radiation storms which

13 9



the Sun, can radio warnings to the Earth of solar radiation storms whichwill soon catch up with the Earth. Since the Sun rotates once each 28 days

and drags its plasma and radiation streams around with it, the Pioneerslagging the Earth are well-situated to forecast interplanetary weather forthe Earth several days in advance. These data are now sent to the Environ-mental Science Services Administration, which then distributes them toabout a thousand users (see Vol. III).

Thus, Pioneer instrumentation has practical applications not foreseen at

the beginning of the program. On the later Pioneers, instrument selectionand design were affected to some extent by this new dimension of theprogram.

INSTRUMENT INTERFACES AND SPECIFICATIONS

In Pioneer terminology, the scientific instruments are considered aseparate system rather than a subsystem of the spacecraft. The forcesexerted on the scientific instruments are considered to be similar to thoseencountered by the spacecraft subsystems. The most important are themechanical loads imposed during launch, the heat from the Sun, the

magnetic and electromagnetic environments extending from the othersubsystems, and the information interface enforced by the data handlingsubsystem These spacecraft subsystems are defined in detail in chapter 4


interface specifications to reemphasize the extra dimensions involved ininstrument design for space research. The following considerations werenecessary:

(1) Mechanical interfaces-dimensions, weights, mounting orientationsand view angles

(2) Electrical interfaces-power levels, voltages, transients, connectorsand cabling

(3) I f ti i t f d l h bi f d i i

140



(3) Information interfaces-word lengths, bit rates, formats, and timingsignals

(4) Thermal interfaces-operating temperatures and surface coatings(5) Electromagnetic interfaces-interference, shielding and grounding

The factors mentioned above are discussed in chapter 4.In addition to matching spacecraft interfaces, each scientific instrument

had to mesh with the interfaces presented by ground-support equipment.

Before even reaching the launch pad, instruments had to be qualified(usually through prior flights on sounding rockets or satellites) and thentested according to the standards described in chapter 6.

A number of military specifications were also applied to the Pioneerspacecraft and its cargo of instruments. One of the most critical wasMIL-I-26600, "Interference Control Requirements, Aeronautical Equip-ment," highlighting the fact that the electromagnetic environment had tobe shared with other spacecraft as well as a host of other aerospace equip-ment at Cape Kennedy prior to and during launch. Finally, the presenceof radioactive sources for instrument calibration meant that federal andstate laws governing the use and transport of radioactive materials alsoapplied.

The scientists flying instruments on Pioneer (o r any spacecraft, fo r thatmatter) had to deal with managerial controls, with specifications more


ments." Experiments for Pioneers A and B were solicited by letter in early1963; for Pioneers C and D, in late 1963 and early 1964.

When experiment proposals have been received, they are evaluated atNASA Headquarters with the assistance of the Space Sciences SteeringCommittee. The members of the Committee and its several subcommitteesare appointed from scientists in NASA, other Government agencies, uni-versities, and non-profit organizations. In the case of Pioneer, the followingf b itt i l d A t S l Ph i I h

141



four subcommittees were involved: Astronomy, Solar Physics, Ionospheresand Radio Physics, and Particles and Fields. The Pioneer Project Officealso reviewed experiment proposals from the standpoints of engineeringfeasibility, cost, and compatibility with the spacecraft.

Usually NASA receives more proposals for experiments than the space-craft can carry. Therefore, the Space Sciences Steering Committee mustchoose those experiments that meet the minimum requirements and thenassign priorities. For example, 18 proposals were evaluated in depth forPioneers A and B, 15 for Pioneers C and D; but only 7 and 8 experimentflew on these two blocks of spacecraft, respectively. The general criteriaemployed in the selection process are: (1) scientific merit, (2 ) ability of theinstrument to make the desired measurement, (3) development status ofthe instrument, anid (4 ) understanding and experience of the experimenters.

The criteria specific to Pioneer that were employed are: (1) pertinence toPioneer mission, weight, data rate, power, etc.; and (2) pertinence withrespect to the solar' minimum.

The Summary Minutes of the meeting of the Space Sciences SteeringCommittee, dated July 22, 1963, typify the selection procedure. PioneerA and B experiment's were divided into two categories as follows:

(1 ) Firm payload,' including magnetic fields, plasma, cosmic-ray gra-dients, and radio propagation

(2 ) Tentative or backup experiments including cosmic ra anisotrop


were received; and ultimately the one submitted by Goddard Space FlightCenter was selected for the Pioneer C and D payloads.

When it was decided to combine the parts left over from Pioneers Athrough D and assemble Pioneer E, the question of experiment selectionwas revived. During the fall of 1965, however, NASA decided to retainthe Pioneer C and D payload rather than making extensive modificationsto the spacecraft parts on hand.

Some of the proposed experiments did not fall within the mission scope

142



Some of the proposed experiments did not fall within the mission scopesuggested by NASA. For example, a proposal submitted by Space/DefenseCorporation was aimed at investigating the influence of electromagneticand gravitational fields on diurnal rhythm. Interesting as such an experi-ment would have been, it would not have measured parameters relatedto the other investigations.

The experiments finally selected fo r the five Pioneer spacecraft arepresented in table 5-1.

THE GODDARD M A G N E T O M E T E R (PIONEERS 6, 7, AND 8)

The interplanetary magnetic field is created by the Sun and modulatedby the streams of plasma that spiral out into the space between the planets.Magnetic field measurements, particularly those that record transientsfollowing solar activity, are critical to our understanding of the spacesurrounding the Sun.

The spin-stabilized Pioneers permitted the use of a unique magnetometerdesign whereby all three components of the magnetic field could be meas-ured with a single-axis sensor (ref. 3). If the sensor axis is mounted at anangle of 54°45' to the spin axis, and if the sensor is sampled at three

equallyspaced intervals during the rotation of the spacecraft, the experimenterreceives three independent measurements of three orthogonal components


TABLE 5-1.-Experiments Aboard the Pioneers

Pioneerspacecraft

Instrument 6 7 8 9 E

Single-axis fluxgate magnetometer I .------------------ X X XTriaxial fluxgate magnetometer a - --_________________ _ X XFaraday cup plasmaprobe X X

143



Faraday-cup plasma probe ---------------- ---------- X XPlasma analyzer ------------------------------------ X X X X X

Cosmicray telescope-------------------------------- X XCosmic-rayanisotropydetector --------------------- X X X X XCosmic-raygradient detector -------------------- - - - - - X X XRadio propagation experiment----------------------- X X X X XElectric-fielddetector ------------.-.-.----------- X X XCosmicdust detector------------------------- - - - - - - X X XCelestialmechanics------------------- X X X X X

I The triaxial fluxgate magnetometer was originallyscheduled to fly on Pioneers Cand D, but because it could not make the launch date, the single-axisfluxgate mag-netometer was substituted on PioneerC.

craft can be taken into account. TheGoddard sensor is rotated by a spring-driven escapement mechanism. Because of their high reliability, elevenpairs of small explosive charges were used to activate the escapementmechanism. Thus, eleven sensor flip-overs are possible by remote control.

The spacecraft Sun sensor triggered the Goddard experiment once eachrotation of the spacecraft. Beginning with this signal, the experiment elec-tronics generated three equally spaced sampling gates which permittedanalog readings from the fluxgates to enter analog/digital converters.These data were then converted into digital words. Each magnetometer

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For this reason, the Pioneers were made as magnetically clean as possible(ch. 4), and the magnetometer sensor was located on a spacecraft boom2.1 m from the spacecraft spin axis. Detailed mapping indicated that the

magnetic interference from the spacecraft was less than 0.125%, 0.3 5 y, and0.2? on Pioneers 6 , 7, and 8 , respectively. The Pioneers were among themagnetically cleanest spacecraft ever built. The data telemetered to Earthhave, therefore, been of great utility in mapping the magnetic structure ofsolar disturbances and, during the first few hours of flight, the Earth's

145



magnetic tail.

The overall characteristics of the Goddard magnetometer are tabulatedin table 5-2. Originally, the magnetometer was to be located at the endof the axial high-gain telemetry antenna, but this proved impractical andit was mounted on a boom (fig. 1-4). The location of the experimentelectronics on the spacecraft equipment platform is shown in figure 4-29.

THE AMES MAGNETOMETER (PIONEERS 9 AND E)

The Ames magnetometer instrumentation consists of a fluxgate-sensorpackage located at the end of one of the 62-in. spacecraft booms and anelectronics package mounted on the spacecraft equipment platform. Likethe Goddard magnetometer, the Ames instrument is based on the fluxgatesaturable inductance sensor; but it employed three sensors mounted along

TABLE 5.2.-Characteristics of the Goddard Magnetometers

PioneersCharacteristic

6 7 8


mutually orthogonal axes rather than a single sensor like the Goddardinstrument. One fluxgate axis is parallel to the spacecraft spin axis and asecond oriented radially. The Ames experimenters hoped that their three-

axis magnetometer would provide a better measure of the interplanetarymagnetic field during disturbances involving large, rapid magnetic fluctu-ations.

The permalloy-core fluxgate sensors were built by Honeywell, Inc., andsupplied to Philco-Ford, the magnetometer prime contractor, as Govern-

146



pp , g p ,ment-furnished equipment (ref. 4). The general construction of the basicsensor is portrayed in figure 5-3. A 6144-Hzdrive signal of approximately1.5 V rms applied to the toroidally wound drive winding modulates thepermeability of the sense-winding core. The sense winding provides thesignal that indicates the direction and strength of the ambient magneticfield. The feedback winding generates a signal that helps to minimizenonlinearities. The three sensors comprise two packages: one single-axisfluxgate is located in a package mounted so that the sensor axis is parallelto the spacecraft boom axis; the second package contains two orthogonallymounted fluxgates with both axes perpendicular to the boom axis. TheAmes instrument includes a flipping mechanism, but it is powered by tworesistance-heated bimetallic motors rather than a pyrotechnic device, such

as that used by the Goddard magnetometer. The motors on the Amesinstrument flip the dual sensor assembly 90 ° upon command from Earth.One motor flips the sensors clockwise; the other counterclockwise. TheAmes magnetometer sensors can be flipped again and again and are notlimited to the number of pyrotechnic charges launched with the spacecraft(the Goddard instrument has 1 1 flips); however, an additional burden is

placed upon the spacecraft power supply by the resistance heaters in themotors.

The electronics package (fig 5 4) has these major functions:

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periodic measurement of the zero level of the sensor aligned with thespacecraft spin axis.

The overall instrument parameters-as originally specified and ulti-

mately attained-are presented in table 5-3. The table indicates a changein dynamic range from -2003y to 4-50y after it became apparent from


TABLE 5-3.-Ames Magnetometer Specifications

Realized parameters

Specifica- DesignParameter tion value Proto- Flight Flighttype unit unit

no. 1 no. 2

Weight (lb)B k 0 7 085 0815 0 837

149



Boom package----------- 0.7 0.85 0.815 0.837

(+0. 1)Instrument ----------- 4.6 6.8 5.85 5.79 5.87(40.3)

Powerrequirement (W)Normal------------- 3.0 5.6 5.72 5.5 5.3Calibrate --------------------------. 8.0 9.16 8.7 8.65Flipcalibrate ------------ 6.36 8.3 8.24 8.3 7.8

Reliability prediction------- 0. 75@ 10 000 hrInstrument dynamic range

Prototype and flightno . 2 4 50'y 50yFlight no. 1 .- ±--------0 0 y 4200y 4-50y 200y 4-50y

ResolutionPrototype and flightno . 2 4-. 05y ±0.05y 4-0.05y 4-0. 2 4-0. 05y

Flight no . I .--....... 0.02y 4-0. 2yRepeatability .--. . O.4-0.2y -0. I - 4-0.2y +0. 2 y 4-0. lyAccuracy(percent)-±------ 4-0.35 4-0.35 40.5 4-0.5 4-0.35DC offset . - - - - - - - - - - - - - 4-0.2y- - 0.3 4-0. 4 y 0.4yStep response(percent)----- <1 . - - - - - - - - 1 1 <ICross coupling ..-. <2y -- < 2 y <2y <22

remaining ionized out to several AU. Further, this plasma is electricallyconducting and interacts in complex ways with solar and planetary mag-


sorts out the particles in the external plasma according to species andenergy. The split in the collector is parallel to the spacecraft equatorialplane to provide directional information about the plasma fluxes in the

meridian plane.The energy spectra of the plasma ions and electrons are measured by

applying square waves at different voltage amplitudes to the modulatorgrid directly in front of the split collector. For example, an 1800-Hzsquarewave varying between V and V2 admits only those particles in the plasma

ith gi b t V d V 2 l t volts F th th q

150



with energies between V1 and V 2 electron volts. Further, the square wave

modulates the stream of particles impinging on the collectors so that thecurrents collected and resultant signals delivered to the electronics sectionof the experiment varies at 1800 Hz, a signal that can be amplified andfiltered conveniently. The amplitude of the square wave is varied between100 and 10 000 V in 14 contiguous intervals to scan the positive ion spec-trum and between 100 and 2000 V in four intervals for the electronspectrum.

The instrument's sensor is sampled once during each of the 32 equal11.25° angular increments that the Faraday-cup sees in one completerotation of the spacecraft. Since the interplanetary plasma flows outwardfrom the Sun, samples from the eight segments within 445 of the Sunline are always used to make up a data frame. Five additional samplestaken during a spacecraft revolution complete the 13 data samples thatcomprise the "fine" measurements. Each of these represents the highestflux measurement from the four (11.250) segments in each of the five(45°) sectors comprising the remainder of the complete rotation. Althoughall 32 (11.250) segments are examined during each rotation, only 13plasma measurements are recorded. The 13 samples are then digitized assix-bit words and stored temporarily in a core memory with a 256-wordcapacity.

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of the summing operation, fine data indicate plasma characteristics onlyin the spacecraft equatorial plane. "Coarse" data are obtained by processingthe current collected by a single half of the split collector. The largest of

the 32 measurements taken from this collector during a spacecraft revo-lution comprises the coarse data word. Comparison of coarse and finedata words yields a measure of plasma direction in the spacecraft meridianplane; that is , the shadowing effect of the Faraday-cup walls combinedwith the view angles of the split collectors permits some coarse resolutionof plasma fluxes arriving from above and below the plane of the ecliptic

152



of plasma fluxes arriving from above and below the plane of the ecliptic.

To summarize, a complete data mode of 256 words consists of 15 framesmade up of 13 fine data words, 1 coarse data word, 1 coarse angle word(indicating which of the 32 angular segments provided the coarse dataword) and 1 high voltage calibration word. The sixteenth frame comprises16 words of engineering data. When the 256-word memory is completelyfilled-as it is at the end of a data mode-power is shut off to portions ofthe electronics, and the entire memory is read out to the spacecraft datahandling subsystem. Then a new mode begins. The power drain of theMIT plasma probe therefore varies in a cyclicfashion, as mentioned in thelast chapter.

The block diagram of the electronic circuits required to accomplish theexperiment's formidable switching tasks is shown in figure 5-6.

Angular resolution of the MIT instrument is better than 5° in theequatorial plane of the spacecraft, which is approximately parallel to theplane of the ecliptic. The useful flux range is between 1 0 6 and 4XI09singly charged particles per cm 2 per sec.

A M E S PLASMA P R O B E ( P I O N E E R S 6, 7, 8, 9, AND E)When the angular distributions of the ions and electrons comprising the



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passes only a certain range of energy-to-charge ratios. If the plates aremade portions of spherical surfaces and the collectors are segmented, theplasma flux arriving from different directions can be analyzed. If the

applied voltages are stepped, energy-to-charge spectrum scanning is pos-sible. Through the use of a mass spectrometer as the particle detector,particles with different masses,but within the same energy-to-charge-ratiogroup, can be extracted, but this was not part of the Pioneer experiment.

When the Pioneer payload was being formulated, the Ames ResearchCenter group had flown quadrispherical plasma analyzers on OGO 1 and

155



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the IMP Earth satellites and was therefore in a good position to proposethe similar instruments fo r the Pioneer series.

Although their basic principles of operation were the same, the plasmaanalyzers flownon the Block-IPioneer spacecraft were significantly differ-ent from those on Block-II spacecraft. These differences are summarizedbelow:

Block-I Instruments Block-II Instruments

Quadrispherical plates Truncated hemispherical plates8 current collectors 3 current collectors16 positive ion groups between 200 30 positive ion groups between 150

and 10 000 eV and 15 000 eV8 electron groups between 0 and 14 electron groups between 12 and

500 eV 1000 eV

Block-I Instruments

An entrance slit on the instrument face, figure 5-8, permits electronsand positive ions in the interplanetary plasma to pass into the space be-


Spacecraft Sun sensor'at Sun pulse t ime

156




have been measured for all 24 energy steps, in all 15 sectors. This mode isemployed when the bit rate is 512 bits/sec. Spacecraft format A is used.

(2) Short Scan (SSM)-This mode is identical to the FSM mode except

that only the eight small selectorsclustered around the Sun line, as shownin figure 5-8, are read out. SSM is used at 256 bits/sec, format A.

(3) Maximum Flux (MFM)-In this mode, the eight collectors orchannels are not scanned in sequence. The measurement read out in thismode is the maximum flux measured during each spacecraft revolution.The specific sector and channel where this measurement was made are

157



pidentified in the telemetry. As with the other modes the plate voltageremains the same during each spacecraft revolution and is stepped when anew revolution commences. MFM is used at the 8, 16 , and 64 bits/secrates, format B.

Block-II Instruments

The Block-II instrument configuration is basically hemispherical, withthe entrance aperture defined by a slit, as shown in figure 5-7. A series ofgrids and ground vanes are interposed between the analyzer plates andthe three current-collectors, which are located so that they can monitorparticles arriving from above and below as well as from along the plane ofthe ecliptic. Fluxeswithin -800 of the plane of the ecliptic can be moni-tored. The detectable range of current is 10-14 to 10-9 amperes.

A sector programmer once again is employed to divide the time betweenSun pulses into equal parts; 2048 timing intervals are used to divide theazimuthal plane into 128 equal sectors. During one of the scan modes, the23 shaded sectors shown in figure 5-9 are selected fo r current measurements.

Most of these favored sectors lie near the Sun line. A comparison of figure5-8 with figure 5-9 showsthe differences in sector widths, especially awayfrom the Sun line




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to sunlight; this makes a light-tight enclosure a necessity. Particle absorbersbetween the telescope elements define the response of the elements tovarious particles at various energies.

The analysis of the pulses generated in the four telescope elements iscomplex, as indicated by the supporting electronic circuitry (figure 5-13).Ignoring for the moment the significance of pulses indicating the passageof particles from the four detectors, let us look at the 6 six-bit words thatthe experiment feeds into the spacecraft data handling subsystem. Five ofthese words are displayed on the main scientific data frames while the

16 2



sixth appears twice in subcommutated scientific data (ch. 5). The Chicagoexperiment constructs six "spacecraft" words from five "experiment"words, which are labeled Aa, Ab, Ac, Ad, and Ae. The 18-bit word Abis composed of three contiguous six-bit spacecraft words that are derivedfrom the following experiment components: seven bits from the pulse-height analyzer associated with D1 and five bits from the D 3 pulse-heightanalyzer; four more bits of information concerning the quadrant in whicharriving particles were detected; one bit indicating the counting rate ofcoincidences from all four detectors (DID2D3D4); and one bit showingwhether the experiment is in a normal or calibrate mode.

Words Aa and Ac each contain three bits from each of the two counting-

rate scalers that indicate counting rates for the following coincidence-anticoincidence situations: DID2D4; DID2D3D4; D1D2D3D4; andDlD2D3D4.17The sixth (six-bit) word is subcommutated twice and islabelled Ad and Ae. This word contains five bits of rate information forthe four quadrants of spacecraft rotation for the DID2D3D4 logic, plus aquadrant flag-indicator bit.

Now, consider particles entering the instrument through the solidangledefined by the plastic scintillator. The particles pass through Dl, pro-

ieI

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102Kineticenergy per nudeon (MeV)

FIGURE 5-14.--Dl detector energy loss vs. particle energy, Chicago cosmic-ray telescope.

of 16 0 keV when D1 counts are considered alone.Electrons can be distinguished in the pulse-heightanalysis of Dl signals because they cause mainly low-amplitude pulses. Counting rates alone-that is ,


and also violated magnetic cleanliness specifications. Consequently, majorredesign work was carried out at the University of Chicago in July 1964to bring the experiment within specifications. Weight was shaved off the

final design by such stratagems as the use of hollow mounting screws, theelimination of unused pins on plugs, and the milling of the magnesiumstructures.

An operational problem cropped up at Cape Kennedy because theChicago experiment used-for the first time-large-area lithium-driftedsilicon wafers, which required stringent humidity control. The experiment

165



had to be flushed with dry nitrogen while it was on the launch pad atCape Kennedy until the last possible moment. Ames Research Centerprovided the control system fo r the nitrogen flushing apparatus.

T H E G R C S W C O S M I C R AY E X P E R I M E N T S(PIONEERS 6, 7, 8, 9, AND E)18

The Earth-based study of cosmic-ray anisotropy has always been ham-pered by the presence of the Earth's magnetic field and atmosphere. Evenscientific satellites do not get far enough away from the Earth to avoid itsmagnetic field completely. The crucial test of one theory that describesthe motion of cosmic rays within the solar system depends upon the careful

measurement of cosmic-ray anisotropy at energies below 1000 MeV. Forsuch measurements, the instruments must be carried well away from theEarth. The Pioneer probes were ideal fo r this purpose.

GRCSW instruments were part of all five Pioneer payloads, but thoseon Pioneers 8, 9, and E (Block II) represented a second generation ofequipment (refs. 7 and 8). The later equipment was more sophisticatedbecause additional low-energy measurements were made in, above, andbelow the plane of the ecliptic


Photomultiplier7010E7151M Platic guard scintillator

16 6



(inside)

FIGURE 5-15.-Axial view of the GRCSW cosmic-ray telescope, Block-I Pioneers. Thedetector dimensions and positions were changed for the Block-II flights; see text.

/ Mean viewing direction (EB)

/!

E


instrument calibration. Two sheets of 0.0005-in.aluminized-Mylar coveredeach plastic cup and protected the detectors from light, heat, and micro-meteoroids.

The same basic scintillator arrangement was employed fo r the Block-IIflights, but it was supplemented with a three-way coincidence telescopeconsisting of four 100-micron, totally depleted silicon, surface-barrier de-tectors arranged and labeled as shown in figure 5-16. EB-coincidencelogic counts particles within the energy range 0.5 to 5.0 MeV/nucleonarriving from 480 above the plane of the ecliptic. AB logic permits moni-

i h l f h li i hil l k k f i l

16 7



toring the plane of the ecliptic, while B F logic keeps track of particlesarriving from 48 ° below the plane of the ecliptic.

The detectors scan the plane of the ecliptic as the spacecraft spins. Inthe Block-I instruments, additional directional discrimination is providedby four electronic quadrant gates (Q1, Q 2, Q 3, and Q 4 in fig. 5-17) whichare opened by electronic gates in sequence during each revolution following

a Sun pulse. Quadrants 1 and 3 look away from and toward the Sun,respectively; quadrants 4 and 2 look fore and aft along the spacecraft'sorbit, respectively.

The goal of the experiments was the study of cosmic ray anisotropies assmall as 10-1 of the mean cosmic-ray flux. Consequently, the count-accumulation times for the four quadrant-registers had to be identical toat least one part in 1 0 4 to provide meaningful experimental results. Aunique and critical part of the experiment, therefore, was the precision,crystal-controlled aspect clock that controlled the gating pulses.

The primary modes of operation of the Block-I experiment are listedbelow:

(1) Dynamic range off-The length of each of the four time periods isalmost one-fourth of a spacecraft revolution. This is used when the Sun isrelatively quiet


1A, Pulse (32 Pulses per spin period)16

168



FIGURE 5-17.-Greatly simplified block diagram of the electronics associated with asingle channel of the GRCSW cosmic-ray experiment, Block-I instruments only.

sotropy pulse-height analyzer, a pulse from detector B must fall within the



1mI.0.007

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experiment's spin counter. To conserve power, the output buffer-a serialshift register-was limited to only 21 bits. During readout of the experi-mental data, as soon as 13 bits are shifted out of the buffer into the telemetrystream, nine bits from the isotropy data accumulator are dumped into thebuffer to complete the 30-bit message. Because the experiment countingrate often exceeds the storage capacity of the nine-bit accumulators, a formof logarithmic storage was employed which allowed 2752 counts to bestored in the nine-bit format.

The additional sophistication of the Block-II instruments increased ex -periment

weight from 4 4 to 5 6 lb; the average power drain went from

170



weight from 4.4 to 5.6 lb; the average power drain went from1.6 to 1.8 W. During the hardware development of the experiment, it wasdiscovered that commercially supplied photomultiplier tubes normallycontained so much magnetic Kovar alloy that the magnetic cleanlinessstandards could not be met. Tubes with the Kovar replaced by a nonmag-netic nickel alloy (Alloy 180) were built especially for the experiment.

standards could not be met. Special tubes with the Kovar replaced by anonmagnetic nickel alloy (Alloy 180) were built especially for the experi-ment.

MINNESOTA C O S M I C RAY DETECTOR(PIONEERS 8, 9, AND E)

The Minnesota cosmic ray experiment had a purpose entirely differentfrom that of the GRCSW instrument. The experiment objectives listedbelow reflect the lack of high precision cosmic ray experiments flown onspacecraft prior to the spring of 1964:

(1) To measure the quiet-time energy spectrum of protons, alphas, andheavier nuclei up to a charge of 14 over a wide energy range with betterenergy and background discrimination than previously obtained

(2) To measure the variations in these spectra, including the features of

SCIENTIFIC INSTRUMENTS 171

Aperturewindow

Two-pieceguard counter(pilot B)

3.5 cm 2 X 1m m BASurfacebarrierdetectors G detector

(11% in. type C7151Nphotomultipl ier

, g/ X = tube)

5 cm 2 X 2m mLithium B2



Lithium B2driftdetector

10 cm 2 X 2m m B3 Each1.50Lithiumdrift |C\ gnm/cm

2 copperdrift cz absorbersdetectors } abho.50

A 3 0.069 gm/cm 2 aluminumD detector (3 in. type /C31009 C photomult ipl iertube with 1 cm thick 1 /sapphire window) \

FIGURE 5-19.-Arrangement of detectors and absorbers in the Minnesota cosmic-raytelescope.

Telescope T5 is essentially omnidirectional and samples the particles inthe cosmic-ray flux with energies greater than 14 MeV per nucleon.Electrons greater than 0.6 MeV are alsodetected. In telescope T4, however,

the guard scintillator G is in anticoincidence and only particles enteringthe aperture are counted. Detector B2 sets the range energy. The pulses


TABLE 5-4.-Minnesota Cosmic-Ray Telescope Arrangements

Coincidence-Telescope anticoincidence Charge and energy ranges

requirements of the particles detected

T1, T2 BiA.Bl1BB2,B3sC Z 2 1 E > 64 MeV per nucleone ± E > 8 .4 M eV

T3 B31A.BiB2.C.G e ± 4.2 M eV < E < 8.4 M eV,H' 39.6 M eV < E < 64.3 M eV2H e4 39.4 M eV per nucleon < E < 64.1

M eV per nucleon

172



M eV per nucleonT4 B i B2.G e ± 0.34M eV < E < 4.3 M eV

(B 1 = B1A + BiB) 1H1 3.5 MeV < E < 39.7 MeV2H e4 6.6 M eV per nucleon < E < 39.7

M eV per nucleonT5 BIA-BIB Z > I E > 14 MeV per nucleon

e ± E > 0.6 M eV

along the radio transmission path between the Earth and spacecraft (ref.11). For successfuloperation the experiment required that a dual-channel,

phase-locked-loop receiver in the spacecraft lock onto signals transmittedfrom the 150-ft parabolic antenna located on the Stanford campus. Whenthe experiment is in progress, two modulated coherent carriers of approxi-mately 49.8 and 423.3 MHz are sent to the spacecraft from the 150-ftStanford antenna. The spacecraft receiver measures the relative phasechange between the modulation envelopes. Since the higher frequency is

relatively unaffected by the presence of ionization, the comparison providesthe information needed to compute the integrated electron-number density,

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TA B L E 5-5.— Stanford Experiment Transmitter Characteristics

Characteristi c Channel 1 Channel 2

Transm itt er frequency 49.8 M H z 423.3 M H zType of transmit ter Trio de linear amplifier Klystron

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following then set the 423.3-MHz carrier at exactly 17/2 times the 49.8-MHz carrier. Both carriers are modulated at either 7.692 or 8.692 kHz fo rdifferential group path measurement. Real-time teletype data are sent toStanford from the DSN showing the operating

point of the spacecraftphase meter. This information is used to keep the spacecraft phase meteron its positive slope.

Both carriers from the Earth are received by the Stanford spacecraftantenna (fig. 5-23) and sent to the dual-channel receiver, which consistsof two separate coherent phase-locked receivers (fig. 5-24). The main

reasons for the phase-lock design are: (1) to increase the sensitivity of thereceiver and (2) to detect the difference in radio frequency cycles bet een

176



( )receiver, and (2) to detect the difference in radio frequency cycles betweenthe 49.8 MHz and the 2/17 harmonic of the 423.3-MHz carrier. Additionalreceiver data are presented in table 5-6.

Because the Stanford experiment must have transmitter operators atStanford in the loop during its operation, real-time teletype data are

relayed directly from JPL's SFOF to Stanford (ch. 8 ). Teletyped param-eters include the modulation phase-difference measurements and the rf-difference counts. The Stanford operator uses this information to adjustthe transmitter frequencies, powers, and modulation phase offset fo r bestoperation. At the experiment design range of 300 000 000 km, it takesabout 33 min for the effects of transmitted changes to be seen in the teletype

messages from JPL.Minor changes were made in the experiment during the program and

these are reflected in the slight weight change between Block-Iand Block-IIPioneers. The spacecraft portion of the Stanford experiment weighed 6 .0and 6.3 lb fo r Blocks I and II, respectively. Power consumption was ap-proximately 1.6 W for all flights.

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THE INTERPLANETARYPIONEERS

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TABLE 5-6.-Stanford Experiment, Receiver Parameters

Frequency 49.8 MHz 423.3 M H z

I.F. bandwidths3-dB bandwidth ------------------ 40 kH z 40 kHzNoisebandwidth ------------------ 45 kH z 45 kH z

Receiverinput noise (including image)Noisefigure- --------------------- 3 dB 3 dBNoisetemperature -.------------ 300° K 290° K

Cosmicnoise(in a dipolewith axis 8000

°

K 100°

Kparallel to galacticaxis)

179



SpacecraftnoisepluscosmicnoiseforPioneer7

Noisetemperature -------------- 17 000° K 400° KModulation frequenciesfo r phase meter_ 7. 692 or 8 .692 kH zModulation phase accuracy----------- Digitized to approximately 3.150 per digit;

additional error estimated to be +20Differentialphasepath --------------- Quantized to 1 Hz of rf phase differenceat49.8 MHz

segment of the Stanford antenna (fig. 5-23) as a capacitively coupled

sensorwith which local plasma waves cat be detected. The sensor is rela-tively insensitive, but adequate for the purposes of the experiment. A num-ber of Earth satellites have carried similar vlf radio receivers for the samepurpose.

The availability of only six subcommutated telemetry words restrictedthe experiment's capability to survey a wide range of plasma waves ininterplanetary space. Even when the spacecraft transmits at the highest bitrate of 512 bits/sec the TRW Systems experiment sends only two words

NN0



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TABLE 5-7.-Differences Between TRW Experiments

Difference Pioneer 8 Pioneers9 and E

Engineering wordsfo r frequencycount --------- 2 4Engineeringwordsfo r step number ------------ 2 1Number ofsteps .------------ ---- - - - - - - - 16 8F requency readingsbeforestep is changed ----- 2 1Time for broadband scan----------------- 7.47 min fo r 56 sec fo r

16 channels 8 channels

18 1



band scan is repeated every 7.47 min on Pioneer 8 and every 56 sec onPioneer 9.

THE GODDARD C O S M I C DUST EXPERIMENT(PIONEERS 8, 9 , AND E)

As related at the beginning of this chapter, no cosmic dust experimentswere initially proposed for the Block-II Pioneers, and the Block-Iexperi-ment proposed by Ames Research Center was not far enough along indevelopment to make the Block-I flights.The Block-II Goddard cosmicdust experiment described below is the result of a specific solicitation oflikely experimenters in this field by NASA Headquarters. The followingdiscussion is adapted from Berg and Richardson (ref. 13).

The experiment objectives were four in number:(1) To measure the cosmic-dust density in the solar system well away

from the Earth(2) To determine the distribution of cosmic-dust concentrations (if any)


Cosmic dustparticle

Front f ilm -grid array

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FIGURE 5-26.-Schematic diagram of the Goddard micrometeoroid sensor.

As a high-energy, hypervelocity particle pierces the front film sensor (fig.5-26), some of its kinetic energy generates ionized plasma at the front, or"A" film. The electrons in the plasma are collected on the positivelybiased grid (+24 V); this creates negative pulses, as shown. The positiveions in the plasma are collected on the negatively biased film (-3.5 V);this produces a positive pulse that is pulse-height-analyzed to measure theparticle's kinetic energy. The same thing occurs at the rear sensor, or"B" film; this generates a second set of plasma pulses. Impact on the plateproduces an acoustical pulse. A peak-pulse-height analysis is performed onthe acoustical sensor output as a measure of the particle's remainingmomentum.

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separate output amplifier. The output signals from these amplifiers areused to determine the segment in which an impact occurred. Thus, byknowing the front film-grid segments penetrated and the rear film-grid

segment affected by the impact, one can determine the direction of theincoming particle with respect to the sensor axis and the spacecraft attitude.The solar-aspect sensor determines the Sun line at the time of an impact.

Each of the four vertical films of the front sensor array, as shown infigure 5-27 is a composite of the eight layers shown in the exploded view(fig. 5-28). Ideally, a thin copper foil (500 it) could be used alone fo r the

vertical strips of the front sensor array, but the foil is very fragile and subjectto corrosion. Therefore, the nickel grid, the parylene substrate, and the

18 4



parylene encapsulation are used as supports and anti-corrosion coveringfor the metal film deposits. The aluminum layers, which serve only asfabrication aids during the preparation of the composite film, reflect theintense heat generated by copper evaporation upon the parylene substrate.

Each of the rear sensor array film strips is a 60-,u molybdenum sheet ce-mented to a quartz acoustical sensor plate.

Extensive calibrations were made using a 2-MeV electrostatic accelerator.Unfortunately the particles used for calibration have been limited to high-density, hard spheres of iron (10 -13 g < mass < 10 - 9 g) with velocitiesmerely approaching the low end of the meteoroid velocity spectrum (2 to10 km/sec).

The plasma sensors respond nearly linearly to the particle's kineticenergy over the limited particle parameter range specified above fo r the

~or~0

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laboratory simulator. The acoustical sensorsrespond to the particle's mo-mentum fo r that same particle range.

The threshold sensitivity of the front film sensor array to laboratoryparticles is 0.6 erg. Time-of-flight is registered fo r laboratory particles hav-ing kinetic energies of 1.0 erg or greater. The electronics of the time-of-flight sensor are limited to particles with velocities ranging from 2 to 72km/sec.

Figure 5-29 shows a block diagram of the experiment. A summingamplifier receives the positive pulse from each "A" film strip. After a gain

of unity,the pulse travels two separate paths. On one path it is amplified

15 times; its pulse height is analyzed; and its amplitude is recorded in theh l f d 1000 d f d i

185



storage register. On the other path it is amplified 1000 times and fed intoa threshold one-shot multivibrator. The output pulse performs threefunctions:.

(1) Its origin identification s impressed directly upon the storage register.

(2) It passes through the NOR gate and initiates a time-of-flight meas-urement.(3 ) It is gated back to the threshold one-shot multivibrator to inhibit

any other "A" film pulse until the measurement has been completed.

An inhibit signal to the other three films is necessary to avoid capacitative

crosstalk for high-energy impact signals. The "A" film pulse is pulse-height-analyzed and the results are stored in the register to await readout.

Positive pulses from the "B" film follow similar, but separate, electronicpaths with the following two exceptions: (1) no pulse-height analysis isperformed on the "B" film pulses, and (2) the pulse is used to stop thetime-of-flight clock. If no "B" film pulse followsan "A" film pulse, the

time-of-flight register goes to the full (63-bit) state and remains full untilanother event occurs.


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SCIENTIFICINSTRUMENTS 18 9

9. LEZNIAK, J. A.; A N DWEBBER, W. R.: Astrophysical J., vol. 156, 1969, p. 173.10. LEZNIAK, J. A.: Ph.D. Thesis, Univ. of Minn., 1969.I. KOEHLER, R. L.: Interplanetary Electron Content Measured Between Earth and

the Pioneer 6 and 7 Spacecraft Using Radio Propagation Effects. Stanford Univ.SU-SEL-67-051, 1967.

12. SCARF, F. L.; E T A L : Initial Results of the Pioneer 8 VLF Electric Field Experiment.TRW Systems 10472-6002-RO-00, 1968.

13. BERG, O.; AND RICHARDSON, F. F.: The Pioneer Cosmic Dust Experiment. NASATN D-5267, 1969.

14. A N D E R S O N ,J. D.; A N D HILT, D. E.: Improvement of Astronomical Constants andEphemerides from Pioneer Radio Tracking Data. AAS Paper 68-130, 1968.



CHAPTER 6

T he Pioneer Test and Ground SupportProgram

HE SCIENTIFIC UTILITY of the Pioneer spacecraft rested ultimately upon:

(1) h i hi h i i i li bili (2) h i bilit ith t d th



(1) their high intrinsic reliability, (2) their ability to withstand therigors of the space environment, and (3 ) their "clean" experimental en-vironments-that is , lo w magnetic field, lack of electromagnetic inter-ference, and favorable temperatures. These attributes did not arise auto-

matically. Experience and good design were vital; so was a deep-searching,comprehensive test program.

The Pioneer test program began at the material and component leveland continued until the final moments before launch. The magnitude ofthe test program, as blocked out in figure 6-1,* is impressive, particularlyconsidering that the Pioneers are relatively small spacecraft. The tests thatlogically fall within the scope of this volume are those that begin with thecomponent manufacturer, continue at TRW Systems, and end with thepreship review and final dispatch of the spacecraft and instruments toCape Kennedy (far right of fig. 6-1). The prelaunch activities and count-down at the Cape are mainly relegated to Volume III, except the electricalground support equipment (EGSE) tests and the integrated system tests(ISTs), which are essentially the same as the EGSE tests and ISTs at

S h f


so a special ground test was devised to simulate conditions during deploy-ment. Thus, the primary testing sequence is embellished by many specialchecks not portrayed in figure 6-1. Some of the more critical tests ofthistype are covered below.

TEST SPECIFICATIONS

Just what constitutes a successfultest of a Pioneer component or eventhe entire spacecraft? Test specifications constitute the standards againstwhich all tests are measured. In view of all of the forces exerted on aPioneer spacecraft and the abundant interfaces, it is not surprising to findthe test specifications rather voluminous (ref. 1). The Pioneer test specifi-

192



cations delineate the 1 1 classes of tests defined in table 6-1. Test specifi-cations include requirements placed upon the test facilities employed andstipulate exactly how the tests will be made and witnessed, and the docu-mentation required of the contractor. A typical development test matrix isgiven in table 6-2.

SPACECRAFT M O D E L S

Throughout tables 6-3 through 6-6, certain spacecraft "models" arementioned; fo r example, the "prototype model" of the spacecraft. Likewise,

figure 6-1 shows the parallel paths of acceptance and qualification space-craft models. The flight models are the spacecraft actually intended forflight. They are identical in almost every respect to the prototype model.The flight models are subjected to the milder acceptance tests, while theprototype must survive the stiffer qualification tests. It is proper to lookupon the prototype model as a machine the engineers could work with, amachine much like the prototype automobiles the car manufacturers sub-ject to grueling tests and design modifications before they commit a design

TEST AND GROUND SUPPORT PROGRAM

TABLE 6-1.-Classes of Tests Used in the Pioneer Program

Type of test Description

Parts, materials, and processes_

Development -. ........ - - - -

Life

Outgassing tests and magnetic properties tests wereused.

Performance verification of breadboards, engineeringmodels, subsystems, etc.-see table 6-2 for a typicaltest matrix.

Tests conducted to establish failure modes, wearout

characteristics, and their effect on spacecraftreliability-many spacecraft parts were subjectedto thermal-vacuum tests for over 6 months.

193



Fabrication.................

Integration.................

Assembly qualification -------

Spacecraft qualification -------

Assembly flight acceptance ..

Assemblies and subassemblies were checked duringfabrication to assure functional integrity.

During spacecraft assembly, compatibility wasestablished by electrical continuity tests, rf inter-

ference tests, etc.Tests conducted on all spacecraft assemblies I underforces usually more severe than those anticipatedduring launch and interplanetary flight-generallyqualification tests were 1.5 times more severe thanexpected conditions. As indicated in figure 6-1,equipment subjected to qualification testing did not

fly. (See table 6-3 for details.)Similar to assembly qualification tests, the purpose ofthese tests was to demonstrate the ability of thespacecraft to meet all performance requirementsunder conditions much more stringent than thoseexpected during flight. (See table 6-4 for details.)They were conducted on a prototype spacecraft

model not intended for flight.Similar to but less severe than the qualification tests,


TABLE 6-2.-Development Test Matrix, Electric Power Subsystem

0~~~~

C~~~~~~~C* o bo

BE~~~~4,.' 0-8

'0'

En CnQ .4M P

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Po~~~~k~~~~~lac~4 4 -~~99 E9 ~ 3 t~

194

Characteristics



Voltage-current -------Power outputImpedance -Magnetic effects

Charge-dischargeAmpere-hour capabilityFloating mode ...... _Switching capability ----Dynamic range ---------ResolutionPower input

Noise generationCharge requirements ------Dropout voltage - - -Sensitivity ----------------Time delay .--------Operation after undervoltageGeneral operation ---------

Noise susceptibility -Solar array operating point

XX

X XX X

X

XXX X X X

XXXX

XXXXXXXXX

X X X X X XX

X X X X X X

X XX XX X

X X

X X X X

XX

XX

X

X

XXX

X XXX

XXXXX

X

TEST AND GROUNDSUPPORT P R O G R A M

TABLE 6-3.-Assembly Qualification Tes t Details


Humidity -------------

Vibration ------------

Acceleration ----------

This test aimed at preventing damage from humidity duringshipping and storage; performance of assemblies was notto be degraded by 24 hr in humidity chamber at:86 4 5 ° F; humidity, 95+35 percent.

Assemblies were vibrated in each of the three orthogonalaxes. The specific frequencies, durations, etc., are toolengthy to list here; they included both sinusoidal andrandom-vibration test schedules.

Thrust was 1.5 times the maximum acceleration expected inpowered flight (See ch 7 for Delta characteristics )

195



Thermal-vacuum ------

Shock ---------------

Magnetic -------------Solar array -----------

powered flight. (See ch. 7 for Delta characteristics.)Spin was 18 5 rpm, as compared to the 60 rpm expectedduring normal cruise.

Pressure was less than 10-5 torr; temperature was 250 Fabove the predicted maximum and 250 F below thepredicted minimum; during the 24-hr exposure, cold-startcapability had to be demonstrated at least three times forcyclically operated components.

Three shocks of 50 45 g peak for 6+' msec were applied ineach of the three orthogonal directions.

See chapter 3.Calculated performance figures were verified in sunlight at

the JPL Table Mountain facility and facilities at PalmSprings.

TABLE 6-4.-Spacecraft Qualification Test Details



TABLE 6-4.-Spacecraft Qualification Test Details (Concluded)


Thermal-vacuum- -- -

Magnetic properties

The spacecraft was tested in thermal-vacuum chamber atpressures less than 5 X 10- 6 torr, with simulated solarradiation and with chamber walls cooled to - 305 4- 150 F;insolation was simulated between the values expected at 0.8and 1.2 AU; test duration was at least 9 days; spacecraftwas spun fast enough to stabilize temperatures of solararray; and the solar wind was simulated. (These thermal-vacuum tests were not carried out on the prototype modelof the spacecraft.)

Magnetic measurements were made at 34 specified operating

196



Magnetic properties ---

Electromagneticcompatibility

Boom deployment -----

Subsystem qualification_

Spacecraft/solar-arrayelectrical compatibility

Magnetic measurements were made at 34 specified operatingmodes. (See ch. 3 for magnetic cleanliness philosophy.) Fourtypes of magnetic tests were conducted on the prototypespacecraft: Type I-spacecraft magnetic field mapping;Type II-spacecraft stray field; Type III-solar arraymapping; and Type IV-solar-array stray fields.

Systems performance tests were to assure that no subsystemswere adversely affected by the operation of the rest of thesubsystems.

These tests under near-zero-g conditions at expected spinrates were to see if all booms deployed satisfactorily; astructural model was.used rather than the prototype.

Some of the more critical subsystems were qualification-tested separately.

Compatibility of complete spacecraft with flight powersupply was tested during simulated operational conditions.

metal or other material. In a sense, these models paralleled the customary


TABLE 6-5.-Assembly Acceptance Test Details


Vibration ------------ Similar to qualification tests except that only sinusoidalvibration schedule used; levels not exceeding expectedflightlevels

Thermal-vacuum ---- Tests conducted within maximum and minimum expectedtemperatures only fo r 1 2 hr; otherwisesimilar to qualifica-tion tests

Magnetic properties --- Same as qualification tests (table 6-3)Solar array ----------- Same as qualification tests (table 6-3)

197



TABLE 6-6.-Spacecraft Acceptance Test Details


Initial balance -------- Similar to qualification tests, except that degree of balancehad to be within 1.5 times the valuesspecified

Vibration ---------- Similarto qualification tests; random and sinusoidalvibrationschedulesused

Thermal-vacuum ----- Similar toqualification tests, except test lasted only 7 days(table 6-4)

Final balance . - - - - - - - -Balanced prior to shipment to Cape Kennedy and (forPioneer 6 only) again beforemating with live third-stagemotor; balance weights added to bring degreeof balancewithin stipulated values

chamber walls are cooled cryogenically to simulate the blackness of space

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laboratory was a dynamic balancing machine upon which degree of balanceand moments of inertia could be measured.

To test the boom deployment sequence under simulated space con-ditions, the spacecraft was first spun up to about 110 rpm on a spin table,as shown in figure 6-6. Zero-g conditions were then simulated by a verticallift of the entire spinning spacecraft (ref. 4). During the coast phase, thebooms were deployed. Deployment was observed visually to check thesmoothness of the operation, and the possible introduction of spacecraftwobble. A nine-channel telemetry system transmitted additional infor-mation on joint angles and stressesduring deployment. The entire deploy-ment scheme was a worrisome point during the program. Fortunately,these tests showed that the design was sound.

The solar array was also subjected to a special series of tests (ref 5)

202



The solar array was also subjected to a special series of tests (ref. 5).A special outdoor test fixture was constructed that could accept eitherengineering model panels or the complete solar array. The test fixture

consisted of a rotating dummy spacecraft, a simulated electronic load,temperature and power output monitoring devices with sliprings, and alarge collimating tube plus standard-cell mount that could be pointed atthe Sun (fig. 6-7). The initial test was set up at the JPL Table Mountainsite in California; the rest of the panels were tested at Palm Springs.

SPACECRAFT INTEGRATION AND TEST PROCEDURES

The test cycle begins with spacecraft component tests and continuesthrough a graduated series of production tests to final assembly. Each as -sembly undergoes the complete regimen of environmental and functionaltests described earlier. Figure 6-1 showsthe overall plan on a broad scale.

Upon completion of the assembly tests, subsystems can begin to be inte-grated into the spacecraft The spacecraft integration and test phase is

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TA B L E 6-7 .— Tests Performed During Spacecraft Integration


Electromagnetic T h e spacecraft, including the instruments, was operated in


whereby changes in performance throughout the test program can bedetected. Deviations between the prototype and various flight models canalso be checked. The IST was repeated frequently so that engineers couldsee trends that might develop due to subsystem degradation. The ISTs

measured the most important system-wide parameters with the spacecraftas near flight configuration as possible. Some of the parameters measuredduring the IST are listed below:

(1) Each receiver was frequency-addressed and its performance verified.The quality of the demodulated command signals fed to the decoder wasverified.

(2) The rf power radiated from the spacecraft was verified in variousspacecraft modes.

(3 ) The CDU operation was verified in all spacecraft modes, including

205



p gordnance-control circuitry and undervoltage output signals.

(4) Simulated Sun pulses were inserted and the pneumatic pressureswitch was monitored to check the operation of the pneumatic equipment.

(5) B it rates, bit formats, and the quality of the data transmitted weremonitored to check the DTU and the DSU.

(6 ) Spacecraft bus current was monitored continuously and comparedwith the nominal power profile.

(7) Each scientific instrument was tested to verify performance duringtypical operation.

In a sense, the IST was a thorough physical examination fo r the completeoperating spacecraft. The IST was repeated at least twice for each space-craft when it reached the launch pad.

ELECTRICAL GROUND SUPPORT EQUIPMENT

Although the operations at Cape Kennedy prior to launch are covered




FIGURE 6-8.-PioneerA spacecraft test history.


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terms of detecting failures. It is rather surprising that the qualificationtests, which were more severe than the acceptance tests, did not encountermore failures. A possible explanation of the difference lies in the fact thatthe qualification tests were made only on qualification units, while allflight units, plus the spares, had to pass the acceptance tests.

REFERENCES

1B

J1

CHAPTER 7

The Delta Launch Vehicle

WHY THE D E LTA ?

tHE DELTA LAUNCHVEHICLE, or Thor-Delta,has been one of NASA'sT most successfullaunch vehicles (fig. 7-1). As of July 1, 1970, 71 Delta

launches have succeeded, with only 7 failures noted on the record books.



This remarkable dossier was not available to NASA mission planners in1961, when the IQSY Pioneers were first brought under discussion. Up to

July 1 , 1961, the Delta had successfully launched Echo 1 , Tiros 2, andExplorer 10, while failing only on its first try (Echo A-l, May 13, 1960).A 75 percent successrecord was extremely good in 1961. Thus, the plannersof Pioneer selected a launch vehicle of high promise.

The Delta had several other points in its favor. It was a low-cost launchvehicle derived largely from previously developed military and Vanguard

Program hardware. Its payload capability for the escape mission wassomething over 100 lb, roughly what NASA wanted for its "precursor"Pioneer. The Delta was also considered NASA's very own launch vehicle,because it had not been obtained through military channels. However,NASAhad procured the earlier Thor first stages directly from the U.S. AirForce and then turned them over to the Delta prime contractor,

McDonnell-Douglas, who had built them in the first place. NASA wasi t t it " t bl " f l h hi l t thi i t th

212 T H E INTERPLANETARY PIONEERS



DELTA LAUNCH VEHICLE

Almost as soonas it was created on October 1 , 1958, NASA began to planits stable of launch vehicles. The favorable record of the Able-II STV ledNASA to sign a $24 million contract with Douglas Aircraft on April 29,1959, for the design and manufacture of 12 Able-based launch vehicles.

Originally, the Deltas were intended only as interim launch vehicles forthe 1960-1961 period-something to fill the gap while bigger boosterswere being developed. As it turned out, the later Deltas could easily andvery reliably orbit satellites weighing up to almost 500 lb. This payload wasmore than ample for most NASA scientific missions. The "interim Delta"did not fade away but became instead a workhorse that has propelled morethan three-score spacecraft into orbit around the Earth or Sun.

The Delta is. basically a three-stage rocket. The liquid first and secondstages are topped by a small solid-propellant third stage (fig. 7-2). Thefirst stage core is the venerable Thor military rocket burning a hydro

213



first-stage core is the venerable Thor military rocket, burning a hydro-carbon fuel similar to kerosene (RP-1, RJ-l, etc.) with liquid oxygen.This stage is manufactured by the McDonnell-Douglas AstronauticsCompany. The three first-stage engines are made by the RocketdyneDivision of North American Rockwell. The solid, thrust-augmentationrockets strapped on the first stages of later models are Castor rockets,usually produced by the Thiokol Chemical Corporation. The fuel for themuch smaller second stage is unsymmetrical dimethyl hydrazine (UDMH),which is oxidized by inhibited red fuming nitric acid (IRFNA). The secondstage is also a product of McDonnell-Douglas Corporation. It employs anAerojet-General engine. The third-stage solid rockets have been manu-factured by various concerns during the evolution of the Delta: AlleghenyBallistics Laboratory (ABL), United Technology Center (UTC), andThiokol Chemical Corporation (see table 7-1). The Delta is one of NASA's

smaller launch vehicles (first-stage thrust, about 175 000 lb; plus about160 000 lb from solidstrap-ons on later models)


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solid rockets for first-stage augmentation; most of the later models wereTADs.

(4 ) Improved Delta is a delta with the "fat-tank" second stage but nofirst-stage augmentation.

(5) Thrust-Augmented Improved Delta (ITAD or TAID) is a Delta "im-proved" in the mid-1960s by increasing the diameter of the second-stagetank from 32 to 54 in. The burning time of the second stage "fat-tank"Delta was increased from about 160 to about 400 sec.

(6 ) Long- Tank Delta consisted of a Delta with the first-stage tank length-ened by 14.5 ft (fig. 7-3). The first-stage burn time was increased from

150 sec (Pioneer 8) to 221 sec (Pioneer E).Future changes may involve the adoption of the Titan-III transstage engineand introduction of the Delta Inertial Guidance System (DIGS). Thelatter o ld impro e the injection acc rac f th first and d stages

216



latter would improve the injection accuracy of the first and second stages.Changes in the third-stage solid rocket motor did not lead to overall

name changes, but the model designations did change as the original

Long tank Delta

-503.475Thrust-augmentedimproved Delta Fairing

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X-248 was replaced by the X-258, which, in turn, was displaced by theFW-4D, etc., as indicated in table 7-1.

The Delta models applied to the Pioneer Program are specified anddescribed in more detail in table 7-2. The Pioneer Deltas came from five

points along the evolutionary history of the Deltas, but only four signifi-cantly different models were employed:

(1) Pioneer 6: Delta E, TAID with X-258 and Castor I augmentation(2) Pioneers 7 and 8: Delta E, TAID with FW-4D and Castor I aug-

mentation(3 ) Pioneer 9: Delta E, TAID with FW-4D and Castor II augmentation

(4) Pioneer E: Delta L, a long-tank Delta with FW-4D and Castor IIaugmentation.

THE DELTA-SPACECRAFT INTERFACE

217



The spacecraft-to-launch-vehicle interface is more subtle than one

might suppose. Voluminous documents detail the design restraints thataffect spacecraft design. These design restraints are essentially detaileddescriptions of the mechanical, spatial, electrical, and thermal interfacesthat the spacecraft designer must match if his spacecraft is to fi t on therocket and survive the heat and other forces applied during the launchprocess.

The hardware manifestations of interface matching are attach fittings,fairings, thermal insulation, and shrouds. Much of the interface matchingbetween spacecraft and launch vehicle occurs where the bottom of thespacecraft physically meets the top of the third stage. Pioneer 6 waslaunched by a Delta with an X-258 third stage, but the remaining four inthe serieshad to be matched to the FW-4D stage.

A brief description of the FW-4D defines the general physical environ-


TABLE 7-2.-Physical Characteristics of the Pioneer Deltas"a

Characteristic Pioneer 6 Pioneer 7 Pioneer 8 Pioneer 9 Pioneer E

Delta popular name TAID TAID TAID TAID Long-tankDelta

Delta model number E E E E LDelta launch number 35 40 55 60 73

Model DSV-2C DSV-2C DSV-2C DSV-2C DSV-2L-IBHeight with

adapter (ft) 60.4 60.4 60.4 60.4 70.3First Diameter (ft) 8 8 8 8 8

stage Weight (lb) 18 6 000Sea-level

th t (lb) 175 600 175 600 175 600 175 600 172 000

218



thrust (lb) 175 600 175 600 175 600 175 600 172 000

Model Castor I Castor I Castor I Castor II Castor IIFirst Height with

stage nozzle (ft) 24.3 24.3 24.3 24.3 24.3aug- Diameter (ft) 2.6 2.6 2.6 2.6 2.6menta- Weight (lb) 27 600 27 600 27 600 29 600 29 600tion Sea-level

thrust (lb) 16 1 700 161 700 161 700 156 450 156 450

Height (ft) 13 13 13 13 13Second Diameter (ft) 5.8 5.8 5.8 5.8 5.8

stage Weight (lb) 14 000 14 000 14 000 14 000 14 000Sea-level

thrust (lb) 7400 7400 7400 7400 7400

Model X-258 FW-4D FW-4D FW-4D FW-4D

DELT A LA UN CH VEHICLE 219

c ra f t , pa r t i cu l a r l y t he l a rge r s a t e l l i t e s t ha t c an be l aunched w i th t he morepowerfu l vers ions of the D el ta . T h e re la t ive ly smal l P ion eer spacecra f th o w e v e r, w e r e a t t a c h e d t o t h e F W - 4 D a n d X - 2 5 8 b y a s m a l l 9 X 8 i n ,con i ca l a t t a ch f i t t i ng ( f i g . 7 -4 ) . A two-p i ece marmon- type c l amp secu redby tw o bol ts held the spac ecraf t in th e a t ta c h f i t t ing. In f light, t he spac ecra f twe re s epa ra t ed f rom the a t t a ch f i t t i ng by o rdnance cu t t e r s t ha t s eve red bo thbol ts ( the sever in g of on e bol t i s ac tu al ly suff ic ient) . S ep ar at io n spr in gsim pa r t ed re la t iv e ve loc i t i es o f 6 to 8 f t / sec to th e space cra f t .




Fairings and Payload Envelopes

During launch, spacecraft were protected from buffeting and aero-dynamic heating by a thin-walled fiberglass fairing or shroud. All fivePioneers used the so-called "standard" Delta fairing (fig. 7-5). Thisprotective shell is 224 in. long, 65 in. in diameter, and weighs roughly53 5 lb exclusive of any thermal insulation the spacecraft may need.

Once the sensible atmosphere had been breached, the two halves of thefairing were freed by firing explosive bolts. Spring-loaded latches thenpushed the fairing halves aside and the spacecraft, plus the second and

third stages,proceeded unencumbered.The interface between the fairing and spacecraft is primarily spatial;that is , the spacecraft must fit within the envelope defined in figure 7-6.Some of the Pioneers required thin layers of thermal insulation on thefairing nose; the amount depended th t j t l t d d th

220



fairing nose; the amount depended upon the trajectory selected and theresultant heating.

Spacecraft M echanical Loads During Launch

The rocket motors propelling the Pioneer spacecraft into escape trajec-tories generate what is termed the "launch environment" for the spacecraft;that is , the sinusoidal, random, acoustic, and shock loads induced duringlaunch affected spacecraft design. The low-frequency sinusoidal excitationsoccurred mostly at liftoff,during transonic flight, and just prior to first-stagecutoff. iMaximum random and acoustic excitations occurred at liftoff andduring transonic flight. Shocks were transmitted to the spacecraft whenexplosive bolts and other pyrotechnic devices detonated. Finally, accelera-tion or g-loads stressed the spacecraft structure during all phases of powered

flight.The time histories of these mechanical forces vary for eachDelta model

DELTALAUNCH

VEHICLE

3.1.11 Inertial Axes. The spacecraft principal axes of inertia shall beperpendicular and parallel to the spacecraft centerline within anangle of 0.001 rad.

3.1.12 Rigidity. Rigidities of the spacecraft in the launch configuration

shall be such as to make all resonant frequencies of the entirestructure and/or assemblies greater than 5 cycles per second.

3.2 Load Factors. The following critical load factors expressed ingravity units will exist at the spacecraft center of gravity duringlaunch. The side loads are caused by both translational andpitching accelerations.

Axial load Side loadCondition factor (g) factor (g)

120 sec after ignition of 6.20 1.05first stage

221



first stageFirst-stage burnout 14.00 0.70

Third-stage spinup 0 0Third-stage ignition 6485 0

58 7 + (spacecraftweight in lb)

Third-stage burnout 677077 + (spacecraftweight in lb)

The specifications above are taken from NASA-ARC SpecificationA-6669.07, dated November 13, 1964. Both specifications and launchvehicle restraints changed frequently during the history of the spacecraftprogram.

Specifications numbered 3.1.10 and 3.1.11 introduce another mechanical

THE INTERPLANETARYPIONEERS

Notes1. Drawing n ot to scale.2. Dimensions are in inches.

DSV-3ESta-106.082

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Clearance for fairingseparation spring cartridgeand cartridge guide.

d)

Access to th e spacecraft attachfitting mus t be provided to installdam p, torque bolts ,and to installand checkout all pyrotechnics.

FIGURE 7-5.-Payload envelope of the improved Delta shroud. From: SpecificationA-6669.07, fig. 3.1.2.

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231




TABLE 7-3.-Events Planned in Pioneer-9 Trajectory Sequence

Event

Stage- liftoffBegin stage- roll programEnd stage- roll programBegin stage- pitch programEnd first pitch rate--stage 1Begin second pitch rate-stage 1Solid motors burnout

End second pitch rate--stage IBegin third pitch rate--stage 1Jettison solid motor casingsEnd third pitch rate-stage 1Begin fourth pitch rate-stage 1End stage-I pitch program

232

Time (sec)

0.0002.0003.6704.0009.670

10.00038.190

64.67065.00070.00089.67090.000

130.000



Main engine cutoff

154.531154.531155.581156.331159.531166.531

167.531169.531460.000460.000534.352534.352534.352

534.721

Stage-2 ignition signalStart VCS a channel 1Stage-I separationStage-2 90 percent chamber pressureBegin stage-2 pitch programEnd first pitch rate-stage 2

Begin second pitch rate-stage 2Jettison fairingEnd VCS channel 1Start VCS channel 2End VCS channel 2Second-stage engine cutoff commandEnd stage-2 pitch program

Final cutoff-stage 2

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DELTA LAUNCH VEHICLE 235

by the Delta contractor (ref. 3). Printouts such as this were prepared foreach Pioneer flight. Detailed scenarios were also prepared for each flighttelling each member of the launch team what to do and when to do it.(See Vol. III.)

A TYPICAL WEIGHT BREAKDOWN

The object of all these tables and computer printouts, of course, is theinjection of the small spacecraft, weighing only about 1 percent as much asthe launch vehicle on the pad, into orbit around the Sun. Since this chapter

focuses on the Delta, it will be instructive to see how 99 percent of thelaunch vehicle weight is applied to the 1 percent payload (table 7-4).

TABLE 7-4.-Typical Pioneer Launch-Vehicle Weight Breakdown



Item Weight (lb)

Launch vehicle at liftoff -------------------------- 152 153Vented liquids and gases ------------------------------- - -39Fuel and oxygen burned, stage 1 I------------------------- --26 316Solid propellants burned, augmentation -------------------- --24 786

Launch vehicle at solid motor burnout ..-...................... 01 012Vented liquids and gases --------------------------------- --49Fuel and oxygen burned, stage 1 -I------------------------ -20 898Burned-out solid motor .---------------------------------- --4 803

Launch vehicle afterjettison of solid motor --------------------- 75 262Vented liquids and gases ---------------------------------- -129Fuel and oxygen burned stage 1-------------------------- --52 791


TABLE 7-4.--Typical Pioneer Launch-Vehicle Weight Breakdown (Continued)

Item Weight (lb)

Fairing jettisoned --------------------------------------- -539.45Fuel and oxidizer consumed ------------------------------- -- 10 313.33

Launch vehicle at second-stage engine cutoff command (SECOM) 3068.02Fuel and oxidizer consumed and lost during engine stop transients -26.50Residual propellants -------------------------------------- -294.91Trapped propellants and gases ---------------------------- -65.15

Spin table ettisoned ----------------------------------- -82.06Dry stage 2 (DSV-3E-3) jettisoned ..---------... .. -- 1665.51TETR satellite released --------------------.--------- -58.00

Launch vehicle at stage-3 ignition ---------------------------- 875.89Start losses, stage 3-------------------------------------- -- 0.05Inert loss during burning stage 3-------------------------- -430



Inert loss during burning, stage 3-------------------------- -4.30

Propellant consumed, stage 3----------------------------- --606.90

Launch-vehicle and third-stage burnout ------------------- 264.64Stage-3 motor, ballast, balance weights, support rings, beacon, -46.62

and telemetry kit jettisonedStage-3 motor case jettisoned ------------------------------ -54.45Spacecraft attach fitting ettisoned ----------------------- - -16.01

Spacecraft injected into solar orbit ---------------------------- 147.56

REFERENCES

CHAPTER 8

Tracking and Communicating withPioneer Spacecraft

TRACKING THE FIRST P I O N E E R S

T H EFIRST GROUP O F PIONEERS P A C EPROBES(Pioneers 1 through 5) werelaunched in the direction of the Moon between 1958 and 1960. The

tracking and data acquisition theories and hardware developed by JPL tosupport these flightsultimately developed into the present DSN The DSN



support these flightsultimately developed into the present DSN. The DSN,

managed by JPL fo r NASA, tracks NASA's unmanned spacecraft launchedtoward the Moon, the planets, and deep space.

The basic problem that JPL had to solvein tracking and acquiring datafrom spacecraft beyond Earth orbit involved the immense distances ofinterplanetary flight. Ten thousand miles posed little difficulty, but at tensof millions of miles spacecraft signals faded away amid the radio noise ofinterplanetary space. The Minitrack radio interferometer stations that theNaval Research Laboratory (NRL) had installed around the world during1956 and 1957 for the IGY could work near-Earth satellites, but they couldnot detect faint signals from deep space with their low-gain antennas.Conventional radars could not track spacecraft much beyond 1000 miles.Therefore, new techniques were needed fo r deep space tracking.

Three fundamental concepts permit the successfultracking of very distant

238 T H E I N T E R P L A N E TA R Y P I O N E E R S

de tec t ion o f s igna ls by t he DS N , bu t i t is i nd ep en de n t o f t he p s eu do ran do mn o i s e a p p r o a c h t o t r a c k i n g .

J P L p u t it s t r ac k in g an d da t a acqu i s i t i on con cep t s i n to p rac t i ce p r io r t oO c to be r 1958 , wh i l e it was st il l un de r U .S . A rm y sponso r sh ip . Exp lo re r s 1th r ou gh 5 we re t r ack ed by a handfu l o f J P L phase - lock - loop , M ic ro locks ta t ions as wel l as the N R L M in i t r ac k Ne tw or k . By la te 1958 , as the f ir stP ionee r s we re l au nc he d , J P L ha d e s t ab l i shed t r ack ing s t a t i ons a t C ap eC a n a v e r a l , P u e r t o R i c o , a n d G o l d s t o n e L a k e , C a l i f o r n i a . T h e b i g g e s t d i s hin t he embryon ic ne twork was t he 85 - foo t pa rabo lo id a t t he Go lds toneP ion ee r S it e (f ig . 8 -1 ) . Sm a l l e r d ishes we re l oca t ed a t Ca pe C an av e ra l ,F l o r i d a , a n d n e a r M a y a g u e z , P u e r t o R i c o . A 60 -f t, D e p a r t m e n t of D e fe n sed i sh i n H aw a i i an d the f amou s 250- ft an t e nn a a t Jo d r e l l Ban k , En g la nd ,c o o p e r a t e d w i t h t h e J P L s t a t io n s d u r i n g t h e e a r ly P i o n e e r l a u n c h e s .



TRACKING AND COMMUNICATING

DSN evolution since 1960 has been expressed primarily in terms ofphysical size (antenna diameter and new stations), electronic sophistication(masers, lower antenna noise temperatures, etc.). The 85-ft dishes, thehallmark of the DSN, are now found near Madrid, Johannesburg,Woomera, and Canberra, as well as Goldstone. A 210-ft paraboloid wasadded at Goldstone in 1966; others are under construction at Madrid andCanberra.

When the Pioneer Program began in late 1961, there was no questionabout network choice. The DSN was the only one of NASA's three networksthat could track and communicate with a deep-space probe. Like the Delta

launch vehicle, the DSN became a basic, general-purpose pillar of thePioneer Program-but a pillar already in place that could be altered-verylittle for any specificmission. Even more than the Delta's, the basic capa-bilities of the DSN helped shape the Pioneer spacecraft design.

S O M E GENERALITIES

239



S O M E GENERALITIES ABOUT TRACKING ANDD ATA ACQUISITION

The three basic functions performed by terrestrial ground-supportequipment during the Pioneer missions were:

(1) Tracking-Spacecraft position was measured with high precisionfrom liftoff at the launch pad to injection into parking orbit, through the

coast phase, to injection into heliocentric orbit, and as far out in deep spaceas possible-several hundred million miles and more if possible. Out toabout 10 000 miles this function was accomplished by the Near EarthPhase Network, which consists of MSFN and U.S. Air Force precisionradars; beyond 10 000 miles, the DSN performed this function.

(2) Communication or data acquisition-Scientific and housekeepingdata were detected and acquired from the spacecraft and routed from the


creased the Delta's payload capacity during the same period. In addition tothe evolutionary improvements, some of the 85-ft MSFN antennas adjacentto the DSN antennas for redundancy during Apollo flights were pressed intoservice tracking Pioneers while they were still relatively close to the Earth.With the Apollo, Mariner, and Pioneer Programs, NASA had so manyactive spacecraft in deep space that it pooled its big antennas to achieveoptimum coverage.

In general terms, the DSN carries out its three basic functions using threedistinct facilities (ref. 1):

(1) The Deep Space Instrumentation Facility (DSIF) consists of the

DSN tracking and data acquisition stations shown in table 8-1.(2) The Space Flight Operations Facility (SFOF) is located at JPL, in

Pasadena, California; it monitors all spacecraft data, issues commands, andperforms all necessary mission calculations.

(3) The Ground Communication Facility (GCF) ties all DSIF stationsto the SFOF with high-speed, real-time communications. The bulk of

240



g p ,

DSN communication traffic is carried via NASA's global communicationsystem, NASCOM, which contributes circuits to the GCF.

One other network of equipment crucial to the Pioneer mission is theEastern Test Range (ETR) run by the U.S. Air Force. ETR radars,optical instruments, and other tracking equipment follow all launches fromCape Kennedy down range past Ascension Island, over Africa, into orbit,

where NASA networks assume the full tracking load. They are considered

TABLE 8-1.--The DSN Stations

DSS Dish Primary during

no Location size Pioneerflight


part of the Near Earth Phase Network during the early portions of thePioneer missions.

To complete the picture of the DSN, JPL engineers often visualize thethree facilities just described as vertical sinews interwoven with six hori-zontal sinews representing the groups of equipment that accomplish thetracking, data acquisition, and command functions as well as those ofsimulation, monitoring, and operational control (fig. 8-2).

When Pioneers 1 through 5 headed for deep space, they were the onlyactive spacecraft beyond Earth orbit. The few trackers in existence couldfind and follow these space probes readily, though without great precision

and not very far. It was a "simple" picture in 1960. Today, however,NASA has stationed almost a score of 85-ft dishes and one 210-ft dish atvarious spots on the globe and filled the adjacent buildings with advancedelectronic gear. It is possible to listen to, track and command spacecraft 200million miles away from Earth. Because of many currently active space-craft, DSN priorities have to be assigned to each spacecraft; and each

241



tracking station, being only a part of a tremendously complex machine,operates on a rigorous schedule. The DSN is a world-wide data collectorfor scientists.

GENERAL D E E P SPACE NETWORK CAPABILITIES

After the Pioneer Program was officially approved by NASA Head-quarters on November 9, 1962, spacecraft design and mission planningcommenced at Ames Research Center. The capabilities of the Delta launchvehicle helped fix the weight and volume of the spacecraft, while theDSN-as it was projected for the 1965-1969 period-had considerableinfluence over spacecraft antenna design, frequency selection, telemetry bitrates type of telemetry and many other facets of spacecraft communication


M i s s i o nn D ,rn r t S !$

S C FMission

testingeting I F 0

242



DSNoperationss u p p o r t FromFIGURE 8-2.-The Deep Space Network can be visualized as three facilities (DSIF, GCF,

and SFOF) interwoven with six systems.

JPL. The spacecraft must carry a phase-coherent transponder for the DSNto track the spacecraft satisfactorily.

The DSN stations listed in table 8-1 were deliberately placed about 1200apart in longitude in a band between 400 north and 400 south latitude.

Overlapping sky coverages result with the 85-ft dishes (figs 8-3 and 8-4)

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two-way, phase-locked mode. A less accurate one-way Doppler mode issometimes used by stations that are merely listening to the spacecrafttransmissions. Drifting of the spacecraft crystal-controlled oscillator limitsthe precision when the spacecraft receiver is not locked onto the Earthtransmitter's frequency. When two separate but intercommunicating DSNstations have the spacecraft in view, three-way modes are possible, with onestation in a two-way mode and the other in a one-way mode. The accuracyof DSN ranging is approximately 4-15 m one way (three-sigma value).Range rate accuracy varies with the magnitude of the Doppler shift.

The standard DSN site with its 85-ft dish depends upon 14 subsystems

(ref. 1). See figure 8-5. A few important features are:(1) Antenna mechanical subsystem-Most of the 85-ft dishes are S-band,

Cassegrain feed, and monopulse in operation. The antennas point with anaccuracy of 0.020 in a 45-mph wind.

(2) Antenna microwave subsystem-Some of the most critical andsophisticated DSIF components are included here: Cassegrain simultaneous

245



lobing feeds, traveling-wave masers, and parametric amplifiers. Beam-widths to the half-power points are 0.32 4- 0.030 and 0.36 4- 0.030 forreceive and transmit modes, respectively.

Acquisition-aid subsystem-DSN stations 11, 41, 42, and 51 are equippedwith S-band antennas with beamwidths of 160 to help lock the 85-ftantennas, with their much narrower 'beamwidths, onto the spacecraft.

The Goldstone Mars station (DSS-14) is equipped with a 210-ft dish onan azimuth-elevation mount. This big antenna is more sensitive than the85-ft dishes and is essential for the tracking of Pioneer spacecraft over 100million miles away. The nominal beamwidths to the half-power points ofthe 85-ft dishes are 0.1350 and 0.1450for receive and transmit, respectively.During Pioneer operations the beamwidths have appeared to be about0 200 Pointing accuracy is 50 arcseconds

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T R A C K I N G A N D C O M M U N I C AT IN G 247



FIGURE 8-6.—Pioneer Mission Operations Center at Ames Research Center.

first on s eve ra l N A S C O M o ve r seas swi t ch ing cen t e r s , w h ic h in t u rn rou t eit t o t h e c e n t r a l c o m p u t e r - c o n t r o l l e d s w i t c h i n g c e n t e r a t G o d d a r d S p a c eF l i g h t C e n t e r, G r e e n b e l t , M a r y l a n d . F r o m G o d d a r d , t h e tr af fic is d i r e c t e dto t he add re s s i nd i ca t e d on the message . In t he ca se o f P ionee r, t he S F O Fa t J PL is t he ma jo r a dd re s se e , a l t ho ug h ind iv idu a l D S IF s t a t i ons c ana d d r e ss o n e a n o t h e r G o d d a r d S p a c e F l ig h t C e n t e r m a n a g e s N A S C O M



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time it would take light to travel the same distance. In effect, "real time"means delays of only tenths of a second at most. The teletype circuits,however, are a little slower, with ¼4-secdelays at each control point (up toa maximum of three control points). The site communication's processors,however, introduce delays of 30 to 120 sec. During the early parts of itsflight, a Pioneer often transmits at its maximum rate of 512 bits/sec,which is greater than the teletype rate of 60 words/min. During this timeit is possible to call up selected blocks of data via the teletype circuits inorder to assess the condition of the spacecraft.

The Space Flight Operations Facility

The focal point of DSN activity is a modernistic four-story building atJPL, in Pasadena; this building is the SFOF. The SFOF terminal of theGCF occupies part of the basement. Above are the computers, displays,controls, and facilities for mission control, a major part of which involves

251



DSN control.Brief descriptions of the eight subsystems that make up the SFOF follow:(1) Data Processing Subsystem (DPS)-The function of the DPS is the

ingestion of DSIF tracking data and its subsequent processing into theformats required for display and control. General-purpose digital com-puters are the mainstay of the DPS.

(2) Computer Input/Output (I/O) Subsystem-Consoles, printers,and plotters provide one interface between the Data Processing Subsystemand SFOF users.

(3 ) Data Processing Control and Status (DPCS) Subsystem-Threeconsoles are used here to monitor and control the Data Processing Sub-system.

(4) Telemetry Processing Subsystem (TPS) The TPS performs real





Special Pioneer Requirements Placed on the DSN

The DSN was better equipped during the launches of the Block-IIPioneers than it was when Pioneers 6 and 7 first probed deep space. Withthe 210-ft Goldstone antenna in operation, the DSN could track andcommunicate with spacecraft most of the way around the Sun-the Pioneersare some 150-200 million miles away before the Sun's radio noise over-whelms their telemetry signals. As successfollowed success in the PioneerProgram, scheduling the tracking time of the big DSN dishes and thoseborrowed from the MSFN became a more difficult task. The swellingnumber of manned and unmanned lunar spacecraft

added to the trackingburden.The finite resources of the DSN dictates careful planning to avoid super-

saturation. The Office of Tracking and Data Acquisition, at NASA Head-quarters in Washington, serves as a focal point where requirements,priorities, and resources are weighed for all NASA missions and all three

NASA k A d d h fl ibl d h d l d

253



NASA networks. A standard yet rather flexible procedure has developed.The project requiring tracking and data acquisition support issues a"requirements document" called a SIRD (for Support InstrumentationRequirements Document). The SIRD collects priorities, requirements, andother important factors for all Pioneers in space and those being readied forthe launch pad. In response to each SIRD, JPL issuesan NSP (or NASA

Support Plan) relating how it plans to meet the requirements. GoddardSpace Flight Center does the same for MSFN support. Each documentmust be the result of considerable negotiation and balancing of priorities.The SIRDs are updated frequently to reflect changing demands. Forexample, the launching of a new Pioneer or the loss of signal from an oldone might be significant enough to require SIRD updating.

Requirements set forth in the Pioneer SIRDs over the years have been

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TABLE 8-3.-General Pioneer Tracking Requirements as of March 1969a

Pioneer Nominal mission

6 DSS-14 daily coverage 4-8 hr/day; absolute minimum, DSS-14 dailycoverage 3 hr/day

7 DSS-14 daily coverage; 4-8 hr/day; absolute minimum, DSS-14 dailycoverage 3 hr/day

8 DSS-12, -42, -51, -62 and DSS-I 1 , -42, -61: continuous coverage; absoluteminimum, DSS-12, -42, -51, -62 and DSS-11, -41, -61; two trackingmissions/day for a total of 16 hr/day

9 DSS-12, -42, -51, -62; continuous coverage; absolute minimum, DSS-12,-42, -51, -62 and DSS-I 1 , -41, -61 and MSFN; two tracking missions/dayfor a total of 16 hr/day, with 1 hour overlap

E DSS-12, -42, -51, -62; continuous coverage; absolute minimum, DSS-12,-42, -51, -62 and DSS-I 1 , -41, -61 and MSFN; two tracking missions/dayfor a total of 16 hr/day

These requirements vary with time, of course. This table is illustrative only.

257



These requirements vary with time, of course. This table is illustrative only.

(2) Thirty-first day to end of mission, two passes per day (coverageperiod 16 hr or greater)

(3) For duration of mission, at least one horizon-to-horizon two-way-

Doppler tracking mission per week not to be on same day of the week(4) Coverage of specific scientific events that offer single time periods

within the flight mission when the data may be retrieved; when in effect,this requirement to take priority over all those noted above

(5) Solar-flare coverage, 30-50 hr from flare initiation for Class-IIBright or greater; upon occurrence, this requirement to take priority over

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TABLE 8-4.--Typical Tracking Requirementsfor a Pioneer Flight

Time/distance coverage Data required Data presentation

Class-I requirement a Time, azimuth,A. Launch-vehicle second-stage elevation, range

engine cutoff (SECO) Data points per sec:to SECO-plus-60 sec /ho minimum,(fig. 8-11). A ddesired,

B. Launch-vehicle third- M / maximum bstage burnout to third-stage spacecraft separa-tion (minimum of 60sec of data if available).

Class-II requirement D

A. SECO to SECO plus 18 0sec.

B. Ascension (ETR Station12) rise to Ascension set.

The data to be converted forpresentation in NRT byteletype to the SFOF asfollows:

(a) Decimal raw-dataformat(b) Orbital elements andinjection conditions ofparking orbit(c) Orbital elements andinjection conditions oftransfer orbit assumingnominal third-stage burn(d) Orbital elements andinjection conditions oftransfer orbit based on

258



Class-III requirement a

SECO to third-stage ignition;third-stage spinup to third-stage burnout; DSStracking coverage

sufficientto define the free-flightorbit (figs. 2-6 through2-9).

Acceleration

transfer orbit based onactual third-stage burn

Voice link and/or single-sideband data link in NRT;initially launch plus ap-proximately

2 hr, and asrequired to meet accuracyrequirements c

a See table 8-4 for priority definitions.b These orbital criteria are stated to fulfill project orbital determination requirements

onl ; they in no way reflect the tracking requirements established by the celestial

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TABLE 8-6.-Pioneer Ground Command Requirements

Coverage

Time or phase of orbitRise to set of DSS equipped with Pioneer mission-dependent GOE.

DSN stationsDSS-12-Echo, Goldstone, Cal.DSS-42-Tidbinbilla, AustraliaDSS-62-Cebreros, SpainDSS-51-Johannesburg, South AfricaDSS-71-Cape Kennedy, Fla. (launch checkout only)

Signal

FrequencyReceiver 1, Channel 6B 2110.584105 MHzReceiver 2, Channel 7B 2110. 925154 MHz

Modulation

260



FSK /PMCoding

150-Hz and 240-Hz tones representing 0 and 1, respectivelyRequired transmitter power

Variable; depends on spacecraft receiver signal strength which is dependent on rangeof 10-kW transmitter at a spacecraft signal strength threshold of -- 150 dBm

Transmission time required to execute commandsTransmission time equals the transit light time to the spacecraft plus 23 sec; the transit

light time varies with spacecraft range relative to Earth; the command message is23 bits in length transmitted at 1 bit per sec.

Format23-bit word (See ch. 4 for details.)



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The above requirements are reduced to 6-8 hr per day to end of missionbecause of spacecraft/Earth distance, spacecraft- and ground-antennacharacteristics, or because only one 210-ft antenna is available for opera-tional support.Priority V (Non-critical).-These requirements are not mandatory for

attaining primary mission objectives:(1) Any tracking coverage in excess of 24-hr coverage(2) Operational integration testing(3) Station time required to test a proposed modification to operational

hardware or software.

SPECIFIC PIONEER NETWORK CONFIGURATIONS

The terrestrial facilities that NASA pooled to meet the requirements ofthe Pioneer flights consisted of parts of the following facilities:

1 . The DSN, which included the DSIF, GCF, and SFOF2. The MSFN, which provided 85-ft-dish support on occasion3. NASCOM, which contributed many circuits to the DSN's GCF

26 3



y4. The AFETR (Air Force Eastern Test Range), which supplied much

of the ground environment from the launch pad downrange 5000 miles toAscension Island, i.e., the Near-Earth Phase Network

Each Pioneer flight could be divided logically into two main phases:

near-Earth phases and deep-space phases. The successfulinjection of thespacecraft into a heliocentric orbit was the event that effectively separatedthe two phases (fig. 8-11). At this point, somewhere over the Indian Ocean,the spacecraft would be handed over completely to the DSN and cooper-ating MSFN stations. Each phase of tracking required a different con-figuration of tracking, data acquisition, command, and ground communica-

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TABLE 8-7.-Configuration of Tracking and Data Acquisition Stations During

Near-Earth Phases

Station Tracking Use duringnumber Location radars Telemetry Pioneer flights

6 7 8 9

AFETR 1 Cape Kennedy ------- - FPQ-6FPS-16,TPQ-18

3 Grand Bahama--.- FPS-16,TPQ-18

7 Grand Turk ------------ TPQ-18

91 Antigua FPQ-6

vhf, S-band X X X X

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vhf X X X X

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12 Ascension -------------- FPQ-18,FPS-16

13 Pretoria - - MPS-25

- Twin Falls (ship) FPS-16

-- Coastal Crusader (ship) ...

MSFN 1 Bermuda -------------- FPS-16,FPQ-6

vhf, S-band X X X X

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TABLE 8 - 7 .- Configuration of Tracking and Data Acquisition Stations During

Near-Earth Phases-Concluded

Station Tracking Use duringnumber Location radars Telemetry Pioneer flights

6 7 8 9

72 Ascension --------------- X X X X

5 1 Johannesburg ------- - X X X X

- S F O F, Pasadena -------- X X X X

- BuildingAO, Cape X X X XKennedy

operationalinterface between the SFOF, in Pasadena, and the Air Forceand Goddard Space Flight Center groups. In view of the manifold opera-

265



p g g p ptions at Cape Kennedy, their complex interactions, and the immensedetail required for effective coordination, such interface groups are essential.The JPL field station also contains an Operations Center with abundantdisplays of different types to help JPL personnel operate range instru-

mentation under their control. Critical tracking and telemetry data are alsorouted to the SFOF through the field station.All launches from Cape Kennedy are under the direct control of the Air

Force until the spacecraft leave ETR jurisdiction somewhere beyondAscension.Because it is responsible for range safety, the Air Force monitorslaunch vehicle status data and tracking information. Commands to

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manded by the primary DSS stations listed in table 8-1. MSFN and otherDSIF stations worked the Pioneer spacecraft on an as-needed basis. Com-munication traffic flowed back to the. SFOF and NASA/ARC over GCFlines; commands, of course, moved in the opposite direction (fig. 8-13).

Each of the primary DSS stations was outfitted with so-called "mission-dependent" equipment that accommodated general-purpose DSIF ma-chinery to specific Pioneer requirements. In Pioneer vernacular, the DSSgear was called Ground Operational Equipment (GOE). No specialequipment was installed at the SFOF, although a general-purpose mission-support area was reconfigured for the Pioneer missions (fig. 8-10). Addi-tional mission-dependent equipment was installed at Ames ResearchCenter (fig. 8-6). Since the presence of Pioneer mission-dependent equip-ment constituted the major difference between a DSS station in the Pioneer

266




configuration and any other mission-dependent configuration, a fewdetails are in order.

The Pioneer GOE was designed to make maximum use of the general-purpose DSS equipment, particularly the Telemetry and CommandProcessor (TCP) equipment (the SDS-910 and SDS-920 computers)(fig. 8-14). Type-I GOE, consisting of five racks of electronic hardware,plus a module tester, a test transponder, and an instructor control set, wasinstalled only at the Goldstone site. The primary .overseas DSSstationsreceived Type-II GOE, consisting of three racks only. The two extra racksat Goldstone were the recorder and display racks employed during thespacecraft Type-II orientation maneuver. The Woomera site (DSS-41)possessesno GOE equipment. The specific pieces of equipment in bothtypes of GOE are indicated in the labels on figures 8-15 and 8-16. As theblock diagram, figure 8-14, indicates, the GOE was actually specializedinterface between the antenna and the D SS equipment.

T E L E M E T RY CAPABILITIES

267



Telemetry capabilities were provided as follows:(1) Provided up to 120 frames of continuous spacecraft telemetry data

for near-real-time teletype transmission to the SFOF (TCP)(2) Provided selected spacecraft engineering data for near-real-time

teletype transmission to the SFOF (TCP)(3 ) Provided local typewriter printout in real time at each DSN station

of selected spacecraft engineering data periodically and upon request(This capability accommodates operational requirements during spacecraftinitial acquisition as well as routine orbital operations.) (TCP)

(4) Drove local DSIF displays for spacecraft parameters necessary for


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A similar situation existed with the SFOF; it is another general-purposefacility that was modified to accommodate Pioneer requirements. A PioneerMission Support Area was set up at JPL's SFOF for use as an operationalcontrol center during the launch phases of Pioneer flights when activity was

high. Spacecraft performance and scientific data analyses were also carriedout there. One of the special SFOF Mission Control Rooms and an asso-ciated Flight Path Analysis Area were made available to the PioneerProject during critical phases of each mission. A data flow diagram (fig.8-18) illustrates the routing of data within and without the SFOF duringthe Pioneer Program. The dispatch of data packages to Ames Research

Center (fig. 8-18) completed the DSN role in data processing and handling.Ames processed data tapes and passed scientific data on to the experi-menters, completing the data link from spacecraft to scientist.

Cruise portions of the flights were controlled at Ames where spacecraftand instrument expertise were readily available. SFOF space was used forcontrol only during the launch phase or in the event of extremely critical

periods.

273



REFERENCES

1 . ANON.: DSN Capabilities and Plans, Vol. I. Description of Deep Space NetworkCapabilities as of 1 July 1968. JPL Rept. 801-1, 1968.

2. ANON.: Support Instrumentation Requirements Document, Pioneers 6, 7, 8, 9, and

E. NASA, Mar. 1969.3. RENZETTI,N. A.: Tracking and Data System Final Support for the Pioneer Mission,

Pioneer 6 Pre-Launch to End of Nominal Mission. JPL Rept. TM 33-426, 1970.(A similar report exists for each Pioneer flight.)



ozP



04Cz?L

CHAPTER 9

Pioneer Data-Processing Equipment

P I O N E E RS P A C E C R A F Tradioed back to Earth two kinds of data: (1)scientific data for the several Pioneer experimenters, and (2) engineering

data to permit the mission controllers to assess the operational condition ofthe spacecraft. Referring to figure 8-18, one sees that the telemetry datafollowtwo paths between the DSN stations, which receive it directly fromthe spacecraft, and the experimenters and Pioneer project personnel.Pioneer telemetry data are recorded directly on magnetic tape as theyarrive from deep space at the DSN stations and airmailed to JPL forverification and then to Ames Research Center. This is the first route, andall data follow it. At Ames, they are processed on the Pioneer Off-Line



, y pData-Processing System (POLDPS) for subsequent transmission to theexperimenters on digital magnetic tapes in formats compatible with theircomputer facilities. Some of the telemetry data, however, also follow a

second route. These are dispatched immediately from the DSN to AmesResearch Center via teletype through JPL's SFOF. These are called"quick look" data; they are used for checking the scientific instruments andfo r retransmission (after some processing) to ESSA to help forecast solaractivity. Data from the Stanford radio propagation experiment are handleddifferently. As described in chapter 5, proper operation of this experiment


sible for building the POLDPS hardware. By the summer of 1965, POLDPSwas ready for operation.

Magnetic tape represented the only practical way to transmit the bulk ofthe data from the active Pioneer spacecraft-teletype facilities could nothandle the volume. At each DSN station, two Ampex FR-1400 tape re-corders operating in parallel prepared analog tapes of the transmissionsreceived from the Pioneers. Tape loading times for each machine werestaggered to avoid the lossof data. One set of tapes containing all recordeddata were selected and shipped first to JPL for verification to ensure thequality of reproduction (fig. 9-1). The tapes were then sent to the Pioneer

Off-Line Data Processing System at Ames Research Center.During 1969, Pioneer tape shipments averaged 400 (9200-ft) tapes permonth, each containing 4 hr worth of data with half-hour overlaps.POLDPS processed and sorted out these data, preparing an average of400 (2400-ft) tapes per month fo r the experimenters. The preparation ofover 15 experimenter tapes per working day indicates that POLDPS was

extremely active during 1969, when four Pioneers were transmitting databack to Earth (fig. 9-2).

The input to POLDPS consists of the FR 1400 magnetic tapes received

278



The input to POLDPS consists of the FR-1400 magnetic tapes receivedby airmail from D SS sites around the world. The following seven channelsare recorded on these tapes at 5.5 in./sec:

(1) Voice (containing station events)

(2) Bit clock data from the DSS demodulator/bit synchronizer(3 ) Universal time and 6.25- and 25-kHz reference signals(4) NRZ-C data (see ch. 4)(5) The biphase-modulated 2048-Hz subcarrier containing the time-

multiplexed PCM-data bit train(6 ) Spare channel

DATAPROCESSING

EQUIPMENT

Microwave to DSS-12 GO E

JPL facility

AR C facility

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279



processingstation

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FIGURE 9-2 .—P ioneer O ff -Line Da ta Processing Sys tem (P OL D PS ) a t Ames ResearchCen te r.

s y n c h r o n i z e r(3) Ge ne ra te s t im e info rm at ion ( in day s , h r, m in , sec , an d msec) f rom

the b i t - c lock channe l(4) D em odu la t e s PC M s igna l s an d sync in fo rma t ion f rom the r aw da t a(5) Dig i t izes an a lo g fu nc t ions( 6) C o n v e r t s t h e F S K c o m m a n d d a t a i n t o d i g i ta l f o r m a t

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FIGURE 9-4.-Block diagram of data processing line at TRW Systems used on occasionfor Pioneer tapes.

DATA PROCESSING EQUIPMENT

10 D SS receiver out of lock04 Bit error probability in excess of 10-s02 D SS bit synchronizer out of lock01 Parity error00 Good data

The experimenter tapes are supplemented by trajectory tapes giving thespacecraft position as a function of time. The basic trajectory tapes areprepared at JPL, but Ames reprocesses them to put them in the formatsand densities requested by the experimenters.

POLDPS has remained substantially the same throughout the Pioneer

program. The only significant change was made for Pioneer 9, which carriedthe convolutional coder designed to improve the quality of Pioneer data.The effects of the addition of the convolutional coder upon Pioneer telem-etry are described in chapter 4.

D ATA P R O C E S S I N G AT TRW S Y S T E M S

TRW Systems receives copies of the FR-1400 analog tapes made at the

283



DSS stations for the first 4 days following a launch. As spacecraft primecontractor, TRW Systems was primarily interested in the engineering dataon these tapes. The output from the TRW Systems data-processing line

consists mainly of tabulated engineering data and automatically plottedengineering parameters suitable for assessingspacecraft performance asa function of time (fig. 9-4).

The FR-1400 tapes are first formated at the TRW Systems Data Reduc-tion Center by the Instrumentation Data Handling System (IDHS).Formating prepares them for an IBM 7094, which next edits, sorts, and

Bibliography

I N C L U D E DERE

ARE all engineering reports and papers of significancedealing with the four Pioneer systems: spacecraft, instruments, launchvehicle, and tracking and data acquisition. Because a great deal of Pioneerdescriptive information is available only in the program specifications, acomprehensive list of these follows this bibliography. Volume III includesthe bibliography covering Pioneer operations and scientific results.

ANON.: Pioneer. TRW Spacelog, vol. 4, no. 18 , Fall 1964.ANON.: Pioneer Off-Line Data Processing System-Systems Manual. Computer Sciences

Corp. Rept., 1965.ANON.: Description of the Deep Space Network, Operational Capabilities as of January

1, 1966. JPL Rept. 33-255, 1966.ANON.:Pioneer B Mission, Manned Space Flight Network Post Mission Report. Goddard

Space Flight Center X-552-66-543, 1966.ANON.: Pioneer Systems Capabilities. TRW Systems, 1966.ANON.: Single-Purpose Concept Credited with Pioneer 6 Success. Aviation Week,



Apr. 1966.ANON.: DSIF Development and Operations. JPL Space Programs Summary. JPL Rept.

37-43, vol. IV, Jan. 1967, p. 17; and subsequent issues.ANON.: Final Engineering Report for the M.I.T. Plasma Experiment on Pioneer 6 and

Pioneer 7. MIT Rept., 1967.ANON.: The Pioneer Spacecraft. NASA Facts NF-31, 1967.ANON.: Tracking and Navigational Accuracy Analysis. JPL Space Programs Summary.

JPL Rept. 37-44, vol. III, Mar. 1967, p. 3.ANON.: Communications Systems Research: Sequential Decoding, Supporting Research

and Advanced Development. JPL Rept. N69-16491, 1968,pp. 171-175.ANON.: DSN Capabilities and Plan. Vol. I-Description of Deep Space Network Capa-


ANON.: Technical Manual for Ames Research Center Plasma Probe. Marshall Labora-tories Rept. ML/TN 2200-106.

BALL, J. E.: The Effect of the Gas Leak on the Pioneer VI Orbit. JPL Space ProgramsSummary. JPL Rept. 37-41, vol. III, Sept. 30, 1966, p. 62.

BARON, W. R.: The Solar Array for the Pioneer Deep Space Probe. Electronics andPower, vol. 14, Mar. 1968, p. 105.

BARTLEY, W. C.; MCCRACKEN, K. G.; AN D RAO, U. R.: A Digital System for AccurateTime Sector Division of a Spin-Stabilized Vehicle. IEEE Trans., AES-3, Mar. 1967,p. 230.

-: Pioneer VI Detector to Measure Degree of Anisotropy of Cosmic Radiation inEnergy Range 7.5-90 MeV/Nucleon. Rev. Sci. Inst., vol. 38 , Feb. 1967, p. 266.

B A RT L E Y, W. C.; ET AL.: GRCSW Final Technical Report on Cosmic Ray AnisotropyDetectors Flown on Pioneer VI and VII. Southwest Center for Advanced Studies,

Rept. DASS-66-12, 1966.B A RT L E Y, W. C.; ET AL.: Recent Advances in Cosmic Ray Payloads for Interplanetary

and Planetary Spacecraft. Natl. Telem. Conf., 1968, pp. 238-244.BAUERNSCHUB,J. P., JR.: Nonmagnetic Explosive-Actuated Indexing Device. NASA TN

D-3914, 1967.BERG,O. E.: The Pioneer 8 Cosmic Dust Experiment. NASA TN D-5267, 1969. Also:

Rev. Sci. Inst., vol. 40, Oct. 1969.

BINGHAM,R. C.; ET AL.: Two Satellite-Borne Cosmic Radiation Detectors. IEEE Trans.,NS-13, Feb. 1966, p. 478.

BOLTON, R. C., JR.: The Present Technology in Space Communications Equipment inDeep Space Probes IEEE Conv Rec vol 4 1968 p 586



Deep Space Probes. IEEE Conv. Rec., vol. 4, 1968, p. 586.BRIGGS,D. C.; ET AL.: Silicon Solar Cells for Near-Sun Missions. NASA TM-X-59774,

1967.BUKATA, R. P.; KEATH, E. P.; A N DYOUNSE, J. M.: Final Technical Report on Cosmic

Ray Particle Detectors Designed for Interplanetary Studies Utilizing the Pioneers 8,9, and 10 Spacecrafts. Univ. Texas Rept. N70-13082, 1969.

CANTARANO,S. C.; ET AL.: Magnetic Field Experiment, Pioneers 6, 7, and 8. NASATM-X-63395, 1968.

CHANEY,D.; CURKENDALL,D. W.; AND MOTSCH,R.: Pioneer VI High Frequency DataAnalysis. JPL Space Programs Summary, JPL Rept. 37-41, Sept. 30, 1966, p. 57.

CROOK,G. M.; ET AL.: Instrument Report, TRW Electric Field Detector. TRW Systems

BIBLIOGRAPHY 28 7

GUNN, C. R.: Delta-A Scientific and Application Satellite Launch Vehicle. GoddardSpace Flight Center X-470-69-406, 1969.

HALL, C. F.; N O T H WA N G ,G. J.; AN D H O R N B Y,H.: A Feasibility Study of Solar Probes.IAS Paper 62-21, 1962.

HALL, C. F.; AND MICKELWAIT, A. B.: Development and Management of the Inter-

planetary Pioneer Spacecraft. NASA TM X-60564, 1967. Also in Proc., 18th. Intl.Astronaut. Congr., vol. 2, M. Lunc, ed., Pergamon Press, Oxford, 1968.

HALLUM, C. E.; AN D WEBSTER, P. W.: New Functional Redundancy Approaches forAttitude Control Thruster Systems. NASA CR-73231, 1968.

IUFER,E. J.: Magnetic Program for Pioneer and Its Scientific Payload. Proc. MagneticWorkshop, JPL, 1965.

KEMP, R. F.; AN D SELLEN, J. M., JR.: Solar Wind Simulation with the Pioneer Space-craft. AIAA Paper 66-153, 1966.

KERWIN, W. J.: Solar Wind Measurement Techniques. Part I-An Improved Mag-netometer for Deep Space Use. NASA/ARC Rept.

KOEHLER,R. L.: Interplanetary Electron Content Measured Between Earth and thePioneer VI and VII Spacecraft Using Radio Propagation Effects. Stanford Univ.,1967.

: A Phase Locked Dual Channel Spacecraft Receiver for Phase and Group PathMeasurements. Stanford Univ. Rept. SU-SEL-65-007, 1965.

LINDBERG,R. G., ED.: FeasibilityStudy for Conducting Biological Experiments Aboard aPioneer Spacecraft. NASA CR-73178, 1968.

LUMB, D. R.; AN D H O F M A N ,L. B.: An Efficient Coding System for Deep Space Probeswith Specific Application to Pioneer Missions NASA TN D-4105 1967



with Specific Application to Pioneer Missions. NASA TN D-4105, 1967.LUMB,D. R.: A Study of Codes for Deep Space Telemetry. IEEE Intl. Conf. and Exhibit,

1967.: Test and Preliminary Flight Results on the Sequential Decoding of Convolutional

Encoded Data from Pioneer IX. IEEE ICC Conf. Publ., vol. 5, 1969, pp. 39-1 to39-8.

MASSEY, W. A.: Pioneer 6 Orientation Control System Design Survey: NASA/ERCDesign Criteria Program, Guidance and Control. NASA CR-86193, 1968.

MATTHEWS,H. F.; AN D ERICKSON,M. D.: The NASA Advanced Pioneer Mission. SAEPaper 857D, 1964.

MICKELWAIT, A. B.; AN D SPANGLER,E. R.: Dolling Up a Space Probe. Electronics, vol.


REIFF, G. A.: Interplanetary Spacecraft Telecommunication Systems. IEEE Spectrum,April 1966.

: The Pioneer Spacecraft Program. AAS Paper, 1969.: Evolution of Deep Space Communications-1958 to 1968. IEEE Conv. Rec.,

vol. 4, 1968, p. 570.SCEARCE,C. S.; ET AL.: Magnetic Field Experiment Pioneers 6, 7, and 8. NASA TM-X-

63395, 1968.SCHUMACHER, C. E.: Pioneer Spacecraft Test and Support Program. AIAA Paper 65-247,

1965.SELWEN, H.: Study of Pioneer Spacecraft Oscillator Stability. NASA CR-73307, 1968.SHERGALIS,L. D.: A Magnetically Clean Pioneer. Electronics, vol. 38 , Sept. 20 , 1965,

p. 131.SPANGLER, E. R.; AND T H I E L , A. K.: Temperature Control of Unmanned Satellites

Remote from Earth. VDI Zeitschrift, vol. 108, 1966, p. 712.-: Structural Design of Unmanned Space Vehicles. VDI Zeitschrift, vol. 108, 1966,

p. 573.STAMBLER, I.: The New Pioneer Satellites. Space/Aeronaut., vol. 40, Nov. 1963, p. 58 .STONE, I.: Six-Month Lifetime Predicted for Pioneer. Aviation Week, vol. 80, Jan. 27,

1964, p. 69.THIEL, A. K.: Spacecraft Development in the Next Decade. In Saturn V/Apollo and

Beyond. AAS (Tarzana), 1967.WINDEKNECHT,T. G.: A Simple System for Sun Orientation of a Spinning Satellite.

IAS/ARS Joint Meeting, 1961.



Pioneer Specifications

H E TWO A-SERIESSPECIFICATIONSwere first used during the solicitationof spacecraft and instrument proposals. They were updated frequently

later. Specification A-6669 was updated with almost every modificationof the contract with TRW Systems. P-seriesspecifications were issued earlyin the program and replaced by the PC-series. The PC numbers under 100apply to BlockI; numbers between 100 and 199 apply to BlockII. Specifica-tions were frequently updated and revised.

A-6669 Spacecraft and Associated Ground Equipment (12-1-63)A-7769 Scientific Instrument Specifications (12-31-64)

P-1 Documentation Procedures (1-2-64)

P-2 Amendments to NASA Quality Publications (12-1-63)P-3 Amendment to MSFC-PROC-158B (6-19-64)P-4 Project Development PlanP 5 L h V hi l P f A l i f th Pi P g (12 1 64)



P-5 Launch Vehicle Performance Analysis for the Pioneer Program (12-1-64)P-6 Classification Guide for Project Pioneer (12-15-64)P-8 An Analysis of the Effects of the Spacecraft Systems and TAD Launch Vehicle

on the Pioneer Trajectories (8-1-64)

P-9 Simulator Description (8-12-64)P-10 Pioneer Time Resolution of Telemetry (8-28-64)P-l 1 Procedures for the Preparation and Processing of Range Documentation

(2-15-65)P-12 Pioneer Canberra GOE ARC/STL Acceptance Test Results, July 1, 1965

(7-1-65)P-13 Pioneer Johannesburg GOE ARC/STL Acceptance Test Results, July 1, 1965


PC-021 Spacecraft/Scientific-Instrument Interface SpecificationPC-023 Spacecraft/Launch Vehicle Interface Specification (7-19-65)PC-025 Scientific Instrument Integration Activities (8-1-64)PC-030 GOE Installation, Integration and Compatibility (4-19-65)PC-046 Pioneer Flight Operations (2-19-65)PC-047 Flight Operations Test Plan (8-2-65)PC-050 Procedures for Pioneer-A Flight Operations Test (11-24-65)PC-051 Pioneer A Flight Operations-Detailed Task SequencePC-052 Pioneer A Flight Operations-Detailed Task Sequence for Participating

GroupsPC-053 Pioneer A Flight Operations, Stanford Procedures (11-15-65)PC-054 Pioneer-B Flight OperationsPC-060 Pioneer Off-Line Data Processing System at ARC (8-28-64)PC-062 Pioneer Permissive Command Tape Program (10-20-67)PC-064 Pioneer-6 and 7 DSIF Operational Computer Program Requirement Speci-

fication (5-23-68)PC-070 Pioneer Solar Array Checkout at Table Mountain (PC-070) (renumbered)PC-071 Activities at the Air Force Eastern Test Range (6-19-64)PC-072 Scientific Instrument Integration ActivitiesPC-073 Pioneer Spacecraft/DSS-71 /GOE Compatibility Test Specifications (8-12-65)PC-080 Tests of Scientific Instruments at ARCPC-081 Scientific Instrument Test Requirements (1-29-65)PC-083 Experiment Tests at Malibu Coil Facility-Master Test Procedures (3-31-65)PC-084 Procedure for GOE/DSIF Compatibility Tests (5-18-65)PC 085 Pi S f /DSIF 71/GOE O S d C ibili T d

290



PC-085 Pioneer Spacecraft/DSIF-71/GOE On-Stand Compatibility Test Procedures(5-27-66)

PC-090 Pioneer-A TrajectoriesPC-091 Pioneer-B TrajectoriesPC-092 Pioneer GOE-DSIF User's List (6-14-65)PC-093 Maintenance and Configuration Control (11-1-65)PC-094 Data Format Generator, Type II-Operation and Maintenance Manual

Ground Operational Equipment (5-16-66)PC-III Pioneer Instrument Specification (12-23-64)PC-121 Spacecraft/Scientific-Instrument Interface Specification (1-22-65)PC-122 Spacecraft/Scientific-Instrument Interface Specification (4-5-67)

PIONEER SPECIFICATIONS

PC-166 Pioneer VIII and IX Operational Computer Program Requirement Speci-fication (11-29-69)

PC-167 Pioneer 9 Operational Decoder (5-69)PC-168 TPS/Convolutional Coder Modification Interface Specification (5-20-68)PC-171 Activities at the Air Force Eastern Test Range (5-24-67)PC-173 Pioneer Spacecraft/DSS-71/GOE Compatibility Test Specification (8-9-68)PC-174 ETR/Pioneer

Compatibility Test Plan (7-12-67)PC-180 Tests of Scientific Instruments at ARCPC-181 Scientific Instrument Test Requirements for Systems Tests of Pioneer C/D/EPC-182 SPAC and POLDPS Checkout Magnetic Recordings (9-27-67)PC-183 Convolutional Coder Test RequirementsPC-184 Procedure for GOE/CCU Installation and CheckoutPC-186 ETR/Pioneer Compatibility Test ProcedurePC-187 Spacecraft

Dolly Proof Load ProcedurePC-188 Pioneer VI and VII DSIF Operational Computer Program Test Procedure(4-1-67)

PC-190 Pioneer TrajectoriesPC-191 Pioneer D Trajectory CharacteristicsPC-192 Pioneer E TrajectoryPC-193 GOE/Maintenance and Configuration Control Specification (8-25-67)

PC-194 Pioneer-C RF Equipment and Trajectory InformationPC-195 CCU Description and GOE ModificationsPC-196 Pioneer-D RF Equipment and Trajectory InformationPC-197 Pioneer-E RF Equipment and Trajectory Information (4-18-69)

291



PT-1 Pioneer Trajectory Group Computer Report I, Coordinate TransformationPrograms (8-16-65)

APPENDIX

Reliabilityand QualityAssurance'

ELIABILITY CONTROLSused as a function of program phasing are shown-in figure A-1 with a time line at the top indicating the approximate

number of months on each phase through the first launch and the controlsapplied during each phase. What is considered to be the more significantof these controls will be described in the following three figures.

Figure A-2 describes the organization designed around the majorcontrols with specific tasks listed. A particular individual within thereliability area was assigned to each of these disciplines. In the case ofdesign reviews this individual was required to schedule, define packagecontent, prepare and distribute minutes, see that action items were ac-complished on schedule, and provide the best reviewers possible. In thearea of specifications, the responsibility was for preparation of environ-mental specifications and signoff on all other specifications. Of utmostimportance should be that the callouts in the specifications are realistic,not too tight, but adequate to account for drift.



g , qFor manufacturing and test surveillance two individuals were assigned.

These individuals covered floor problems and attended all problem areameetings held by the functional activity. They were responsible for moni-toring changes that affected their area and had an input as to the dis-position of material that failed. The test surveillance monitor reviewed thetest procedures and reviewed all test setups prior to the start of testing.

In the area of failure reporting, instruction was given as to the use offorms and pickup points where the forms were to be deposited. A failure

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Design ManufctgFailure esignreviewpecifications and teat

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Schedul ing Signoff, a ll Test procedure Forms Weight

Package content (realistic) Test setup Pickup points Power

Minutes Floor Failed partsproblems analysis

Action i t ems Prepare Change Appropriate fix Simplicity

accompl ished environm enta l control board Failure review Proven~~~~~Reviewer Material review board

selection board Interfaces

Materialsan d Parts

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Specifications SpecificationsWelding and Vendor selectionpotting an d monitoringschedulesProcedures Specia ltestsSpecial tests Burn-in data

evaluationTraining Magnetically

cleanDerating criteria



Derating criteria

F IGURE A-2.-Reliability organization and functions.

and knowledge of interface parameters. Types of redundancy were eval-uated, with consideration continually being given to weight and power.

Materials and processes personnel provided specifications and weldingand potting schedules. They were continually involved in special testingand training when problem areas occurred.


TABLE A- l.-Signficant Reliability Elements (Design)

GeneralPart specificationsbased on Minuteman criteriaReliability engineerassignedto each major designareaRedundancy selectedon greatest improvement/lb

DesignreviewsKept small (12 to 25 knowledgeableengineers,including NASA)Data packages,conciseand earlySeparate reviewsfor drawings (producibility)

High reliability parts program100 percent burn-in (100 to 250 hr)Parameter drift screening

100 percent environmental and life test samplingNo new types,processesor production linesPart typeslimited (Parts Deviation Board)

TABLE A-2.-Signicant Reliability Elements (Manufacturing and Test)

GeneralReliability surveillanceduring manufacturing and testEnvironmental test evaluation criteria established fo r each equipment

296



estab s ed o each equipmentFailure evaluation system

Failure reporting forms simple, drop stations convenientCauseof failure definedrapidly (dissection,X-ray, etc.)Determine and implement correctiveactionConcurrency by Failure ReviewBoard (TRW and NASA participants)

Test philosophyEquipment tests--development, life, qualification, acceptanceSystemtests-development, qualification, integration, acceptance (space simulation)Maximum test time possible (minimum 1000 hr/system)

APPENDIX

particular discipline contributed to seeing that the producibility of theequipment was at an acceptable level.

The high reliability parts program was considered of major importanceand contained 100 percent burn-in, that is , burn-in of all parts from 100 to250 hr depending on the part type. It also included parameter drift screen-

ing; i.e., monitoring the drift of various parameters over the interval thepart was in burn-in, and 100 percent environmental and life test sampling.No new types of parts were allowed with one exception which will be dis-cussed later. Table A-2 describes the significant reliability elements asso-ciated with the manufacturing and test phase of the program. Manu-facturing and test surveillance was provided by two reliability engineers.This consisted primarily of assembly and test setup procedure evaluation,evaluation of test procedures to determine that the environmental testcriteria fo r each equipment was correctly assigned to disclose design weak-nesses, participation in the disposition of failed material, and control ofchanges as well as assignment of test requirements following a failure.Preship and prelaunch evaluation of flight scheduled equipment consistedof reviewing each equipment's history carefully before it was assigned toflight status. In the area of failure evaluation, failure report forms were keptsimple and drop stations were located convenient to the manufacturing

297



and test stations. Failure definition was determined expeditiously (within2 days normally) by means of dissection, x-ray, etc. These results were usedto assign fixesand served the Failure Review Board in making their decision.

The Pioneer program test philosophy was very simple. The equipmentwas made to operate; therefore let it operate as much as practical. Aminimum of 1000 operating hours were accrued per system prior to eachlaunch. On occasion at the Cape, when the launch was postponed forlaunch vehicle reasons, the spacecraft was turned on and left operating in

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TABLE A-3.-Post-launch Anomalies

Six anomalies noted in eleven spacecraft years of operationOne anomaly degraded mission

Cause-New innovation device (PSCR Sun sensor)Threshold sensitive to ultraviolet radiation

Experiment viewing direction lost on Pioneer 7 (two experiments of little value)Fix (ultraviolet filter) included on Pioneer 9

Five anomalies produced no mission degradation becauseRedundancy available by ground command (1)Self-heal (2 )Function not required following orientation (2)

probably the test sets did not satisfactorily simulate interface conditions.Another major fault of the reliability program can therefore be assigned tonot including test equipment in the design review program. The final chart,table A-3 in the Pioneer A-E discussion, showsthat there were sixanomalies

noted in 1 1 spacecraft years of operation. Only one of these failures causeddegradation to the mission, and this was because of the use of an unprovendevice, something that a good reliability engineer should avoid if at allpossible. The reason for using this device, a photo silicon control rectifier,

299



p g , p ,as a Sun sensor was to save approximately three-quarters of a lb in weight.The viewing direction during spin has been lost on Pioneer 7 which makes

two of the experiments of little value. A study subsequently revealed thethreshold of this device to be sensitive to ultraviolet radiation and since afix could not be implemented until Pioneer 9 it is anticipated that Pioneers6 and 8 will fail in a like manner. Five other anomalies have been noted,none of which have degraded the mission. It is entirely possible that theself-heal anomalies will reoccur again as failures.

Index

Aerojet-General Corp., 213, 215 design, 91-100Allegheny Ballistics Laboratory, 213, 215 functions, 7Ames magnetometer, 136, 149 Communication subsystem, antennas, 12-

design, 145-148 13, 55, 57, 66-68telemetry, 76-78, 86 bit rates, 40, 57

Ames plasma probe, 136 design, 55-68, 135design, 152-160 distance capabilities, 71

telemetry, 76-78, 86 , 157, 158 functions, 7Ames Research Center, 2, 8, 9,11, 27 , 55, interfaces, 58

95, 139, 141, 146, 155, 165, 181, 192, phase-lock operation, 59-60, 61, 63241, 245, 247, 254, 255, 256, 259, 260, power budget, 62, 64

293



266, 273, 277, 278, 280,282, 283, 293 receiver, 61-65(See also Ames magnetometer, Ames reliability, 14-46plasma probe) transmitter, 65-66

Anderson, J., 258 Computer Sciences Corp., 277Antennas, spacecraft, 12-13, 55, 57, 66- Constraints, system, 3, 39

69, 132, 133, 194 Convolutional Coder Unit (CCU). SeeAntennas, tracking. See Deep Space Data handling subsystem.

Network. Cosmic dust detector. See Goddard cosmicAstrodata, Inc., 27 7 dust experiment.

Cosmic ray experiments. See Chicago

INDEX

reliability, 46 , 73word structure, 74 , 91

Data processing, 259, 277-283Data Storage Unit (DSU). See Data

handling subsystem.Deep Space Instrumentation Facility

(DSIF). See Deep Space Network.Deep Space Network (DSN), 3, 9, 11, 15,

40, 41, 55, 57, 58 , 61, 63, 64, 65, 6 7,70 , 85, 87, 90, 93, 95, 115, 116, 176,191, 194, 277, 278

antennas, 245design, 237-273DSIF, 241-245station configurations for Pioneer, 264,

274-275station list, 240tracking requirements, 21, 253-263tracking functions, 59-60, 239, 242-243(See also Space Flight Operations

Facility)Delta launch vehicle, 3, 9, 11, 14, 40, 41,

52 , 58 , 103, 127,191, 239, 240, 241configurations, 215, 218description, 211-236

Environmental Science Service Adminis-tration (ESSA), 2, 139, 277

ESSA. See Environmental Science ServiceAdministration.

Experiments. See specific experiments.

Feasibility study, 8-11, 14, 15-17, 18, 39,44, 47, 48, 49, 90 , 103, 129

Goddard cosmic dust experiment, design,135, 181-188

selection, 141-142telemetry, 76-78, 86 , 186, 187, 18 8

Goddardmagnetometer, design, 136,142-145, 146

telemetry, 76-78, 86 , 143Goddard Space Flight Center (GSFC),

74 , 217, 265(See also Delta launch vehicle, Goddard

cosmic dust experiment, Goddard

magnetometer)GOE. See Ground Operational Equip-ment.

GRCSW cosmic ray experiment, design,136 165-170

302



p ,interfaces, 217-225launch requirements, 21

launch sequence, 231-235Pioneer-9 launch description, 27-29trajectory design, 225-229, 230, 232weight breakdown, 235-236

Digital Telemetry Unit (DTU). See DataHandling subsystem.

DSIF. See Deep Space Network.

136, 165 170telemetry, 76-78, 86 , 168, 170

Ground Operational Equipment (GOE),

191, 253, 260, 266-271, 279

Honeywell, Inc., 146, 148Hughes Aircraft, 66

IGY. See International Geophysical Year.Instruments, design, 137-190

INDEX

Launch requirements, 14, 26-27Launch vehicle. See Delta launch vehicle.Launch windows, 26-29Lifetime, spacecraft, 14, 40

(See also Reliability)

Magnetic cleanliness, 7, 39, 40, 42, 44,47-49, 144-145

tests for, 196, 197-198, 200, 204Magnetometers, 48 , 49, 129

(See also Ames magnetometer, Goddardmagnetometer)

Manned Space Flight Network, 237, 239,

240, 253, 257, 263, 264McDonnell-Douglas Astronautics Co., 27 ,

211, 213, 215Minnesota cosmic ray detector, design,

136, 170-171, 172, 173telemetry, 76-78, 86

Mission objectives, 1-3, 19-20

MIT Faraday-cup plasma probe, design,101, 136, 148-153

telemetry, 76-78, 86 , 100, 109, 150, 152

303

Philco-Ford, 146, 148Pioneer mission operations center, 245,

247, 259, 260, 271Pioneer A. See Pioneer 6.Pioneer B. See Pioneer 7.Pioneer C . See Pioneer 8.

Pioneer D. See Pioneer 9.Pioneer E, 23, 52, 142

launch vehicle, 21 7scientific objectives, 20, 255, 256, 265tracking requirements, 253

Pioneers 1-4, 211, 237, 241Pioneer 5, 12, 211, 237, 241

Pioneer 6, 52, 101, 130, 299gas leak, 177launch trajectory, 230launch vehicle, 217scientific objectives, 19, 256test history, 206-207

Pioneer 7, 130, 138, 29 9

gas leak, 117launch, 22 , 21 2launch vehicle, 217scientific objectives, 19-20, 255, 256

Pioneer 8 138 266 271 272 29 9



NASA Communications Network(NASCOM), 93 , 240, 245-251, 263

NASA Headquarters, 8, 181, 241, 253NASCOM. See NASA Communications

Network.Naval Propellant Plant, 215Naval Research Laboratory (NRL), 237,

238North American Rockwell, 213

Pioneer 8, 138, 266, 271, 272, 29 9launch vehicle, 217scientific objectives, 20, 255, 256

Pioneer 9, 102, 283, 299launch sequence, 231-235launch vehicle, 217scientific objectives, 20, 256trajectory analysis, 25 , 27-34

Plasma probes. See Ames plasma probe,MIT Faraday-cup plasma probe.

INDEX

Solar array. See Electric power subsystem.Spacecraft design philosophy, 3, 39-44Spacecraft evolution, 48-53Space disturbance forecasts, 2, 34, 139,

277Space Flight Operations Facility (SFOF),

3, 176, 240, 241, 242, 256,258, 263,265, 266, 267, 273,277

Space Science Steering Committee, 17,141

Space Technology Laboratories (STL).See TRW Systems.

Stabilization, 12, 41, 129, 132

(See also Orientation subsystem)Stanford radio propagation experiment,

13, 80, 87, 88 , 132, 133, 255, 277design, 136, 171-176, 177, 179telemetry, 76-78, 86

Structure subsystem, booms, 131, 132,133, 134

design, 135, 128-136dimensions, 129, 132functions, 7reliability, 13 4

STL See TRW Systems

Test and Training Satellite (TETR), 21 ,41, 53, 227

Texas Instruments, 106Thermal control subsystem, design, 15,

121-128, 13 5functions, 7

requirements, 121-122Thiokol Chemical Corp., 213, 215Tracking. See Deep Space Network.Trajectory, constraints, 20-21, 225-229,

230, 232design, 20launch phase,.21-22, 225-229, 230, 23 2

Pioneer 9, 27-34TRW Systems, 8 , 11, 12, 42, 45, 50, 51,

63, 65, 66, 85,90 , 100, 103, 107, 121,122, 123, 125, 126, 134, 191, 192, 196,198, 199,277, 278, 282, 283, 293, 296

TRW Systems electric field detector,design, 136, 176-181

telemetry, 76-78, 86 , 176, 179Type I and Type II maneuvers. See

Orientation.

304



STL. See TRW Systems.Sun sensor, degradation, 49, 114

design, 114-117, 118, 119, 121, 127functions, 143

System and subsystem definition, 3-8, 9,55, 56

Telemetry. See Communication sub-system.

Test program 44 134 191-210

Unified S-Band, 238United Technology Center, 213, 215

U.S. Air Force, 239, 263,264, 265U.S. Army, 238

Watkins Johnson, 66Weight, spacecraft, 23 , 39, 40, 41

detailed breakdown, 134-136evolution 48 49

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Documents

The Interplanetary Pioneers. Volume 2 System Design and Development