the small astronomy satellite (sas) program

8/7/2019 The Small Astronomy Satellite (SAS) Program

1/25

N A S A T E C H N I C A L N O T E

THE SMALL ASTRONOMYSATELLITE (SAS) PROGRAM

by Marjorie Re TownsendGoddard Space Flight CenterGreenbelt, M d .

N A S A LN D-5099I*

LOAN COPY: RETURN TOAFWL (WLIL-2)

KIRTLAND AFB, N MEX

N A T I O N A L A E R O N A U T I C S A N D S PA C E A D M I N I S T R A T I O N W A S H I N G T O N , D . C . M A R C H 1 9 6 9


2/25

TECH LIBRARY KAFB,NM

0131788NASA TN D-5099

THE SMALL ASTRONOMY SAT ELL ITE (SAS) PROGRAM

By Mar jo r i e R. TownsendGoddard Space Flight Center

Greenbelt, Md.

NAT ION AL AERONAUTICS AND SPACE ADMINISTRATIONFor sale b y the Clearinghouse for Feder al Scientif ic and T echnic al Information

Springfield, Virginia 22151 - CFSTI price $3 .00


3/25

ABSTRACT The Small Astronomy Satell ite (SAS), one of NASA's

newest Explo rer-c lass pro grams , can provide at relativelylow cost much previously unavailable ba si c information onlow and high energy radiation fro m so ur ce s both inside andoutside the galaxy. This rep or t gives sufficient informationconcerning the SAS volume and weight, and the power, command, telemetry, and control syst ems to per mit the rea de rto ascertain its suitability for specific astronomy experiments. Pro je ct reliability requirements and environmentaltest conditions are included in the appendixes.


4/25

FOREWORDOur cur ren t decade has seen many remarkable advancements in astrophysics.

Some of the most outstanding are the discovery of over two dozen x-ray so urce sand the detection of gamma radiation from the galactic center. These findingsrepresent a majo r st ep toward the goal of understanding the high energyphenomenain ou r galaxy and those which govern the principal physical pr oc es se s of theuniverse.

It is now time fo r deta iled stud ies of the phenomena alre ady discovered, formuch more thorough sur vey s of the sky to determine the numbers and types of obje ct s emitting high energy photons, and for intensive studies of c eles tial objects ofparticula r inte rest . Supernovae, for example, appear to be the most catastrophicevents within galaxies and are al so thought to be the origin of cosmic rays. Theydeserve examination in the ultraviolet, X-ray, and gamma-ra y s pec tra l regions toprovide a fuller understanding of such processes as nuclear synthesis and possiblemechan isms of cosmic ra y accele ration. Further, high energy photon stud iesshould provide some importan t piec es i n the puzzle of dynamic balance among thevari ous fo rces in our galaxy, including those of gravity, the cosmic ray gas, magnetic fields, and the kinetic motion of i nt er st el la r matt er. The resolut ion of thi sproblem, together with the determination of the composition of in te rs te ll ar matter ,is al so fundamentally important in the study of star formation. Probing further intothe universe for objec ts and phenomena still unknown to u s is yet another very important facet of high energy astronomy; it should be remembere d tha t none of thecelestial phenomena known before the 1960's led astr onom ers even to predict theexistence of celes tial x-ray source s.

The principal a im of the Small Astronomy Satellite (SAS) pro gra m is to provide an Explorer-class s a t e 11i t e capable of s up p o r t i n g experiments inhigh energy astronom y, and thereby providing one means of accomplishing the goalsjust outlined. Designed for launching by the Scout vehicle, SAS is intended particularly to f i l l the gap between very small or short-duration expe riments and thevery lar ge ones to be flown on observato ries and lar ge spacecraft. This documentdescribes the SAS satellite and its general capabilities s o that the potential u se rmay determine whether it is compatible with his needs.

Carl FichtelProje ct Scientist

iii


5/25

THE SMALL ASTRONOMY SATELLITE (SAS) PROGRAMby

Marjorie R. Townsend S A S Proj ect Manager

Goddard Space Fl ight Center

INTRODUCTIONThe object ive of the SAS program is to study the cel estial sphe re above the eart h's atmosphere

and to sea rch for s ourc es radiating in the X-ray, Gamma ray, ultraviolet, visible, and infraredregions of the spe ctru m both inside and outside of our galaxy. Surveys per form ed from such asmall, Scout-launched satellite (Figure 1)w i l l provide valuable data on the position, strength, and

Figure 1-Artist 's conc eption of the Small Astronomy Satell i te (SAS): experiment, above; spacecraft, below.1


6/25

time variation of sour ces radiating in these regio ns of the spectrum and lead t o the selection ofthe mo re i nteres ting ones for detailed study by later missions on this o r more sophisticated spacecraft. It is anticipated that , beginning in 1970, one or two of these satellites will be launched eachyear (with an estima ted lifetime of a year, or as limited by the experiment sensors).

A nominally cir cu la r 300 nautical mile (555 kilometer) equatorial orb it is planned for theSAS series. The equatorial orbit is highly desirabl e fo r an astronomy mission, because thesatellite will mi ss the South Atlantic anomaly where the radiation belt s com e fa r down into theearth's atmosphere (Figure 2). Th is will avoid degradation of the spacecraft operation andminimize the background count which can affect the data fro m seve ral different types of se ns or sapplicable to satellite astronomy. Thi s orbit, with an incl ination of 2.9 degrees, can be achievedby launching fr om the San Marco platform located three miles east of Kenya in the Indian Ocean.The maximum payload capability under these conditions is 333 pounds.

The prime contractor for SAS is the Applied Physics Laboratory of the Johns HopkinsUniversity, which is responsible fo r developingthe spacecraft and for integrating and testingeach spacecraf t and experiment.

BASIC SPACECRAFT DESIGNThe satellite is designed to have a clean

interface between the basic spacecraft s tru ctureand the experim ent payload (Figure 3). This isintended to minimize costs for follow-on programs. The basic spacecraft is a cylinder approximately 22 inches in diameter and 20 inch eshigh, weighing about 180 pounds. The light-fo r the experiment. Four sol ar paddles, hingedto the outer shell, provide raw power to thespacec raft and experiment. At the tips of thepaddles are the command and telemetry =tennas. Inside the shell (F igure 4), and thermallyisolated from it, is a honeycomb deck which is

NORTH GEOGRAPHIC LATITUDE SOUTH

Figure 2-Longitudin ally averaged map of the earth'srad iat ion belts. The isolines refer to omnidirectionalfluxes (cm-2-sec-1) of elec trons wit h energies > 0.5 MeV.

itself a good thermal conductor. On it are the sys tems to provide the basic spacecraft functions:a rechargeable battery with its charge control and regulator systems; command receivers and decoders; a 1000 bit pe r second teleme try system with one-orbit storag e capability on tape; and amagnetically torqued commandable control system which can point the spacecraft stably to anypoint in the sky. The electronic circ uitry is packaged in "books" mounted, with the nutation damper ,on the underside of the honeycomb deck. The battery, tape r ecord er , transm itt er , command receivers, and roto r are mounted on the top side.2


7/25

STA. 25.00

I I I I l l ,STA. 18.5 1 AFT EDGE -INST RUMEN T PACKAGE:J4 STA. 19.89 Q UMBILICA L CONN. & A/C DOORSTA. 21.45 Q HINGE

, , I ,

CONTROL PACKAGE \ STA. 27.00 FWD. EDGE- SOLAR PANELSlrri I 1 I 'SPACECRAFT ADAPTER

4TH STAGEFW-4 ROCKET MOTOR

1 1 1 STA. 38.39 SEPARATION PLANESTA. 43.21 ORDNANCE ACCESSIfi----STA. 47.77 ROCKET MOTOR INTERFACE% STA. 52.00 Q DESPIN CABLES & FWD.

STA. 81.90 AFT EDGE- SOLAR PANELSSTA. 84 .5 2 ROCKET MOTOR/"D" SECTIONINTERFACESTA. 85.40 ANTENNA ELEMENTS

Figure 3-SAS in launch configuration on Scout with standard shroud.


8/25

STA 18.89-- STA 21.45

HEATSHI ELD

Figure 4-Cross-section of basicspacecraft structure.

STA 38.39ORDNANCEe ACCESS-STA 43.21

STA 47.77

SOLAR PANEL hPAYLOAD ENVELOPEOne of the SAS design requireme nts was the ability to opera te in any orientation, thus com

plicating the therma l design as well as that of the power system. The sur fac e of the cylinder is theonly pa rt of th e spacecra ft portion which radiates the intern al heatloads to space. Since the experiment is isolated fro m the spacec raft section with a high the rmal resistance, we will consider thespacecr aft section independently of the experiment section. The most critic al items are the taperec ord er and the battery, which are co-located for efficient th erm al design. The judicious use ofpassive thermal coatings and multilayer insulation, with a smal l amount of act ive therm al controlin the form of heate rs, pe rm it s operation in any random orientation. While most of the spacecraftcomponents can be operated satisfactorily from -20" C to +70" C, th e battery prefer s a temperaturebetween 0" and +30"C. A sepa rate arr ay of sola r cell s on the base of the spacecraft providesheate r power for the experiment in the worst-case cold condition.

POWER SYSTEMThe power sys tem co nsists of: four paddles with sol ar c el ls on both si des ; a 6-ampere-hour

8-cell nickel-cadmium battery; and a shunt regulator. Still assuming a random orientation, thepower system mu st provide 27 watts average power over the entire orbit, both night and day. Thisis particularly important i n astronomy experiments, which frequently get their most useful dataat night, when scat te red sunlight cannot interf ere with the sen sor s. The basic spacec raft functionsrequire 17 watts average power, allowing 10 watts for the experiment. Figure 5 shows the variation4


9/25

--in available power as a function of sun angle.The total power includes operational needs plus 10 ,---enough to charge the battery to provide the VI / \ /*/c-- x MAXIMUME 8 - // \', ; \\INSTANTANEOUSsame power through the dark pa rt of the orbit. Wa / \\ I \

A block diagra m of the power subsystem isv I-'g6

6 '/

\ I

d ,' AVERAGEshown in Figure 6. The solar cell ar ra y is con- w 4 - \ \ ,------\cc \ I \nected in two separate sections, the main arra y LK \ / \ /3and an auxiliary array. The auxiliary array 0 2- \\ '\\ / / \\I I,'/iMINIMUMINSTANTANEOUSprovides redundant power for the command sub- I ' d ' , I , 'b,: Isystem, and trickle-ch arges the batteryif it hasbeen disconnected f rom the main bus by the lowvoltage sensing switch.

Figure 5-SAS solar cell array current.

C O N V E R T E R C O N V E R T E R

L O WV O L T A G E TEMP.S E N S I N G C O N T R O LC I R C U I T- 4 I

I M I T E R

Figure 6-SAS power system.The 6-ampere-hour nickel-cadmium ba ttery is protected by two redundant charge-control

devices, a coulometer and an enhanced zener.The coulometer s en se s the total energy taken from and returned t o the battery, by integrating

the voltage developed acr os s the battery telemetry resi stor . When the energy taken from the5


10/25

batter y on discharge has been returned, the coulometer reduce s the battery charge cu rrent to a"safe", pre set trickle-ch arge level by shunting the remainder. (This unwanted cur ren t is shuntedinternally if the batter y is cold, or externally if it is hot.)

The enhanced zener is a voltage limiter. It actuates a shunt, d ecreasi ng the battery chargecurren t, when the battery voltage reac hes a pres et level. Thus the charge curren t decr ease s toa level which makes the battery voltage equal to the enhanced zen er limi t voltage. As with thecoulometer the shunted curr en t may be dissipated internally or externally, depending on batterytemperature . The power subsy stem normally operate s with both the coulometer and the enhancedzener. However, it can operate with either one alone.

The low voltage sensing switch continuously monitors the voltage of the main bus. If it dropsbelow 8.8 volts (normal voltage is 10.7 volts), power is switched from the main converters and thebatte ry to prevent damage due to possible shorted loads. In th is "solar only" mode, the batte rywill be trickle-charged by the auxiliary array , while the only loads on the main array ar e the dualcommand sys tems and the heate r circui ts. Each circui t can be switched out by command i f it becomes defective.

COMMAND SYSTEMSAS employs the NASA standard PCM command syst em and car ri e s redundant receiver s and

decoders. The redundancy continues through the relay coils themselves, but a single set of contacts is used. Figure 7 shows the. command system. Independent dipole antennas are mounted onthe ti ps of two adjacent sol ar paddles. The command signal, a pulse code modulated, frequency-shift keyed, amplitude modulated (PCM/FSK-AM/AM) signal , is demodulated by e ither o r bothcommand rece iv er s and app ears at the video output. Bit synchronization is obtained from a signalthat has been amplitude-modulated onto the data subcar rie r. The video signal is sent to a commandecoder, which te st s the tones for authenticity by examining the timing and frequencies. If thetones a r e proper, the decoder converts them to digital levels and sends the levels and appropriatel

timed clocking pulses t o the logic circuitry.D A T A

D C / D C C O M M A N D SS I C EXP. TIMER por tions of the logic , beginning with the first-. correc tly received tone bu rst and remaining on

long enough to complete the command. Thelogic c ircuit ry accumulates the decoded levelsin a shift register and, if the satellite addressbits ar e proper, it activates the proper row andcolumn of the re lay switching mat rix to switchthe desired relay or relay cluster.

4T-E A TM The decoder a l s o applie s power to the standby

From the redundant interconnections it issee n that only one rece iver, one command logic,

Figure 7-SAS command system. and one relay switching matrix need be operatin6


11/25

t o send commands. One bit i n the commandword (Figure 8 ) addresses the particular logiccircuit and one bit the particula r r ela y switching circuit t o be used, so tha t only one logic andone matrix are activated during any given command. Each half-system h as its own power converte r. The sys tem' s magnetic latching rel ayseach have two coils, one for ei ther sys tem half,and consume power only during a command. Thedirection of the current in any given coil determines the state in which th e rel ay will latch.Because of these dual coils, actually located in

1 15 16 17 23 24 25 2627 32 33

' w T E L ! T E LO bIC IMATRIX DESTINATIONSYNC ADDRESS COM MAND BITSTYPE50 56 57 58

Figure 8-SAS command word.the rel ay switching matrix, the sys tem 's redundancy continues up to the magnetic f l u x circuiti n the relay pole piece, and no single relay coil open or short can disable the system.

The command sys tem provides SAS with 35 on-off rela y commands: 25 fo r the spacecr aftand 10 fo r the experiment. In addition, i f bit 25 in the command word specifie s a "data command"instead of a "relay command", then bit s 33 through 56 are routed, as specified by bits 27 through32, to a location other than the relay m atr ix fo r separate decoding. Th is per mit s an experimenterto increase his allocation of 10 on-off commands by decoding up to 2 2 4 within the experiment.

TELEMETRY SYSTEMBecause of the low equatoria l orb it chosen for SAS, the spacecra ft can be se en regularly by

only one of the NASA Satelli te Tracking and Data Acquisition Network (STADAN) sta tions. Forthis reason , the spacec raft tele metry subsystem (Figure 9) mus t include on-board sto rage of thedata collected during the remaining 90 perc ent of the orbit. Therefo re, the pulse code modulated/phase modulated (PCM/PM) tele metry system ha s two basic modes of operation-record andplayback. In the normal re cor d mode, digital data from various sources are multiplexed and sentto be stored serially in Manchester Code on an endless loop tape re co rd er and to phase-modulatea VHF transmi tter . Thi s continuous output, radiated through a turnstile antenna at th e tip of asol ar paddle, is used as a tracking beacon and to collect real-time data. A s the satellite passesover its data acquisition station, however, the tape re cor der is commanded into playback modeand the data from the full 96-minute orbit phase-modulate the same t ran sm itt er with a 30 kbpssignal. Simultaneously, the transmitter is switched into its high power (2-watt) mode. The re cord er and transmit ter r eturn to the r ecor d mode automatically after completion of the 3.4 minuteplayback cycle, o r upon ground command. In a special operating mode, real-time data can betransmitted at 2 watts. Th is sys tem me et s the requ irements of the NASA/GSFC Aerospa ce DataSystems Standards fo r PCM tele metry systems. The maximum phase deviation is *64", with area l-t ime bandwidth allocat ion of &15kHz and a tape re co rd er playback alloca tion of *45 kHz.Linear phase filters, having their 3-db points at 1 kHz and 30 kHz respectively for the record andplayback modes, are used f or premodulation filtering.

7


12/25

------

THERMALINPUTS SUB-COMMUTATOR I JI It

MAINANALOGCOMMUTATOR CONVERTER COMMANDL TIME DATAPOWER1LO POWER

COMMAND

COMMAND EXPERIMENT t3 MULTI&PLWER TRANSMITTERINP S ENCODER TRANSMITTERCRAM DAMPERINPUTS MEASUREMENTREADOUT GATES INPUTS

MAJOR D S C - PRE-MODUIATI ONNO. 1__------ COUNTER~

I FRAME I I IDENTIFICATION 1 SI C TAPE RECORDERI ---T----1J 1 TELLTALES 1 PRE-MODULATI ONFILTERTOTAPERECORDER 7-6 x IO A B IT

COMMANDRECORDlPMYBACK-WHEEL S P E D EXPERIMENT MONITOR AUX CMDVERIF INPUTS

Figure 9-SAS telemetry system.

The real-tim e data rate of 1000 bps was det ermined by optimizing exper imente r and housekeeping data requirements with a reasonable length of tap e (300 fee t) and a reasonably conservative bit packing density for single trac k recording (1667 bit s per inch). The tape re co rd er , of flight-prov en end less loop design, weighs only 7 pounds and requi res 1.5 watts in re cor d mode and 3.0wa tt s in playback mode. Wow and flutt er will be kept below 2 percent peak-to-peak, with jitter below 3 perc ent peak-to-peak. Commands will bypass the output reclocking circu itr y and energizeth e erase head, as well as provide the obvious record/playback and on/off cont rol. Digital datarecording was selected for its better signal-to-noise properties and for the ease of providingdigital data to the telemet ry system.

The basic forma t is shown in Figure 10. The minor fra me is composed of 32 words, e achwith three 8-bit syllables, an arra ngemen t that per mi ts easy handling of 8 bit, 1 6 bit, o r 24 bitdata words. Figure 11 shows the breakdown of a minor frame. Of th es e 96 syllab les , the firstthree are allocated to fr am e synchronization. The fr am e synchronization word used is the optimum 24-bit code word determined by Maury and Styles1, 111110101111001100100000. One syllableI J e s s e L . Maury, Jr., and Frederick J . Styles, Development of Optimum Frame Synchronization for Goddard Space Flight Center PCM

Telemetry Standards, Proceedings Na t io na l Te leme try Co nference . June 1964, Los Angeles, California.8


13/25

WORD

15 II FRAME IDENTi

78 II9

SYLLABLE 1 I SYLLABLE 2 SYLLABLE 3I FRAME SYNC23

10I 1I213 i41516 1 1 M U M I FRAME = 64 MINOR FRAMB 17 DSC 'I 1 DSC '2 EXP. HSKPG.181 MINM) FRAME = 32WORDS I9 I1 S Y M L E :8 BmS BINARY 20

1 WORD = 3 SVUAELES 21 I PARITY CHECK2223

II 4igure 10-SAS PCM telemetry system frame format. 242526 27 28is allocated to fram e identification, thre e to 29 I064-position analog subcommutators, two to dig- 31ita l su bcommutator s, one for a parity word, and n I

the remaining 86 for expe rimen t data. The Figure 11-Telemetry minor frame format. period of the minor fr am e is 0.768 second, and of the major frame is 64 times that, or 49.152 seconds. The basic timing source for the system is a temperature-com pensated 1.024 M H z crystal oscillator whose frequency variation is less than one par t i n l o 6 per orbit. It se rve s as th e source for all readout ga tes, clock sign als to contr ol the multiplexing of spacecraft and experiment data, motor drive frequenc ies for the tape rec ord er and rotor, and pulses for th e 20-bit accumulator, which is incremented by the minor f ram e identifier eve ry major f ram e, and sampled every eight minor fram es. Unambiguous time co rrela tion for about 20 months is provided.

While mo st of the data orig inate in digital fo rm, many of the housekeeping functions are analogand mus t be converted. A dual-slope integrating analog-to-digital conv erter , acc urate to 8 bits,is used for this purpose. It is clocked at 16 kHz and has an aper tu re of 8 msec. While its normalinputs are 0 to 250 mv, the tele metry syst em can provide a varie ty of a tten uator s so that signallev els of *0.5, &1.25, *2.5, *5.0, or *7.5 vol ts can be accepted. An additional fe at ur e of the analogsubcommutators is that it is possible to command ei ther of them to stop and samp le continuously-at the minor frame rate of 0.768 second-any one of its 64 positions. After receiv ing the RUNcommand again, it will automatically resynchronize itself at the beginning of the next major frame.

One of the dig ital subcom muta tors has 16 channels. Twelve of thes e accept 8-bit paral le l data,while the othe r four accept 8-bit serial data. The second digital subcommutator ha s 8 channels,each of which a cce pts 8-bit serial data. These digital subcommutators provide command verification, the unambiguous time code, a monitor of the r otor speed, and time corre lati on to 24 mse cfor the digital solar aspect sensor data.

9


14/25

This telemetry system occupies less than 0.2 cubic foot, weighs 14 pounds, and consumes4.5watts. It is built in modular f orm for easy expansion, and the design is such that the presenthard-wired forma t could be replaced by an in-flight programmable unit.

CONTROL SYSTEMThe most unique fea tur e of the Small Astronomy Satellit e is its control system. An evolution

ary system is planned which will per mit the present scanning system to prog re ss to an accuratelypointed, three-axis-stabilized spacecra ft-a mos t valuable tool in astronomy.

The basic control system use s the earth's magnetic field for torquing the satellite. An obviousadvantage of this is the elimination of the need fo r expendables and th ei r inheren t life limitations.

Experiments can look along the Z - a x i s of the SAS spacecraft or perpendicular t o it. If anexperiment looks out perpendicul ar to the Z - a x i s , and the spacecra ft is rota ted slowly about theZ-axis, then a swath is swept out of the celesti al sphere. The vec tor made by the Z - a x i s in spacecan be moved, and another swath swept. Thus, ultimately, the en tire celes tial spher e can bescanned.

Thi s sweeping technique is valid even if th e Z-axis wobbles, but data reduction is much easieri f the Z-axis is stable. One way of achieving stabil ity is by spinning the spacecraft. This is quitesatisfactory fo r many experiments looking along the Z- a x i s . However, a high spin rate will reduce the effective time-on-target fo r perpendicular-looking expe rimen ts and, consequently, thesignal-to-noise rat io of the cel est ial sou rces. An alte rnat ive method of adding the requiredangular momentum along the Z - a x i s is to include a momentum wheel whose axis is parallel to thez-axis. -AGING MECHREGULATEDFREQUENCYINVERTER

M x A O < M x c +ZOOCHARGEABLE MAGNET SYSTEM

CAGING MECHEFFECTIVESPRING RATEADJUST TELEMETRY

NUTATION DAMPING SYSTEM

Figure 12-SAS-A spin-axis orientation control system.

Figure 12 shows the vario us components ofthe control system. Z - a x i s orientation is controlled by energizing a torquing coil. Thiselectromagnet, acting like a compass needle,attempts to align the spacecraft with the earth'smagnetic field. At the maximum rate , using amagnetic dipole of 5 x io4 pole-centimeters,the Z - a x i s of the spacecra ft c an be moved 1.72deg ree s per minute.

Figure 13 shows how a maneuver might beperformed to point the Z - a x i s to a new directionas requested by an experimenter. This particular sequence is designed to r es ul t in no changein righ t ascension , but only in declination. Innormal operation, a combined motion i n bothright ascension and declination will be

10


15/25

60 : tI I

I IDESIRED TERM INUS

p 301- DELAY TIME AFTER RISEa FOR 2nd MANEUVER (minutes)

AND START 2n d0: -10 0 i o i o i o i o i o BO i o i o i o 100 i i o i i o A o

RIGHT ASCENSION (degree s)Figure 13-SAS-A declination maneuver sequence.

accomplished with the a id of a computer. The computer mus t sto re the values of the ea rth' s magnetic field in each predicted satel lite position. With thi s set of values , the dipole moment of thesatellite, and the de si red change of position, th e computer will deter mine at what point in the orb itthe torquing coil is to be turned on and for how long. Time delay s of up to an hour are availablein the satellite, so the maneuver can begin at the optimum time in the orbit, not jus t over the command station. Time increme nts for this delayed maneuver are 2.5 minutes, both for the delay andfo r the maneuvering period. Fine r con trol can be obtained by turning the torquing command onand off while the satellite is within sight of the station. Once the position is achieved, it is heldpr im ar ily by the momentum of the rot or (Figure 14), which is 2.12 X l o 7 gm-cm2/sec. Driftingfrom th is position is caused by gravity gradient, aerodynamic, and magnetic torques, the latterresu lting fro m uncompensated dipoles. Thetota l of these d rif t ra te s should not exceed 5degr ees pe r day on the average. To minimizeits residual magnetism, the spacecraft is provided with a degaussing coil and a chargeabletr im magnet system to compensate in orbit forresidual magnetism along each axis. It is possible to use this system to produce a specificdipole moment for controlled spin-axis drift.

The nutation damper is used to dissipatethe lateral components of satellite angular velocity c rea ted by the magne tic torquing. It consists (Figure 15) of a torsion-wire suspended

SlMPLlFlEO DRAWING OF ANGULAR MOMENTUM FLYWHEEL

LOWTR-BEARINGFigure 14-Simplified drawing of angular-momentum

fl y wheel.

11


16/25

"f" NU TAT ION DAMPER PERMANENTANGLE DETECTOR MAGNETSPRING RATEADJUST MAGNET, LIGHT S O U R C E ~ L L

COPPER VANE

FIELD COIL

A O R S I O NC A G I N GMECHANISM

Figure 15-Simplified drawing of SAS-A nuta tion damper.

X MAGNETOMETERSENSOR X MAGNETOM X MAGNETICELECTRONICS FIELDAMPLIFIER

- - Y AXIS COILY MAGNETOMETERSENSOR A

X-AXIS COIL

Figure 16-Block diagram of spin rate control system.

a rm with an end ma ss and a copper damping vane.If nutations occur, the a r m o scillates, causing thecopper vane t o swing back and for th through thegap in a permanent magnet, thus inducing eddycu rre nt s in the copper vane. With the rot or on,the r es id ua l nutation should be less than 0.2 degre e; with the r oto r off, it may be a few degrees.The motion of the nutation damper, detected optically by a coded digital mask mounted on themovable a r m is telemetered every 8 secondsvia a 7 bit binary word.

The spin rate control system is shown inFigure 16. Signals sensed by the X- and Y-axismagnetometers are amplified and applied inquadrature t o the Y - and X-axis coils. T his system, in conjunction with the ea rt h' s magneticfield, operates e ssentially as a motor to increa seo r decrease the ra te of rotation about the Z - a x i s .The maximum rate of change of sp in is 10 rpmper day.

Many astronomy experiments can be pe rformed well with t hi s basic SAS control system.

Used in conjunction with a good attitude-determination system, it is ideally suited for sky surveys.Data telemetered from the magnetometers and a digital sun se nsor on board the spa cecraft can beused to determine attitude to &5 " or better.

PLANNED EXPERIMENTS AND FUTURE DEVELOPMENTOn SAS-A, the principal investigato r, Dr. Riccardo Giacconi of American Science and Engi

neering, Inc., is flying star and sun sens or s accurately aligned with the experiment X-ray collimators so that attitude may be determined to a minute of arc . Th is expe rimen t will provide anaccu rate map of X-ray sou rces throughout the cele stial sphere.

After such a survey, which will determine source stren gth, position, and spectra l composition,later X-ray experiments will req uire mor e accura te pointing fo r detailed study of individualsources. The SAS design per mi ts building on this basic control sys tem, by adding a star trackero r star camera system, to provide three-axis stabilization and thus maintain the satellite attitudewithin one minute of a r c or better. This would permit useful experiments in the ultraviolet andinfrar ed regions as well.

Rough positioning of the space craft would be accomplished by the basi c control syst em, andthe exac t attitude would be determ ined by ground computer. Then, fo r example, a program could12


17/25

be transmitt ed for s torage i n the spacecra ft defining a pattern that the star ca mer a should see,and fine torquing would be used t o find and maintain the pr op er position. Two sets of optics withabout 10" fields of view perpendicular to each other could pe rmi t three-axis attitude detection toa resolution of 10 seconds of ar c, and control to 1 minute of arc. E r r o r sig nals would control thespeed of the basic roto r, whose exc es s momentum would be dumped into the earth 's magnetic field(again avoiding expendables). Small gy ro s o r a worm-gear tilting mechanism would be suitablef o r fine control of the X- and Y-axes. A sys tem of this general type is planned for SAS-C andlater flights.

SAS-B will perform a sky surv ey of gamma-ray sources, with Dr. Ca rl Fichte l of GoddardSpace Flight Center as principal investigator. The basic control system is adequate for thispurpose, as the Gamma Ray Spark Chamber ha s a 45" field of view and does not requi re very finepointing.

Goddard Space Fligh t CenterNational Aeronautics and Space AdministrationGreenbelt, Maryland, October 4 , 1968878-11-75-01-51

13


18/25


19/25

Appendix ASummary o f SAS Characteristics

launch VehicleScout with FW-4 fourth stage

OrbitAltitude: 300 n. mi. (555km) *40 Inclination: 2.9" -+0.15"(lo) Period: 95.8 min. Eccentricity: 0.008

SizeSpacecraft: Shape - cylinderDiameter - 22"Height - approx. 20"Volume - approx. 4.4 cu. ft .

n. mi. (74km), or lo

Experiment: Shape - cylinder topped by truncated cone (Figure Al)Base diameter - 30"Height of cylindrical portion - approx. 14" o r 29"Height of t runcated cone - 15.2"Top diameter - 17.7"Volume - approx. 6.4 cu. f t . or 12.5 cu. f t . (for the "pointed SAS" about 1 cu. ft.must be made available from the experiment volume)

WeightSpacecraft: 180 poundsExperiment: 150 pounds - must have low c.g. (see Figure A 2)(130 pounds on 'poin ted SAS")

Continuous Average PowerSpacecraft: 1 7 watts Experiment: 10 watts (9 watts on "pointed SAS") Voltage: +10.7 vdc +lo% -15% Ground: Three -wire system-separate chassis, signal, and power grounds

15


20/25


21/25

30 .OO" D IA . __ !O"- -T'D7i2"15.20" 47.39"32.39" 19 "

14.19"-STA. 18.51 - I- 1STA. 18.51 -

I

STANDARD SCOUT SHROUD

Figure A1 -Outline

.I Y18

v,nz 1730a f 1 E t-ZW

1 5 u Wa X2 140t-I g 1: YI

I W %

1 2

28 2 6 24 20 18 16

ELONGATED SCOUT SHROUD

drawing of SAS experiment package.

-

14 12 10 8 6 4 2 0C.G. O F EXPERIMENT(MAX1MUM POSITION FW D . OF

S P AC E CR A FT /E X PE R IM E NT I NT E RF A CE '- I N N C HE S )*STA 18.39

Figure A2-SAS experiment weight vs c.g.

17


22/25

--

Table A 1Qualification-Level Vibration Specifications fo r SAS

Experiments (based on Scout launch vehicle data).~~Random - 4 Minutes Each Axis

Direction Frequency (Hz)

20-5050-100

Thrust Axis 100-150150-370370-2000

I

Power SpectralDensity (g*/Hz)

0.1Increases at 10 db/octave

1.oDecreases a t 1 0 db/octave

0.07Increases at 10 db/octave

1.oDecreases a t 10 db/octave

0.07

20-4040-150

150-370370-2000

Later al Axes

DirectionI--hrust AxisLatera l Axes

Sinusoidal - 2 Octaves per Minute-

Acceleration -0.187 inchesdouble amplitude

403015

7.50.13 inchdouble amplitude

857.5

.-~

Frequency ( H z )20-9595-175

175-250250-400400-2000

15-3636-200

200-400400 -2000

18


23/25

Appendix BSpecial SAS Requirements

In designing experiments to be flown on the Small Astronomy Sate llite (SAS),in general therequ irem ents of NPC 200-2, "Quality Prog ram Provi sion s fo r Space System Contractors," datedApril 1962, must be met. In addition, pa rt s and mate ria ls must conform to the cur ren t GSFCPreferred Parts List. While specific details must be worked out for each experiment, this list isprovided for guidance.

Parts ScreeningIn general, parts screening will be as described in the curre nt GSFC Pre fer red Pa rt s List.

Only hi-re1 parts will be used. Specifically, fo r all semi-conductors, Le., diodes, tra nsi sto rs,and integrated circuits, the following screening sequence is required as a minimum:

a. Visual inspection before sealingb. Tempera ture cycling -65" C to maximum rated sto rage temperaturec. Centrifuge te std. Electric al te st with variables data recordede. 3 3 6 (+3 6 ) hours burn-in at 100C and 80% of pa rt ra ted powerf. Electric al t es t with variables data recorded ( par ts will be rejected i f outside acceptable

variables limits)g. Fine and gr os s leak tes tsh. Fina l inspectionAll pa rt s must be approved by GSFC before they can be used. Li st s must contain the

following information as a minimum:a. Type of componentb. Value and rat ingc. Manufacturerd. Manufacturer' s type and modele. Manufacturer's screening proc es s specifications to which pa rt is being boughtf. Maximum anticipated electrical stress level

19


24/25

--

ReviewsA design review, a prototype readiness review, and a flight unit re adine ss review will be held

on each exper iment and the spacecraft. Malfunction repor ting and configuration control will beimplemented after the design review and pri or to the beginning of prototype fabrication.

Reliability AssessmentA reliability assessment of the experiment will be performed by th e experiment contractor.

Noise and Transient ProtectionIf the experiment req uire s the u s e of any high voltages, the experimenter mus t provide pro

tection fo r the experiment and the spacecraft sys tems against high voltage sur ges o r radiatedpulses. The experiment must be designed to be insensitive to noise, e.g., that generated by clockpulses o r dc/dc conve rter transients, and the experimenter is responsible f or eliminating unneces sary tran sients and noise from his lin es by means of suitable filt ers , shielding, etc.

If opening the telemetry output would develop a la rge voltage surge to the multiplexer input(more than a factor of 2 over nominal), the output must be paralleled with protective circuits(zener diode). Isolation shall be provided from the subsystem monitoring c irc ui ts in cas e oftelemetry shorting.

Current tran sients in the main power line to the experiment must not exceed the followingvalues:

Current Time4 a 0-5psec1.75 a 5psec - 0.1 sec

(Normal) 1.0 a After 0.1 sec

20 NASA-LangIey, 1969- 31


25/25

AND SPACE ADMINISTRATION POSTAGE AND FEES PAERONAUTICSD. C. 20546 NATIONAL AERONAUTICSWASHINGTON, SPACE ADMINISTRATIOFIRST CLASS MAILFFICIAL BUSINESS

POSTMASTER:

T h e aeronaiiiical and space activities of th e U nite d States shall beco n d w ted so as t o contribute . . . t o th e expansion of hutnan knowl-edge of phenomena in the a tmosphere and space . T h e Adminis tra t ionshall provide fo r the widest practicable and appropriate disseminationof inforiiiation concerning its activities and th e re sd ts thereof.-NATIONALAERONAUTICSN D SPACE ACT OF 1958

If Undeliverable (SectioPostal Manual ) Do Not

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONS TECH NICAL REPOR TS: Scientific andtechnical information considered important,complete, and a lasting contribution to existingknowledge.TECH NICAL NOTE S: Information less broadin scope but nevertheless of importance as acontribution to existing knowledge.TECHNICAL MEMORANDUMS :Information receiving limited distributionbecause of preliminary data, security classification, or other reasons.CON TRA CTO R REPOR TS: Scientific andtechnical information generated under a NASAcontract or grant and considered an importantcontribution to existing knowledge.

TECHNICAL TRANSLATIONS: Informationpublished in a foreign language consideredto merit NASA distribution in English.SPECIAL PUBLICATIONS: Informationderived from or of value to NASA activities.Publications include conference proceedings,monographs , data com pilations, handbooks,sourcebooks, and special bibliographies.TECHNOLOGY UTILIZATIONPUBLICATIONS: Inform ation on technologyused by NA SA that may be of particularinterest in comm ercial and other non-aerospaceapplications. Publications include Tech Briefs,Technology Utilization Reports and Notes,and Technology Surveys.

Details on the availabil i ty of these publ icat ions may be obta ined f rom:SCIENTIFIC AND TECHNICAL INFORMATION DIVISION

NATIONAL AERONAUTICS AND SPACE ADM INISTRATION

the small astronomy satellite (sas) program

Documents