FINAL REPORT October 1968
Prepared for National Aeronautics .and Space Administration Manned Spacecraft Center Bow ton, Texas
Prepared by:
Martin Marietta Corporation P.0. Box 179 Denver, Colorado 80281
MABTfIII MARlmTTA eORPORArBQN DENVER DIVISION FOREWORD
This report is submitted in ac:ordance with the requirements of Contract NAS~-8343, Article X, Reprts of Work. The report is prepared in two voluacs. Volume I contains the basic report. Volume I1 contains Appendices A and B.
OWAmr##U MAR#- ~'QRPoIBA~#~# DENVER DIVISION PR34-113 iii
CONTENTS
Page
Foreword ...... ii Contents ...... iii I . ~ntroduction ...... I-1
11. Summary & Recommendations ...... 11-1 111. Scientific Objectives. Experiment Groupings and Requirements ...... 111-1
111.1 Scientific Objectjves and Experiment Groupings ...... 111-1 111.2 Experiment Requirements ...... 111-8 IV . Experiment Cost and Availability ...... IV-1 V . Mission Constraints ...... V-1 V.l Lunar.Coverage ...... g-1 V.2 Payload/Delta Vt?locity Cansiderations. 8-15 V. 3 Crew Systems ...... V-26 VI . Subsystem Design Requirements ...... VI-1 VI..1 Data and Commurtication System .... VI-1 VI .2 Power System ...... VI-12 VI .3 Guidance and Con.trol ...... VI-25 VP.4 Thermal Control ~istem ...... m-40
MARIIAl ~LIRP.P~~ABORPQAP4r8SCV DENVER DIVISION F i ~urc Page
IIT.. 1-1 Basic, Block Diagram for Lunar Orbital Science Experiment "KoaJn,ap" ...... ,. Lunar Orbital Science Experiment Groupings for Subsurface Regime .....
Lunar Orbital Science Experiment Groupings for Surface Regime ......
Lunar Orbital Science Experiment Groupings for Atmospheric and Iono- spheric Regime ......
Lunar Orbital Science Experiment Groupings for Cosmic Regime ...... Typical Experiment Operating Timelines. .
Experiment Cost and Schedule Estimate Summary ...... Lunar Surf ace Coverage ......
Areas Accessible for Orbital Experiments ...... Experiment Utilization ...... Subsatellite Turn Analysis ...... Subsatellite Analysis - AV .....'.. Subsatellite Analysis - Propellant Required ...... Sensor coverage of Lunar Site ...... C&SM Camera Coverage of Landing Site . . Extended Sensor Coverage ......
Extended Sensor Viewing Time ..., ...
MAA~INnwAmeI"rra UQ~P~~PA~~OAP dENVER DIVISION Sitc Analysis ...... V-16 Site Analysis -Av ...... V-18 Plane Change Requirements ...... V-19 Site Analysis - Payload ...... V-22 Inclination E£Eects ...... V-24 Crew/Equipment Timeline for Lunar Orbital Period ...... '. . V-31 Crew Visibility ...... V-33. C&SM - Sector I Configuration During EVA...... e. .'J-36 Typical Timeline EVA ...... V-37 Stable Satellite - Lunar Orbiter - Data System ...... VI-4 VI. 1-2 Pioneer Spinning Satellite Data System ...... VI-7 VI. 1-3 C&SM - Group I Experiments - Data System ...... TI-3 VI. 1-4 C&SM - Group I1 Experiments - Data System ...... VI-9 VI. 3-la Orbital Stabilizat3.cn Coordinate Systems ...... VT-27 VI. 3-lb Orbital Stabilization System and RCS Mechanization ...... VI-28 VZ. 3-2 Propellant Usage as a Function of 04-G . VT-30 €3 VI. 3-3 Auxiliary Command and Service Module System for Lunar Crbital Experimeat Orientation ...... VI-32
AIART'8N MAR#-A CPBbPPOR&76OM DENVER DIVISIQ,N Figure Page
VI. 3-4 Lunar Orbiter Block Diagram ...... vi-36 VI.3-5 Lunar Orbital Relationships ...... VI-38 VT.4-1 Heating Loads - C&SM Group .I Experiments ...... VI-45 Heating Loads - C&SM Group I1 Experiments ...... VI-46 Heating Loads - C&SM Group I11 Experiments ...... '...... 1,71..47 Maximum Heating Loads Condition - C&SM Group I1 Experiment Coolant System . . . n-48 VI. 4-5 Heating Loads - Stable Satellite Group I11 Experiments ...... VI-51 VI. 4-6 Maximum Heating Load Condition - Stable Satellite Group 111 Experiment Coolant System ...... , . . VI-52
IWAATffW tW~~8-a OlbAPOWANBICI DENVER QIVISION Table
111.1-1 Satellit2 Requirements tor Lunar Orbital Science Experiments ...... Potential Experiment Groupings ..... Experiment Development Status, Other Programs ...... III. 2-2 Lunar Orbital Science - Experiment Requirements Summary ...... ' .... Experiment Contamination Summary .... Experiment Orbit Requirements Summary . .
Crew Reauirements Summary ...... Experiment Deployment Requirements Summary ...... Peculiar Ground Checkout and Servicing Requirements ...... Peculiar Subsystem and Support Requirements ...... v. 2-1 Payload Comparison Abulfeda ...... V. 3-1 Crew/CM/SM Experiments Tasks ...... V. 3-2 Crew/Ex~erimentsTasks .. :...... v. 3-3 CM Stowage Requirements ......
VI. 1-1 Summary of Data System Modifications Required to Imp\ement Experiments .... Groupings Presented ...... Sample Experiment Grouping - Input to North American Rockwell - ~u~ust1968 . .
MART#N MAR#-& eO4WP.ORAr#@N DENVER D'VISION PR34-113 viii
-Table Page VI. 2-1 Stable Satellite Subsystem Power Requirements ...... VI. 2-2 Stable Satellite Power.Summary . . . . . VI. 2-3 Spinning Satellite Subsystem Power Requirements ...... v1-19 V1.2-4 Spinning Satellfte Power Summary . . . . VI- 20 VI. 2-5 C&SM Subsystem Power Requirements . . . . VI- 22
VI. 2-6 C&SM Power Si~~nrnary- Group A Experimen tS , . , ...... VI- 23 VI. 2-7 C&SM Power Scwnary - Group B Experimgnts ...... VI- 24 VI. 3-i. Comnand and Service Module Attitude Control Add-on Alter~lates......
VI. 3-2 Subsatellite CharacteristYcs Attitude Control ...... VI. 4-3. Thermal Control Requirements for Sstellite Missions ...... VI.4-2. C&SM Experiment Grouping ...... VI .4-3 Stable Satellite Experiment Groupings . .
VI. 4-4 spinning Satellite Experiment Groupings ......
IWAAIYIIV bW4BOm-A O&~BORAT~ON DENVER DIVISION * Appendix A ExpGriment Description Documents
Metric and Stellar Camera
Panoramic Camera
Laser Altimeter
R.ldar Al t-ipetez
Ccrchemistry Experiments
Mass Sprctrometer
LK P!apping Radiometer
IR Spectrometer
Microwave Radiometer
RF Transponder
Bistatic Radar
Photoelectric Photometer
RF Reflectance
Electromagnetic Measurement
15. Radio Noise Survey
16. Radar Imager
17. Meteoroids
18, . Solar Wind Foil
19. DCMagnetometer
20, AC, Magnetometer
21. Plasma Probe
MA#PIN JW&RBS1;1a C6RPOIPAf #QIAY DENVER 31VlS10N . 22. Elect.ric Field
23. Low Energy Particles
24. Solar Energetic Particles
25. UV Spectrometer
26.. Laser Retroflector'
27. Gravity Gradiometer
28. Swept Frequency UHF Radiometer
Appendix B ~x~erirnentCost and Schedule Estimates
AIdRBIN MAAelSrrA 06RP8RLOr#8U DENVER DIV!SION PR34-113 Page 1-1
I INTRODUCTIOM
In June 1968, Martin Marietta ~or~orakionwas awarded a three month study as part of a NASA-MSC Phase I3 Lunar Orbital Science Study. The study resulted from fhe 1967 Summer Study of Lunar Science and Exploration at Santa Crux. The total Phase B study is to further define the scientific use of lunar flights. The Martin Marietta study was directed toward the further definition of the scientific use of the orbital portion of the lunar flights.
A majority of the effort, in accordance with NASA-MSC direction, was applied to baselining experiments coopera- tively with experiment scientists and accomplishing a Eunet imal flow grouping zn~~ysis of large numbers of experiment -combinat ions identified as principal goals at , the 1967 Santa Cruz Conference. The experiments were baselined and .a new analysis technique devised based on pa.yload, orbit, mission and data management constraints . The study confirm that a lunar orbital science program can be developed for achievement during the 1972-74 period.
MAWrIN MaRlS-4 00RPORArJON DENVER DIVISION PR34-113 Page 11-1
I I SUMMARY Am RECOWNDATIQNS
This contract was accomplished as a part of. the NASA-MSC Phase B Lunar Orbital Science study resulting from the 1967 Summer Study of Lunar Science and Exploration. The contract has as its objec- tive further definition of the scientific use of lunar orbital flights through analysis, review, updating, and consolidation of experbent requirements, development status and feasibility , grouping, and miss ion,requirements . The Phase B study represents a furthe,r step toward the preparation of detailed science mission plans, a process which is necessarily iterative in nature.
The Martin Marietta study placed primary emphasis on achievement of the Santa Cruz Conference objectiv,es through a fundamental scientific analysis of the experiment objectives and their data requirements, and limited iteration of these with proposed exper- iment sensor capabilities and groupings derived during the study. Experiment mission and subsystem requirements were analyzed, based upon sample experim~ntgroupings. 8unrmarized results of the stu'dy are:
1. With the Lunar Science and Exploration Conference as a baseline, scientific objectives were grouped into four regimes: a) Subsurface, b) Surface, c) ~tmos~hericand d) Cosmic. The sub- surface regime ;;as the least defined and required particular atten- tion during the'study.
2. Ar, analytical grouping ratisnale (roadmap j was developed which introduced a more.fundamenta1 technical experiment grouptng approach than has been used during the early phases of previous experiment programs. The primary result of the roadmap analysis is the !.rfentification of correlative experiment groupings required to produce unambiguous data toward the scientific goals. WFth this roadmap, each experiment was analyzed against a series of seven engineerfng and scientific constraints. The Phase B analysis shows that all but four of the originally proposed experiments are generally compatible with both scientific objectives and mission constraints. Significant clarification and combination of sensor requirements were achieved, and one new experiment was identified for consideration. It is a swept freque~cyUHF radiometer, for study of the lunar subourface regime.
3. Of the 30 experiments analyzed, 6 were in flight hardware configuration, 17 exts t in nsn-flight hardware configuration with a high confidence level that lunar flight can be achieved; 7 require new development. Contact with responsible scientists and industrial sources in each experiment area provides assurance that instruments P1234-113 Page IT-2
can be avail.able for a lunar program in the 1972-74 time period, A preliminary cost estimate for the 30 experiments indicates a toea1 cost span of 60 to 85 million dollars for flight hardware. . Further def inition work, e. g., specifications and a detailed pso- curement cycle is recommended to arrive at a closer price defini- t ion.
4. Fif ty-seven experiment groupings were gener-ated us ing . the roadmap technique. These groupings are categorized into minimum, medium and maximum scientific data accuracy groupings for achieve- ment of the object Fves sf each regime (Subsurface, Surface, Atmos- pheric or Ionospheric and Cosmic). Neither selection .of specific groupings for particular missions nor programmatic optimization of the number 0.f missions was worked. However, early flights would be expected to use groupinds of minimum accuracy since their ins truments are generally less amplex and more readily avdilable. Additional work in experiment grouping and detailed instrbment definition remains for accomplishment after the program constraints are evolved by Phase B.
. The large number of groupings (57) indicates the advisability of computerizing the sya tem interface requirements (weight, vol- ume, power, data) for flight grouping since more than one group can be fXowm by each flight, once the payload allowables are better defined. Computerization allows optimization toward flight schedules and experiment costs. Furtheanore, the split of CGcSM- borne versus subsatellite-borne experiments (some of the experi- ment groups can go either way) can be optimized.
5. An opt hum mission for most experiments would be polar and of at least 30 days duration to provlde adequate coverage of the lunar env%ronment. This points clearly to a requirement for subsatellite usage in the program. The near. equatorial orbit con- s traints of the Apollo Spacecraft permit sens ing of only about four percent of the lunar surface during the experiment operating periods; however, substantial achievement of certain experiment objectives can be accomplished. A large amaunt of data can be collected by use of the Apollo Spacecraft only which for early flights will represent an acceptable data return and will provide a significant contribution to the more cornpiehens ive exploration planned for later flights. .
6. There is a. significant problem area in the requirement for data reduction and analysis following.each flight. Further attention to the entire experfment data management area is advis- able. Application of advanced data handling technZques has favorable program cos t implications which should be inves tigated
MAmr8W MAffIPW.4 8QRPQmAr88AI DENVER DIVISION PR34-113 Page 11-3 prior to a ?hase C start. An analysis of allowable data com- pression within the MSFN on the basis of intrinsic experiment data accuracy requirements is also recommended. It will result in an almost real time experiment data availability with a tremendous impact on the visibiltty of mtssion and experhent success. If there are new discoveries in lunar science they could be as sertained for announcement immediately after the flight - the present time delay is estimated to be 1;; years.
7. Preliminary timeline analysis shows that most of the crew time is required for basic mission performance. However, the crew requirements for experiment operation are within the design control cognizance of the experiment integrator. EVA to retrieve camera film was found to be the only mandatory require- ment for crew participation. Deployment and actuation prccesses for either expertments and/or subsatellites can benefit from crew control. EVA for film retrieval is probably best performed at a miss ion time subsequent to the return of the LM when all three astronauts are avaf lable.
8. ~a~loadcapability varies depending primarily upon the landing site selected for the basic mission. Within the mission constraints of this study, payload available for scientific experiments varies from zero to approximately 1900 pounds.
9. Subsystem studies were made on the basis of selected groupings and show a requirement for various subsystem changes in botn the CSM aqd the subsatellite eandidetes. Pinal defi- nition of these changes is dependent upon experiment gr.oup selection and will require additional investigation and iteration. General conclusions can be summarized as follows: . Data System requirements on the C6rSM inczease moder- ately, while on the subsatellites , they range from about a 40% change (Lunar Orbiter) to complete re- design (Am) . . Power S,ystern requirements on the G&SM also increase moderately, while the subsatellites require almost double their present capacity and significant sys- tem design change. Most components are usable, however.
. Guidance and controi appear adequate for most exper- . iments,.with the most stringent requirements being -+lo on attitude control and .-$. 05O fsec. on angular rates. Some experiments require special pointing, , PR34-113 Page 11-4
such as for a stellar reference or earth pointing antenna. These experiments may impose some modi- fication requirements on the guidance and control system, although these appear to be well within state of the art capabilitiee.
. Thermal control will be required by some experiment .' groupings ; for example, one stable satellite group- ing studied requires an active thermal control sys- tem weighing approximately 190 pounds. In this particular case the subsystem would represent a significant portion of the total subsatellite weight, indicating a requirement for further study in the experiment grouping and in the thermal control sys- tem itself.
MLIR7tN CWAIPITWA UORPORAFION DENVER OlVlSlON PR34-113 Page 111-1
111 SCIENTIFIC OBJECTIVES , EXP1:RIMENT GROUI'INGS AND K~.:QUIRFMENTS IlI.1 Scientiiic Objectives and Experiment Groupings - Using the Sonto Cruz Conference as background, four broad regimes for lunar science exploration were identified, each regime dealing with a different part of the moon and the lunar environment. The four regimes are:
1. Subsurface 2. Surface 3. Atmosphere and ionosphere 4. Cosmic
Although these subdivisions have some overlap, they represent a logical sys tem when considering the grouping of scfentif ic ob jec- tives and experiments for lunar orbiting sensors.
Responsible scientists and industry sources were interviewed to ee tablish the objectives and instruments which could meet those objectives. The data oti objectives and instruments provided a ' functional flow or roadmsp approach (Figure 111.1-1) developd to establish experiment groupings. For each of the four regimes, basic scientific objectives were listed. These were in turn sub- dividad into specific sub-objectives. Candidate experiments which could meet the objectives were then listed, and fed through a series of seven constraints or roadblocks which defined estimates of 1) feasibility', 2) availability, 3) measurement constraints, 4) instrument constraints, 5) orbital requirements, 6) data re- . quirements and 7) data interpretation. If an experiment failed to satisfy the requirements of a roadblock, it was eliminated from further considerations. The roadblock which was responsible for elimination of an experiment is indicated by a blacked lower right corner. The requirements of surface structure, atmosphere and ianospheric, and of the particles and fields (COSMIC) of the moon are well understood. The Santa Cruz Conference report gives coherent rationale for those regimes. Therefore, as expected, all experimdnts for these groups pass the ro.adblocks as ahown in Figures 111.1-3, 111.1-4 and 111.1-5 respectively.
The subsurface regime of the moon is the leaet understood of the baseline exploration considered by the Santa Cruz Conference. Only recently has the possibility of exploring this regime shown 'up (e.g,, the "Raisin Bread" phenomenon has dram the attention of the scientific community), Therefore, the subsuzface regime "as emphasized by MMC and the ~antaCmz baselice has been updated in cooperation with the experimenters concerned. Four experi- ments in the subsurface regime were eliminated by the roadblocks, as shown by Figure IZPsl-2. The elimination of four experiments . led to identification of the swept Erequelicy UHF radiometer.
MAIII8m AOARIIIC'WA Od9llPOmAlrlON DENVER OIVISIOF' 1. Subsurface 2. Surface & Nenr Surface REGIMES: 3. Atmosphere & tonospirere. 4. Cosmic
MAJOR SCIENTIFIC MAJOR SCIENTIFlC OBJECTIVE
SPECIFIC SCIENTTFIC SCIENTIFIC SCIENTIFIC SCIEhTI FIC SCIENTIFIC SUBOBJECTIVE
R.B. $1 - Can experiment be performed? I I R.B. /k2 - Will flight qualified hardware I be availahle by 1972? . MEASUREMENT - . R.a. +'0 Resolution and sensitivity CONSTRAINTS - ,,-,&--, LLmLLac L ons, etc.?
K.B. *4 - Antenna size,power requirements. etc. ?
R.B. 115 \:'hat orbits should I>@ flowr;? REQllIREMENTS -
R.R. *!6 - Vhat data rate is required? REQUIREMENTS
ACCURACY ACCURACY ACCtitiACY GROUPING GROUPING GROUPING
I / / TOTAL DATA I / TRANSMISSION -1 CONSTRAINTS
BASIC BLOCK DIAGRAM Fnfi LW4R ORBITAL SCICNCE . EXPERIMENT GROlIPING "ROADMAI'" ------*- ---
KtClnk
TO ' temlnm varlatlm 1n umm and #hare of .ooa h 1oc.I don,- Lty vsrlqlono 8
ma h.V" used in ponders on lunar sur- satrlllta tr.cklng facr or drop and land Iran .arch unlts from LJ,
ntnf work for ..,.s. Ll&htu L. . -. l~.l.lil~~ r. I'ravldcr data an 112 Rovldem, data on 112 ,,. of lunar murhcr only8 trscker ptm:, . complete orblt ultP 0f 1un.r nurhct only t 25 It. nl,t',4'. rufflcltnt transyondrr .ever0 earth etro~pher- accureey riIc camtralnts No ritema1 accrl.ra- tlon. p-RLLted; Ion8 tnte(lr.tlon ?Lac. rn- Any orblt deru r.tln~ on r-nl~slonh < lla bm -I NuIm -0Up Track tram *arch. IImhLrlmht mer a8 Imrga port of the shed4 allow accur.ta uFQlnL r.6- cdriqup to 111 th8 bri&tn*s. t entire 1-r aurfac. a0 poaoibla aor Itnitel lumar ouutara ace. lmr #urf*ae (far .Id. mlrtd) .o Iarp a p ourface .a p ------w------.- > I"'-* on s,rh1tt r F7-lPO' den.. f 15. acr. rep* or color Ill.: 10 Llloblt~l~ec. J-&&dlo~t.r 6 K-band RI- my reglre ulcl-u4velanph IR rad1aat.r to pln dmkin-tic murlaer I. q%.racure. tcale channel U.H.F. unlc at eurhca aad tw dapth. with Match rmmolutton (around mratb) to fly-6 n~a~anatnarad101tona. orblt hal@t to provld* conclwoua mm sbmal par rallmtar Sma. dnlm accur~ynoup. brl#hcmaaa tq.raaura b &a- ucept Vlctmr brl&cnma~ depth prolllaa 6 muparlor denth Inparacur* profLLas roaolucla mLUty war u larp a part ef aurlarm rtlra lrur mrlna ma watblo Swept heaumacr URf Radl- ~1mr.r' .ad CO.POO~CLOR of 4*pooLts Csll r Sub. mlltta VahlcIm. 1b- ~n:'t ch.Al.-61.3 ~b1~;7;.s Mu: 4 ckennalr-122.5 Iblt.l..c tr*- truedttd krk La "real rlr* ruordrd 6 r1tres.dct.4 ep fa: ;,r -- C - r-.... wd..at,nd .tru!trrc rf lunar interior I I-I I tla. of lunar Interlor 1 of lunar Inrt*rlur Orvclopmet't i.ce.'cam Gp~ntgrogr.~ rdqulrcd IH no. - 2 yr r~~ulrod2 ye.rs. 1-1 -4.1 1.7; 1 Altlmvter function h.1 10s - perlodlc Cnolirg nf opt~clby antmn~mrequlrcd for P.,,~YC I.d~.c~VC lr frr-q.. mccnna #LIP 11dt.tIn 10: 1 bandvldth . 2 pol.rI~.tl~nai aide , ~~~'~~~el~r~~~:~ lob* sunt.rrralun 1 i SInp11~c.llhratim qu1r.d for lnterpr*t.- qu1r.d; ard.; slmulntr rlon of data no trurh sItr rrq.; color 111111 with .I. nee1 trmp*ractrre dom to I depth oI LOO It. 5 k~lobitalmmct.Corn a tape a 10 kirebita/mac recot6 en tar far aid. 02 moo and r.tranl.ir on far did. of &n snd rltranrdv 71 IU-20 klLoblt.l~*c -1 FplFP\ klyrq m n.ms rtr 90' dr.. , i 15. sec. h - 50 10 IUO k. 90' dm.. + 15. ace. 7.5 kllohlts/..c~ Car sldc of mun Rwqeirrr trucls hltc Jet., Lor lmh~r Obtsln aynoptlc ups of lu-r Hatch ~porlol reaolutlon of a tarrain ra~lllmadurn" of "lmfocuo~d" Record data on tap* on far a11 radar lmr*~. airtd pphnible radar I.M~IL-datr poaoible of won and r.tronmmlt I 7- - - To datarmlnc lunar surlrce :o daterdnc theraul PLOP. and I Rznta Rznta -and 7 . . -1 I 1 \ /w\ / 1 - fWW\ f U'\ 4 fU (nnn-acannlns) NICKOVAYt RADILW ?I R IXI'tKIHtnT IR 5l'kEl (non-mcannlny) Car f1llr.d (1rap0rtlon.l eauntcr* and electron multl- --- "dual elm me1 d' ~tmpicllorll a'nl.ll.r.; ~Stat~-of-th~.C~t" duvalr,swnt prayram 1.r.d muPftd. detsctorm held to 25.C; 12 to 50 channe1.i IOU hltnhec., record .*cord data on Iar on data far .Id* of .Id. 01 moll on tape mocm on tap" and and rctrananlt Ircl ran-ult HeIsr.nc. 1.h dm:. on eartt,; rcqulrcdi trutb .Ice required m.t 'slulate col0.r r-1shift ulth as* K-lnd Y diortrr (non"1c.n) d da Iu u (untoc~s~d) A-Lnd ;#Cr Imccr (unfecu.ed) ?~~-IIXG%TGO~.scan) %%pp2a:.dl&Fer K-Band rad1aet.r-1"p.r i'~nfoc~~~d"mde imam- 80 1 IR DPLnC H.~I-C.P I I I reiolutlon prov1d.a brat ramolu- 11 Suhaatrlllte Veh:cle carrier and radlnrter n and r.transmlt record data on ra*. at a11 it-. CI't.CTRUI(LT1 R cnuntrrs 6 argsnlc mclmlll.tor. 1 to 110 ....a. 10-13 torr Senior. channel rprrl ronrtor. avallahlr Itv 1112 In.lrummt dsvelnpmml specs. to he drlernlnsd '40. da*. . + 15. mcc. I h - 100 km 25 blt.1~~~. !Ired. controlled rel-ale of tracer 8,s Ira aurface to of flux. mmentua callbrala Ieshnlquc ObeaLm crude up* of ~lectrlcml on nrmr mid* of won only ~urthcmam waalblr. propertlaa amr am laram a part of lunar surface m. pommlbtr pommlble 100on bltalaa=. ,ldm of loon Photatar data on rap. or Ill" 6 rod7,k. data ]rjdlvp on "ntaraide" of wont ermnm- rl~{or mod to b. ratranm.lttod when orbit; r and r*lrmnnlt Ile r-lnlnl 4.1. in "realtlw", v.hlelr r-mcham near mid*. 1 except rhar on far slda af .wn. atrornld flux and VI~u~1and near prclral type 1R Lwslnl I 0.35pIO 1.rp I men1 on CEH; Ll#ljt- vel6hI llnlta to hr AIL cqul~.nvnt cur- nrmnrlv .vaLlahla (~cr(wndlcul.r lo LIZ of lunar I pol, on rccrlvs rurfacr only 4lllneltr funcllon X-band requlrcd lo reduce slzc of Yo data llnk I.? klloblt.l*uc' not opllllc.l,ll ; to'..l I no o.p.ra1. data (2 cbsnnrl* ruqulred) Hnk rcqulrsd I -MXIN'Y ACCIMCI GROllP UINlllU ACOMCK OnWP ACCUMC\ mra6s07rs - . RP 1r.n.wnd.r. &-hnd Udar Iumr(unfocu.ad) 2-2~) Altimrtar Trach f~4lunar orhlrvr Track Iraq earth K-Lnd H.dlacler-Iuscr IR Lpnlnl Radlwlor lzll 1:11lia. a~1ttatIcm1 and qdar~lv. ' IR I(.~nl'yRIdLmrttt mclhodo to obtain up01 trrraln ' shape cat=porlaa 1-n) ll Suboatllllte Vahl le t. rrl nt Submatclllle V.hlcle Experlwqt rsllon covarln# dp to 112 of lunar 4urf.e.e 65 k1lobll.l~;~. Take Jatr on wary mecond or Xecord data on I thlrd orbit; r.sord &ta on tap. a1 a11 tira b hrp &tm on Of won. To d=termlne lunar murfacc dnr. lor Iandln~rlte cer,tlflc.tlon 6f the Lunar murfec*. and future mlaslon planning. I- - ro ?.5 n*trr* lor ".I"" 1.we.r plnot~~raphy .tll:ud. lntornatlon dcvelolrvnt hy modlftcallon to THC. Inc. rxI~Ilngcanyos L~~IIIuLL.:~~Unit. . could hc developed Yllsht quallfll?d unlt 2 par d*vmloplmt ready by I972 .ffort; unlt rrrdv hy I972 t 15' lnc. Lraorponder vlslbl1Ity la IUD km ,"d k. - LOO + Is' IIIC., h - LOO km PI 1. rn hlrn of .~rbitlor rlv:( nnlnn rlon of hi~lwrur.1b.r b.1~upplng nl.slon of lunar surfwe Rscord all 4at? on Ill. bnd t. a tone 1 .-- - I r-- 8: a:-..., r: . . Y.. 'T'\? :T X'. V : - L---- ir rGr*;-il?a sp:ar b%,s' .Ihc;s I( a decror .f roqLrac1m O f ir :ete;-l?* rrl:. )06 I 10 d.tsi-&n. iv and :rol:o, I' +unar lm@~--.r. . : :. ;,-; er* LS**L.L~~Ln :. ar ialos--.a I ,.:-. ., I I I I I. . . , . . . ...... 0 nrklr dor . :- ra hL. .. s . . . .. I5 b5ts r=c i-1J : PR34-113 Page 111-7 shown in FICURF: 111.1-2. Now this regime emerges with n I'ull con~plement01' groups necessary to solve the subsurface ex plor- ntion problem. The result of the roadmap development was the identification of a large number (57) of experislent groupings. These groupings reduced the extremely large number of possible combinations of 30 experiments to more manageable proportions. A further reduc- tion of combinations was achieved by defining ninimum, medium, and maximum accuracy groups. These groups reflect estimates of experiment sevitivlt:/, and additionally provided a measure of system complexity. Minlmum. accuracy groupings allow early flights to obtain some of the data useful in meeting the lunar science objectives. Maximum accuracy groupings provide more and better data, but generally at somewhat higher cost, weight, and data transmiss ion rates. The number of mFssions required to meet the lunar science objec- tive~was not estimated. However, .the groupings presented as the roadmap output allqw future mission estimates to be made, based on the ground rules and vehicles available at that time. The type of satellite on which each experiment candidate should be flown in shown on Table III.1-1. - Table 111.1-1 sateilite Requirements for Lunar Orbital Experiments - I Connnand and Semice Module CSM or Stable Satellite Metric and Stellar Camera XR Mapping Radiometer Panoramic Camera IR Spectrfmteters Laser Altimeter Microwave Radiometers Radar hager Radar Xmager (if image on film) Radar A1t ima ter RF Reflectance Swept Frequency UHF Radiqmeter Bistatic Radar Gaoch~istryExperiments* Electromagnetic lfeasurements Mass Spectrometertr Photoelectric P.hotometer -- RF Transpondem Spinning Satellite W Spect~mttW Meter- . qs AC Magnetometer Solras d Foil DC Magnetometer Laser ,etroref lactor Plasma Probe Gravity Gradiometer Low Energy Particles Electric Field* Solar Energat ic Partdcles hdto Noise Survey 7- -- *Can use on W3M; however, possible background noise, contamina- tion problem. Prefer more extensive coverage provided by higher inclination, longer mission. *Can also use on spinning satellite. Four missions were identtfied by NASA as csndidates for the lunar science program. Four representatbve group%nis, shown on Table SfT.1-2, were developd. Subsequent sections of this report dealing with subsystem supwrt requirements are based on these groupgnga. It should be noted that the groupings, derived jointly by NASA and MMC early in the study, hold up quite we11 when compared to the roadmap. I 'I Table 111.1-2 Potential Experiment G~oupfngs [ 1 )-nd and Service Module [ Stable Satellltc .-----.--- .. - . -.---- -. ------. ---..---- .------I Metric and Stellar Camera Geochembtry Experiments Panoramic Camera Mass Spectrometer Photoelectrtc Photometer RB Transponder Laser Altimeter ' Meteoroids Solar Wind Foil Experiment Weight 712 Ibs Experiment Weight 526 1bs Inclination 0-15O Inclination 90° Covgrage 1- 4% Coverage 100% ------.------. Stable Satellite Spinning Satellite IR Mapping Radiometer AC Magnetmeter = Spectrometer DC Magnetometer Microwave Radiometer. Plasma Probe RF Reflectance Electric FLeld Bistatic Radar Low Energy Particles Radar Altimeter Solar Energetic Particles RF Transponder RF Trans ponder Elec tromagnetf c Measurement8 Radio Noise Survey Experiment Wetght 644.5 Iba Experbnt Weight 73.5 lba Xnclination 90° I Incl fnatfen 0-90° Coverage 1007, t Coverage 1002 ?!here will be further groupfngs in terma of payload capability which can be achieved by combining the scjlentific discipline groups identi'fied on the roadmap. The final selection of the groupings will be based on constraints such as weight, lunar orbit inclination, nodal points and duration desired. - -- 111.2 Experiment Requirements - Experiment requirements were obtained by a survey of responsible sciantisto and potential induetry vsnr dors . The requirements obtained, -re titen compiled an? documented MAlYN#V IUARJIWA OOffCBRA~#OM DENVER OIV1510N PR34-113 Page 111-9 in Experiment Description Documents (EDD) enclosed ae Appendix A of this report. Tho ED& were prepared to the f12rmat of .the NASA/ MSC Phase A report, Appendix A, and irkcludo experiment objectives and measurement concepts, with the requirements of bpecific hard- ware which most nearly satisfies the objectives. Many states of development are represented by the experiments. Of the thirty experiments surveyed, six are of flight hardware configuration, requiring only adaptation to the potential space- craft. .Another seventeen experiments are state-of-the-art. These have not necessarily been previourvly flight qualif Fed, bui: suitable instruments are available which can be modified for lunar flight. The remaining seven experiments will require sensor technology ' improvement and development. The state sf sensor development can be deduced Prom Table 111.2-1 * Experiment requirements are sumnarized in Table 111 -2-2. These requirement6 are also presented in somewhat greater detail in the EDDs of Appendix A. The majority of the experiments have mqny common requirements; however, some differences can be seen. Power requirements for.most experimants are compatible for use with 28 volt dc systems. The K-band microwave radiometer is an exception since it is an application of a Nimbus unit which uses -24.5 volt power. Another exception ie the bistatic radar which requires 50 volt, 2400 Hz square wave primary power. The Mariner 1967 Venus probe unit is reflected, again an application of exist- ing hardware. Temperature lhits are reasonably broad, with the exception.of film cassettes, some of the geochemistry units, and the IR spectro- meter. This latter unit presents the tightest temperature control (23'~ -+ 5O) limit. Data storage and transmissi0.n requirements range from zero to a video link requirement of 12 megabits/secsnd, f~rthe photoelectric photometer. All instruments requiring greater than eight $it resolution have individual analog to digital converters, and pxe- sent a digital data stream to the spacecraft data system. Capability to hold the spacecract within 2 lo of local vertical should be adequate for all sensors requiring this pointing mode. Control of attitude rates on the oxder of 0.05 degrees per second appears to be the most stringent requir-nt. Several of the sen- sors require pointing modes other than local vertical. The solar X-ray and cosmLc ray portions of the geochemistry experiment and' MAdWT8AI MAAIRWA UORPUBRAr8OdV DENVER DIVISION' PR34-113 Page 111-10 Table 111.2-1 Experiment Development S*atus, Other Programs 1 ------..-.- --.------I Exper ime n t 1 * Development Progrnm -I - Metric and Stellar Camera Fairchild study for MSC Pa~~oramfcCamera 1; None for space application Laser Altimeter 3 Cohtro? Data Corp. development I contract Radar A1 time ter 2 None orbftal, SIB unLt flown 1 to near orbit Geochemistry Experiments* 1 2 AShB study fcr GSFC Mass Spectrometer /Z UEP (similar) IR Mapptng ~adiometer ' 3 MSC s tudv LR Spectrometer j 2 None Microwave Radiometer 2 Nimbus E (Space General) RF Transponder 1 C & S?, Lunar Orbiter aistatic Radar 2 Mariner 1967 Photoelectric Photometer 2 None RF Reflectance ( 2 MSC aircraft units ; no orbital development Electromagnetic Measurements 3 None Noise Survey 3 RAE Explorer in earth orbit Radar' lmager 2 Venus probe (JPL - Mariner) Meteoroids ' 3 MSC study Solar Wind Foil 2 ALSEP Magnetometer (D.C.) Magnetometer (A.C.) Already developed for Plasma Probe many space miss ions IAW Eraergy Particles '1 (Am, Pioneer) So1,ar Endrget ic Particles 1 } ElectrSc Field 3 None . UV Spectrometer 2 Mars probe - C.U. Mariner Laser Retroref lec tor ' 2 AZSEP Gravity Grdiometer 3 None d *Category 1 - Adaptable Flight Hardware Avai1abl.e 2 - State-of-the-art 3 - New Wveloprment WCombination of four experiments . . . - Cot SS!? LLtiAK 0Kl rTZ SCIY..:t - t:XIF.%l.%lii -S':'':rC5.::TS AJ:'.. i: * ~Ph,szcc: 2~c~rI'al Pars r(eqdlr~.=-n.rs -.- -- -. .-- I I -- Bxperlm~nt Yrlfct 2c:p:lt Di.:car Ions t:.:ctr,,.3~-2>rt:. .%.dl. : ' !on Z: ( 711 . ,or at "I : 7 (och) i>:. .. a 0 b.' X '9' x 11" I - ' L>SS Spec:zon.ts.~ -1 Xars Spec~ror~rler .Is Y 1.x 7.5 x 12 led . li -20 to 1 "11 I' I I. j .,>I : I. Pave* bYli -0 I +$ :P ki.tlul ~;1'.PS ':*:r. , rrit. Z. Drpl.;. Stal Corer (a,?.I I .%&nrLorattr (AC) Sc nsor -(3 rcqd) 1 Elrcrr~?rics 5 .\naIq 1 UPS :izk*i - - rn~b ' ' 5 >..- .---- - >a~i~riamelrr(DC) . Srnrur 1 .J ."x II" r 1.5" -50 co Arfc 01 Ppjnnin. ..r.llit* corm1 I to pls~aa! rrllpric Elertronics . -L ! t" x I>" ,, 1" 28-+ 1.5 -5 ro L Jij~vdl 2- LPS I 1i.x-d 3 *+IJ" 2. &lit.. Sequrnct. Po.itio- know!. J,. :P. -; G11n caniro: ~1x6~4Probe 0.5 -"*guIreu ..:I: ILCI:U. Lnoilc..'.. itr.!r;. run- .'. anel,.: lo, ro p,>sltiar irtril..dr. n. : W. -- - Ele~trlcFrrll hcel,?.~*(_ .c: , -.. . * .+is a? .jlt~nIr. . ':.??i:e ~..rrrl / ( " . -:+' . " . ., 1, pla..%...I , .i:pt:, I LCOTC~I Elrrrrunlrb -.;su . . . 2" L~:, 1 -)a, t<>-,,. dl . Jt. R1.5 -- ~- .- >k?roralda . laprcr/flajh lnrt. .- 4'' I .-"dll Dlciz~l J.3 BPS 4% Pr.!?rr..! cr:c.nt.:t. ..- i .rag tr,a (",i;birr iular .~rIai . 1.- .:. r my Jrirrrativn (:i nezqssari 16 tat. At1 1Lt11:*, rs,niic?,e zcqulrrd hscr R.tr~rafiects,r Rrrioref1e;tor Art+ 1L '1 ;5 1 if" r .' : :;O,AQ -50 to Sonv - +qzlra - . rrr . rr..! rr:vsccp. vit',in -5tJ0 Earti lrU'ot rttr~-..!. c.l r Q 3-$5. eal. Orient. . ------?.0fke Sur,!ey .,o 2 (J 28:. 5 -1u to 11.5~ 1 ~z<.t..l sr UPS . 1,-..r:. r I.." 1. PT~c~m- ':: Ibc1.1 'e.r (') . * :~cl~Jts .~cpl~ - .J. -c.,a~~s~ :1,.5''! ' I' I-to-tip Tlter.wl Cont. Antelmr ill lb - - . . * &dtinel . - Lravlt) Gradiam tcr t r I . 1.1 1 1 i I- 28:- i -ir, 1. 'R DF-IJ~I attlrudc kn~zl~t:.. tu .\" Solar Lnrrr. tl~' .;LO Elrctrot;lc Ortest~r 10 u 6- -0 ." 1," Aftilu~r6 p~.':.v !.ls:ir1cs 7' UPS required. ~t:lt,:,;, *~IG~.~L;:5O - - . . . . - - - Low Enrri, lbl t DF~CSLOI!+ .ltl 1) 4,' x 7" )i ir" * it:ir~jc L t.,~:. :- istories rrcuirrd. Artltw?. r:.u.:.dne 25' 2:: A ?, n *c - ,.,w - - I""r,w L.IU,,-rrI CO.....O".,.IOY Z9.1~:IL.L-I (Cont'd) %St4 I,UNA!4 OORRIThL SClENLE - EXPERlEhT KE~U~KEXKTSS?.'L!bP Physical Electrical Data Pequirrmenta Attitude Control ,I Bit Uax Experinen.: Dimans iona TYPr Drv rfincrola Hruurks IR Uapping Radiometer I I I*bdidtor on radiometer sust viru bdionr tar Electronics rpac . (,om mounrrd probably) -- -- 20 0" x 6" x 13' shown. AtTitudc knoilcdpc of 8" x 6" x 13" -lo rt quirrd. 0 :40 BPS 1R Spectrnmntcr Spac troaecer 6 X fi 112'' X ? 2~210~ 1 P!,J~LB~ orbiqxl~i (min) Vert i4~i~1. [email protected] Roll SIC to .:quire sun 1:. rrr.sor 1 151 1 1 1 bfis/l~C 1 1 F.O.V. tor calil~rrrioi~:r lcrat Electronics - 8" x 8" i 12" 1x10 bits1 per ncc once; prefer oncr!day. 3 - orbit (max) -7- - 1 L: I*" x 14" dia 2a2n 1, Power On-Off Attit~dkknouledy. 2 . . 1 Analog 1 BPS Vrrt 1. Fllrrr Select re+uirtJ for posr fLirlt :: .-.t.z 0 - 5V (tatal) tor Xd:ti8pec:ral I I ----- I I 0 118" x 196" x 3" 23 0 1-2U r. +'>'I Z ' Digirai 20 JPS :. Powrr On - Standb> - j 2 ft. ndriqua Lrtu Interface with the esrSM display and control panel should he mini- mal. Of the experiments in the WMgrouping, power on/off switching is the only requirenur,nt. If the geochemistry exyeri- merits are considered as part of the WMgrouping, then an add- itional boom deployment function may be required. Subsatellite groupings of experiments can be operated bv use of the sub- satellite sequence system. Environmental control requirements identified to this point in the program show a concern for electromagnetic interference, principally caused by shultaneoue operation of experbents and supporting subsystems. Several requirements, peculiar to speci- fic experiments, are seen. The neutron albedo portion of the geochemistry qxpesimenr must be shielded from potential neutron sources, such as the SM propellant tanks. (Note: !Che propellants have not been positively identified as sources of background noise. However, an evaluation should be made.) The laser altim- eter will probably require so- spacecraft provided coolant. The instrument is not designed at this point to dissfpate all of the heat generated within the unit. Operation of the IR mappint: radiometer on the dark side of the moon will require a cold ref- erence source. Use of a radiator for this purpose is presently specified, which may present a difficult integration problem. The experiments listed in Table 111.2-3 will require some form of protection from spacecraft contamination. This is ?art&cularly true for those experiments operated on the CMM. C~ntamination mechanism and sources are not fully understood. However, thone experhnts operating at wavelengths approaching the depth of potential particle layers should be protected. Apertures and surfaces of all optical syatems should be shielded while not in we. 4 MACPT'#N.MARlErr# OO~PCIUATDOAI DENVER DIVISION PR34-113 Page 111-15 Table 111.2-3 r:?cperimetlt Con tmiriat ion Sunanary --- - ! ~xp?riment --.- ~equir~L merit -*-- C-----L .--- I . Metric and Stellar Shield lens; inhibft Reaction Control Sys- Cameras tem during exposure ' Panoramic Camera I Same as Wtric Laser Altimeter Shield aperture of transmitter and receiver I teleseopeh Mass Spectrometer . Shield entrance aperture from attitude con- I trol jets and known outgasstng , IR Mapping , Shield entrance aperture . I Radiometer I LR Spectrometer ' Protect telescope and mirrors Photoelectric j Shield optics I'hotometer Solar Wind Foil Protect from acidic atmospheres ; do not expose to rare gases UV Spectrometer Protect entrance aperture from attitude control jets and known outgassing Laser Shield surfaces from jet plume, outgassing LRetroref lector - The orbit requirements of most experiments arb, compatible with an 80km (approximate) , circular, low inclination oxb.it . While many of the experiments could obtain more data with a high inclination, long duration mission, the objectives of most experi- ments are substantially met using the Apollo orbit. Some excep- tions can be seen in Table 111.2-4. The crew requirements for operation of experiments in the WM grouping are summarized in Table 111.2-5. As seen in the table, with ths exceptio,~of EVA for data retrieval, the requirements on the astronauts are minimal. Crew requirements'for all, other experiments are shown in the experiment deseriptlons of Appendix A, should it be desired to include them in the CbSM grouping. Otherwise, these expriments are actuated automatically or on ground conmnand. Potential astronaut functions are therefore limited to deployment of experiments 'and/or subsatellites and switching on/off. Other functions such as experiment scientific data monitoring are considered to be options only at this the. bUARr#W MARJEITA OQRPORATJON DEN\ ER DIVISION --- - - Table 111.2-4 Experhen t Orb it Requirements Sum-iary . - -- f 1 - -r ---- Exper iuent A1 tltude Eccentricity Inclinstion ! Remzrks I IIF Trans ponder, 600, 100, 50Km ' O to 0.3, altitude 30, 50, 70, 85' 1 Des.Lres mltiple Laser Retroreflector dependent missions to I achieve objectives I AC ar,d DC Magnetom- 100, varying to circular changLng o0 Initial circular, eters, Plasma Probe lOOOKm to elliptical I 100Km; boost in ! steps to lOOOKm apogee Gravity Gradiometer 65Km or less . polar preferred, 15O acceptable I lerO I I 1 Table 111.2-5 I Crew Requirements Summary ---- .--- -.-- .- - . ------Experiment Func t ion -1 Netlic and Stellar Cameras Switch s tandby/on/off. Retrieve film at . experiment conclus icn . Panoramic Camera Same as 2uletrLc Cmera . Laser A1 t ime ter Switch standby/on/off RF Transponcier None required Photoelectric Photometer switch' standby' on/& f Solar Wifid Foil Actuate experiment boom deployment, fire cover squfbs. Retrieve foil ac experhent ' conclusion. I iI . MABrICV MAAICrrA CORPORATION DENVER DIVI$ION PR34-,113 Page 111-17 Deployment requirements are summarized in Table 111.2-6 Below. Of' .the experiments in the WMgrouping, only the solar wind foil requires deplopent . Tablc 111.2-6 Exper hent Jkployment Requirements Sumary 7 Experiment Requirement .-- --*. -. - . . - -- IR Ma~pingRadiometer Must extenc?, experiment so that I radiator can view space Microwave Radiometer ilust erect L-band antenna I. i (12l x 12: dimens ions) Electromagnetic Deploy two mutually perpendicular Measuremen ta 1500 ff, dipole antennas Noise Survey Deploy and extend "V" or dipole antennas (-60 to 120' length) Radar Imager Deploy L-band antenna (3 x 5 meter dimensions) Solar Wind Foil Deploy such that foil surface is not shielded from the sun AC and DC Magnetometers Deploy at least 10 feet from space- craft to minimize interference Must rotate experiments if space- craft is non-rotating Electric Field Deploy antenna mechanism, then extend antenna elements; antenna must spin Gravity Gradiometer Ikploy gravity meters 15 feet from spacecraft Operating timelines for all experiments are shown in Figure I11 .?-I. The tlmelines reflect operating periods which allow complete coverage of the lunar surface beneath the spacecraft or complete accomplishment of the experhefit (as far as mission .duration and inclination will permit). As seen from the figure, operation of the metric and panoramic cameras is not required every orbit. In fact, due to slow rotation of the moon and the wide fields of view (74'1, the film capacity is not approached for complete coverage of the area under a 15' inclined orbit of three days duration. Operation of the laser altimeter is shown as continuous, althcugh its prfme function is to obtain altitude for camera correlation. The accurate alclmetry data obtained can be used with other experiments. The radar Fmager and the microwave imager cannot operete concurrently. However, some mutual surface coverage can be achieved Yy operating the two. MAWYJN MARIEmR CORPORAnOAR CEN'JFR CIVISION Metric & Stellar Camera - H Panoramic Camera . . i Radar or Laser Altimeter -* Geochernf s try -- I Ranging Transponder -H-C-.lC-(lC--.lW 3. Solar Wind Foil Solar Oriented ! -I IR Happing Radiome cer - - - - - I- I Ricrowave Radicpeter _1 8- s- - a 4 -~ I IR Spectro~ecer t C4-- Photoelectric Photometer ,- - I- . - 'd cd (b P m W Bistatic Radar fD * -1 e-.--. - e4 -* - - .-- .-a .- .---. - .----.I - J H I H - H I Radar Inlager CI, I .- - b-4 - - i-----., aD W Spectrometer - - - - .-# - .- Figure 111.2-1 Typical Exper!.ment Oprating Timelines MARIIN MARtET7A CQRPOmATCOW DENVER ,DIVISION I Day 1 ---4~ay.2------gqJ5q v 'h Laser Retroflector Operate nearslde only (illuminated by laser from earth) . , I ---.C-.---.-r-- I I i AC Magnetometer I 4 ! i Electric Fie13 I 1 Electromagnetic Pleasure. ------. - -f;3 i ! i / Plasma Probe ----.-- -- -.-- ) Meteoroids ---- - . -- . .------.- - t f 1 Noise Survsy I i Low Energy Particles -0 m a- Solar Energetic Particies I ------. ------. ------r------lag I i j GravLty Gradi meter Continuous unless RFI problem 1 ------.- C ! l2F Reflectance ------I Pigure 111.2-1 Typical Experiment Oprafing Tiinelin,es (ccrntd) ~~34-113 Page 111-20 instruments a1 ternately on every other orbit s ince their Eields. of view will overlap. As seen from the scient5fic roadmap (:;ec- tfon TIT. 1) an IR radiometer and the microwave radiometer should operate concurrently. Ground se~vicing,checkout and launch requirements are not yell defined at this point, Some general observations can be made, however. An optical alignment facility must be provided to per- mit alignment of experiments to the spacecraft and to each.other. All experiments requiring precise allgnmenr will be delivered with an optical surface which defines the centerline of the field of view. Laboratory and checkout space which is temperature and humidity controlled should be provided. However, no requirement has been identified for clean rooms. There are a number of experiments which require deployment due to size and/or operation location on the spacecraft. An area will be requtred in which t~ checkout the deployment devices. Pyrotechnic simulators must be provided for vehicle checkout to simulate deployment. A combined systems test of experiments (or simulators) and the spacecraft should be planned, not only to confirm system oper- ating adequacy, but to test interference among simultaneously operated experiments. Experiment data should be recorded in these tests to be sure that no experiment affects the, data collected by any other. Contamination covers should be actuated during i'ne systems teat and during prelaunch activities. - Feculiar checkout and servicing requirements are shown in Table IS1 2-7 --- T~ble111.2-7 ---- Pecul far Ground Checkout and Servicing Requirements Exper hnt Requir-ment L - -- Metric and Stellar, Load film within 3 daye of Paunch'. Panoramic Cameras I4aintain film +lwrperatures.1 Check intervalme ter , camera operat ion. Laser Altimeter Provide ground cooling if experiment operated more than a few cycles during checkau t ., IR Mapping Radiometer Provide cryogenfcs for detector cooling (or cryogenic heat sin,k for experiment radiator) I AC knd DC "Magnetically clean" area required I Magnetometers I far calibration, checkout. Possibly 1 1 use manufacturer '3 facility during I I experiment acceptawe for this pur- pose. Additional study required as necessity of degaussing satelltte.--- MaIIr0AP MR ffI.C'-4 0 0RPOIP4 r8OAI DENVER DIVISION PR34-113 Page 111-21 Peculiar subsys tern and support requirements are suwrized in Table 111.2-d. Reference should also be made to the other tables of Section 111, since some peculiar requirements are s,hown there and arc not repeated below, --- --.-.- -- .------,.L- -.,--.--.- _- - -- Table 311.2-H - r Peculiar Subsys tern and Support Requirements I Laser Altimeter Spacecraft coolant at 15 + 7O C required. 1 Mass Spectrometer , Controlled emiosion from lunar eur- I face desired (not required). I RF Trans ponder, ' Spacecraft orbit and attitude should be Laser Retroreflector unperturbed by onboard systems. I AC and DC Antennas must be spinning during oper- I I Magrletometers , ation--spinning subsatellite candidates. 1 Plasma Probe, Plasma probe requires accurate sensor I Electric Field for knowledge of spacecraft-sun line. I ! Metric and S t:ellar, frtervalometer required, either using Panoramic' Cameras sp..cecraf t sequence sys ten1 or. incf uded in experiment. i I Micrawave ad iamc Ler Requires 100 pps timing signal from spacecraft . 1 0pt:onal Operat ion: Radar Imager If operated in high resolution focused mode, fL1~recorder (or ability to handle 10' ' BPS) required. EXA will then be required to retrieve film. Microwave Radiometer I If included in C&SM grouping, provision of on-board processing and display will permit astronaut to make observati bns and adjust antenna scan, FOV, and gain to obtain data on areas of particular 1-nterest. -, Experiments not operated manually will require control by the spacecraft sequence system. Furthel-more, based on past experience, interface with opacecraf 1: t fm3.ng will be required for experiments in addltian to the microwave radiometer lnsted above. IV EXPERIMENT COST iWXD AVAILABILI!! ' A summary of cost and schedule data is presented in Figure IV-1. These data are based upon the detail cost and schedule estimate sheets, for each expeziiuent, found in Appendix B of thLa repbrt. These data are presented by experiment and portray a cost range fron an es thted minimum to a maximum. SLr~cethe quarltity of hardware varies among experiments,.the cost should be related to the hazdwqre required per the detailed estimate. Estimates have been influenced by the confidence level qf the organization(s) contacted to translate the experiment definition into cost in- formation. Fkny experiments are well defined and hardware avail- . able, while other experiments, not as well defined, require a degree of con jecbure ; theref ore, the coat "range" allows for this adjustment. The schedule spans reflect quarters necessary from the order.to delivery of the flight article(s) for each experiment. For reasons stated above a schedule minilhum and maximum the is presented. Various NASA agencies, university personnel, and manufacturers were contacted in gathering these data. In some cases it was detenniucd that simllar items. are (1) in use on an existing satellite, (2),about to be launched on a satellite within the next six months, (33 in the development stage for a funded pro- gram, or (4) undergoing a design and feasibility study. One example of existing hardware is the Laser Altimeter built by Control Data Corporation and scheduled to be placed in an Earth Orbit during the spring of 1969. At the other extreme, the Gravity Grisdiometer Experiment is still undergoing a feasibility study to determine the technique to be,used. M&IPIIIV MARIeWA Q0HP01FLsnON DENVEA DIVISION PR34-113 Page V-1 V MISS JON Cfj?T:Y.!'IIATNTS v.1 L,. - I:CZ:S!! a11d Su1)~at~lllt~-- - TIIc' II~IS~C1u1iar Lanci- ing mission is not com?atfble c*l! th total lunar surface survey by S:I 1 ., I-. I~IL~sur'..~c~ covc.rar:cl tror:) 'I Ltln;lr L:~~~tli~~;; orbit wjth u;, * -r 2!100 pounds of 1:,1yload capabilitv and a miiximum ti.mc of 3 tl:\vs LS -1 very s~nal.' fraction of *he total. Therefore, subsntcl 1 l.tc. st s t cSals should bc colrsidcred to cnlii~nccthe covcBragca througi~h.11;h i IIC! l~iiif!o! R i11ld Ion\; sLay timcs or the cxt endcd vi~win~o! cerf.I ;n lun;ir slits should be elevated in value above the tot:1l surfac:c covc.ral;c. objcctives stated at the start of this ,)roeram. The coverage of the lunar surface is discr~esedin terms of general area coverage, specific orbit accessibility and the impact on experiment utilization and specilic target coverage. Subsatel-. lite requirements and cnpabil~itiesare discussed when they vary fsom the basic (XSt.1. The portion of the lunar surface that is sqd~lablefor viewing is presented in Figure V.l-i as a functicn of orbit inclination and lunar stay time. A polar orbit w,ith a stay time of about 14 days is rcquired f3r complete surface coverage. The coverage areas for the 10.1 and 14.3 da.y total time missions (1 day and 3 day stay time, respectivsly) are noted along with the 14 day . stay time data. An orbit altitude of 60 n m ia'assumed. Se1c.c t t hc. 1'1 tlc~p~i.i~inc 1 i.11'1t ion t113t is requirecl ,Lor I hc Abulfeda landing and estabLCah the lighting such that it is favorable to thtb landing mission. The lunar surface accessibility for the 10.1 d~y(1 dnv stay) and 14.3 day '(3 day stay) missions are shown in FIgurc V.? -2. The effect of lighting on e-~pesiment utilization is shown. Experiment utilization is shown in Figp~re V.1-3. No restrictions have been assumed on the available time due to crew duties or spacecraft pointing capabilities and the data represent a maximum. Very low percentages of the' lunar sur- face are available to the C&SCI t~nn nc7nlinol landirl;: mission. To imprqve lunar surface coverage a subsatellite could be deployed at high inclinations. The size of the subsatellite is a function of the propulsion system requirements. The subsatellite propulsive . requirements to achieve a high inclina.tion from the initial C&SN orbit irclination can be defined from the turn requirements. Figure V.l-4 presents the minimum turn raquirements to achieve a given orbit inclination from any f:&Shl orbit. An orhLt aLtitvdt. oL- 60 11 m is assumed for the subsatellite data. The turn angle requirement can be converted to a velocity increment requirement as shown in I4DAY STAY TIYE 100- - W LD 80- 4 Cr m W > 0 60- C) - 3 3AY --- I DAY --- -. - -- __-I-I II 0 38 60 90 INC!.l NATION-DEG. Figure V. 1- 1 Lunar Surface Coverage 1 I L -j-6o0 1 77 I I CAMERA I !I t I I ~T\S'GEOCHEM.(PFD~ ; I I I-] PHOTOM'R (PF~,) / . v-7 I NO ' I 1'4 ILLUMINATlON ! A3ULFEi)A I- - 3. I oi . . 15 Q@ 96 0 9 0' 1 50' a -FI WEST EAST B P 3Q W (P C- I e P 1 t-' W W Figure V.1-2 Areas Accessible for Orbital I MAIPrfN MARlEYTA CORPORATION :.t...-r. : . Inclination: 15O - - eurleet Larding Site: Abul feda ------Subsolar-pgin t- --'-- - . --- -I------.- --- XCoy~rngc - , Experirccn~al -- - 11.1 ~iminat'lon-(*) - . --- 1.01: day:; - -I --3. days Equi?mcnt Acccptnhlc I Preferred I Acccp~ Prcf ( Acccp: PreE . ------. , ! -=.- . ------. ---. I ------ ~ndar/LascrAleilcrcr ~?YO -I I 7.0 -- Gcochcnis try i -tC5O light side I 7-0 1.2 - I-- /--- I - Rangins Spondcr a/~ I I -- -- 4- -- -- I i Solar Wind Foil ------.-- - rR Ikippi:ig XadLona tcr -?-90° dark side 1 1.2 ' 3.5 - - - -- I -.- .- - --1-90° light side j 3.5 I-- i9C0 ligbi aide ! +:Go light side 0.1 ; 3.5 0,2 - I - -- - . -6 - --. .-a Bis tatlc kzdar --anyd . ! 2.5 .----- 7.0 I I I U.V. k.ager : any 4 I___2.5 7 1 I I I Figure V. 1-3 Ilxperiment Utilization iLW= IWCLINATION OF COMMAND 143CC4; LE INCLINATION OF SUBSATE LLI TE ADEG ~i'gureV. 1-4 Subsatellite Turn Analysis MARTIN MaRIETTZ CORPORarKOAP a .. . ? K 1. b', c,-I Figure V.1-5. Thcsd data, in turn, can be presented in terms of the mass ratio rcquired to develop a given velocity increment. The mass ratio is expressed as a precent of the subsatellitc weight that must be exper~dedin propellant to achieve a given turr, in Figure V. 1-6. Two specific impulse values are noted for compar- ison, The subsatellite mass shown does not include the ejection mechanism to dep!.oy it from' the C&Shf. 'I'hc C&Sbl ccrvr.1-a ,c. pictilrc lor general surface char:ictcristics has been noted. Subsetell~teswill improve the coverage characteris- tics of the system. Now, consider the coverage ot specific targets or landing sites 'for multiple viewing and study. Figure V.l-7 shows the general coverage problem for a circular orbit at 60 n m. A maximum sensor viewing half angle,/Yof 37 degrees is no-ted. Coverage of any chosen landing site is given in Figure V.l-8 as a function of the latitude of the landing site. A sensor view- ing half angle of 37 degrees is assumed. The n~isaionplan for this coverage is such that the inclinatio~of the orbiting C&SN is only slightly (less than 2 degrees) higher than the latitude of the landing site. ~iso,the illumination angle a.t the time of lunar ' module landing ( @ = 70°-83O) is also approximately satisfactory for camera viewing ( = 4S0-75O) . The illumii~ationangle becomes smaller at the rate of 12.2 degrees/day due to surface rotation. Thus the ianding site receives satisfactory Llluqlnation not on1.y at lunar module landing but at all times thereafter for the stay times considered . Coverage of the lunar landing site starts when the C&SH passes ap- proximately over the landing site and continues until the landing site passes out of the field oE view of the sensor as the orbit moves westward. The c&SPI' orbit is at an inclination s Ll1,htLy greater than the latitude of the landing site. The number of orbits passing within sensor coverage of the landing site (based on 74 degrees field of view) ranges from approximately 49 orbits for a site at 10 degrees north or south latitude to sbme 39 orbtts at 45 degrees north or south latitude. Si~ce,however, the 3 day surface stay time of the lunar module limits the number of crbits to some 36 orbits, it id evident that the landing aite is withfa view of the C&Shl-motnted camera part of each orbit. ilsr entire stay time. Figure V.1-8 also relates the number of orbits to the amount of viewing time of the camera. For the 3 day surface stay time (with 36 orbits), the amount of viewing time varies with site latitude and inclination of the orbit. For all landing sites the total viewing time is about 1 hour for a 3 day stay time . ALTITUDE - 60 N.MI. IoGoor TURN ANGLE-DEG- Figure V. 1-5 Subsatellite Analysis, A V MARTIN MARIETTA CORPORATION DFnlVtR 1J'vISION TS P - S-EC, Figure V.1-6 Subsatellite Analysis, Propellant Required C Circiunference of ntoctn -= 5900 n mi T Time of satellite orbit =- 2 hr 6 Velt.~itsc.f satellite lIFOC/hr i Inclinationoforbit .-., T & angle cf sensor ficlt of view = 37' @ Equivalent arc on m~r~surface of sensor coverage = ?.8', -46 n mi h Apparent motioc of node to west = l.lC/orbit, -18.0 n mi/orbit Figure V.l-7 Sensor Coverage of Lunar Site MARTIN MMRIETlA CORPORANON 1)t NVER DtVlSlON SENSOR :CAMERA FIE LD OF VIEW: 74'(87 N-MI. AT 60 N.MI.ALTITUDE) LATITUDE OF SITE DEG Figure V. 1-8 C&SN Camera Coverpge of Landing Site MARTIN MARIETTA CORPORATION is I~IVI'.IOW PR34-113 Page V-11 The previous coverage numbears assuned that the initial pass is directly over the landing site or target. In this case only one half of the ground viewing circlo is available for the coverage sweep with time? since the target: starts in the middle of the circle and will not be viewed once it has moved beyond the circle or returned to the center where nominal LM pickup is achieved. Now consider the case where the target is started anywhere wlthin the coverage circle or as a maximum on the edge of the circle. This condition tands to maximize the accessibi.lity of the target. The site longitude and lhtituda will fall within view of u p,ivan orbit inclination and nodal position if the dist~ncebetween the site and the orbit (angle ) is less than the radiua of the viawing circle (W ), The smallest angle, , between any orbit and any site is computed as & sl;, rJ&dL --kn/~jh,~5Lnn where d is the site latitude, i is the orbit inclination, and A is the difference between the orbit node, XL) and the site longitude, ,4 ( C\ R = I lo - 1 \ ) . If I (P I L Y , the arc of the lunar surface cut off by Y ineludes the target site. is computed PS where R,, is the radius of the moon and h is the altitude crf the orbiting vehicle. Figure V.l-9 gives the maximt~m number of orbits with sensor co-ler- age as a function of inclination and site latitude. For maximum coverage, the inclination should be within several degrees of the site latitude. Figure V.l-9 is based on a sensor field of view of 74' and an orbiting vehicle altitude of 60 n m. When the initial orbit is eastward from the site by exactly degrees ( @ = \il ), the coverage is a maximum and Figure V.l-9 can be used. However, when this is not the case, some modifica- tions need to be made. For instance, when the initial orbit is over the site, (Q = 0 and the coverage C is approximately one- half of CmX. In general, C = 2 where C is the coverage in orbits, and 0 is negative if the initial orbit is westward from tlie site. The viewing time for the orbiting vehicle can be obtained from Figure V.1-10 where the parameter CmX is taken from Figure V.1-9 and is computed as above. If the orbiting vehicle is a sub- MARTIN MARJE77A CORPORATION DENVER DIVISION UDEm 60NMI. Of- V IEWm 74' LATITUDE OF SITE, DEG Figure V.1-9 Extended Sensor Coverage MARTfN MARfEITR CORPORATCON DENVER DIVISIOW Figure V.l-10 Extended Sensor Viewing Time PR34-113 Page V-14 satellite, there is no stay tima limitation on viewing tima aa there would be for the LMM. Also, ~Ercerappuoximo tely 14 dayar, tha initial conditions are repeatad but from @. southerly dirac- tion if the original orbits had bean from o northerly direction, and convarsaly. MARTIN MARIIWA (CORPORAWON DENVER DIVISION PR34-.I13 Page V-15 V.2 PayI~stZl/DeltaV ConeLdezatio - The payload capability is dependent upon the srelaocion of a speoific landing site and ds- sion date. The maximm aslpability is 2800 pounds. This capa- bility Fs r~cluaedto 2300 pounds for a landing at AbulP~da(in- olhation about 15 dagraes) . The tradeof f s between payload and the udssion variables are disaussed in the following pa,rsrgraphs. Xh@ payload 5s a function of the velocity increments required for lunar orbit insert ion (ZOX) , technique of LM pickup and trans - earth insartion (TBI). These velocity Fnorements, in turn, are a Cunction of the tznnslunar trip tb,and lunar orbit notlal posi- tion and inclina.tion required for the mission. For a mirssian whare the principal objective is a lunar landing the nodal posi- tion and inclination are established by the landing site location and technique of LM piokup. For any case the velocity increment is minimized by utilizing a three impulse technique that employs a highly elliptic initial orbit for plane changes before achiev- ing the final circular orbit. Consf,der the landing site locration and its effect on the mission. Twenty landing sites have been identified and are shown in Figure V. 2-1. These sites are: Littrow 11. Schrster's Valley Hyg hus 12. Aristarchus Tyc ho 13. Dionys ius Gas s end i 14. Sea of Alexander Censarinus 15. Tobias Mayer Abulfeda 16. Copernicus C Rim Bode 17. Marius Hills MG-~tingC 18. Hadley Rille Pra Mauro 19. Alphonsus Hipparchus 20. Ilarbiniger Mountains The first four sites have been designaped by NASA-MSC as those of primary interest. Data for Abulfeda are included as a comparison baseline to take advantage of the extens ive site analys is already completed for this site. For specific trip times, a nodal posi- tion can be defined that will yield a minimum LO1 velocity incre- ment. Similarly the maxfqum LO1 velocity increment is associated with a specific nodal position. The positions of the descending nodes for these limiting cases and a three impulse insertion tech- nique are also shown in Figure V.2-1. The surface traces shown are representative of 60 n m orbits at an inclination of 30 deg. Data are shown in Figure V.2-1 for total translunar trip times of 84 hrs, 96 hrs and 108 hrs. This total thincludes the time in the arrival ellipse that is a part of the three impulse tech- nique and corresponds to actual trip times of 66 hrs, 72 hrs and MARrllV MARlEmA UORPORAnCIN DENVER DIVISION DEG. NEST LONGITUDE DEG- EAST figure V.2-1 Site Analysis MARrIM MARIETrA DORPfORATfON DEUVER TIIVIStOW PR34-113 Page V-17 78 hrs respectively. The nodal positions will not change with fnalination and can be either ascending or descending nodes. The corresponding ascending node picture can be achieved by a rota- tion about the lunar equator. These data do not reflect a re- quirement for free return trajectories since the study was not constrained in this way. The magnitude of the WI velocity increment (three impulse tech- nique) is presented in Figure V.2-2 as a function of inclination. The data are presexrted for nodal positions yielding minimum and maximum conditions and it: is apparent that the nodal position is a significant factor in the energy requirements at high inclina- tions. Trip time is also a factor with increasing trip times resulting in reduced energy requirements. The transearth insertion (TEI) velocity increments will be s hi- lar to the LO1 fncrentents. For stay times of about three days the maximum TEI increment will occur with the maximum LOT, incre- ment and the minimum will occur together. The orbiting weight equivalent to a change in the velocity increment can be computed on the bas is of about seven pounds per fps for LO1 and three pounds per fps for TEI. Specific site analysis, then, will define the velocity requirements and ehe attendant payload capability in lunar orbit. Two approaches that are different in the LM pickup technique have been Fnves t igated in determining the pay1 oad capr* ability. The first method considered consists of picking an initial inclin- ation that passes through the lunar site of landing and utilizing a plane change for LM pickup that increases with stay time. !She mission profile would show a wait of i8 hrs after LO1 for checlc- out and landing preparations. At the end of the wait the site would pass under the orbit plane and the LM would descend and land. A stay time of three days is utilized and the lunar rata- tion has moved the landing site out of the orbit plane. A final orbit inclination is determined that passes through the landing site and intersects the orbit plane at a minimum turn angle. This is the plane change required for LM pickup. Also, the initial and final orbit inclinations are determined (graphically in most cases) as well as the initial and final orbit nodal positions. The plane change maneuver is presented in Figure V.2-3 as a function of stay time a~dCM orbit inclination for a lunar orbit altitude of 60 n.m. Two specific stay times are noted to provide a comparison between the requirements for the 10.1 day and the 14-3 day mission. The landing site is assumed to be Abulfeda. MARrlN MARJETTA CORPORArlON OENVER DIVISION Figure 77.2-2 Site Anslgsis -bV MARNN MAHI-A CORPORATION 3E' Figure V. 2-3 Plane Change ~e~uirements MARrIN MARIEWA CORPORANON DENVER DIVI>ION PR34-113 Page V-20 The LO1 land TEI velocity increments are evaluated Prom Reference 1 which assumes that the three impulse technique is used. Pro- vision is made for the pickup plane c!~ange and adjustment is made for midcourse corrections and a L2.I rescue mode sf operation. The final velocity requirements set the propellant needed and the resulting payload available for scientific experiments. For a landing at Abulfeda this method yields a payload of about 1700 pounds. This appco~challows for any orbit abort within its nominal per- formance. The plane change increases to a maximum at the nominal departure time! therefore, early Dl pickup requires less energy. However, an abort results in a shorter mission time and nodal pos- itions displaced from the nominal case. Depending on the choice of landing sites the new departure node could result in either higher or lower TEX velocity increments. The capability of any orbit abort, then, will require a payload penalty for some land- Lng sites or possibly the provision for waiting in lunar orbit for the nominal departure time. The second method is faster computationally than t'he first and is generally less restrictive in terms of payload. In this case, an orbit is determined that passes over the landing site both initially and at the nominal IM pickup time. No pickup plane change is required for the nominal flight plan, However, a small plane change is required for any orbit abort. The initial and final orbit inclinations are the same and slightly greater than the latitude of the landing site. This condition is noted in the fol- lowing sketch. \- Lunar Equator MARrlN NPARtmWA C)ORPQRAZION DENVER DIVISION PR34-113 Page V-21 The initial and final nodal positions can be determined from simple expressions that are dependent only on the longitude of the landing site and the lunar surface stay timas. The equations for the inclination required are: Sin 1 = Sin& (inclination) Sin w where w = (90-X) and SinxeCosd Sin ah 2 The initial node can b& determined from: 20= Longiite - 98 + (uncorrected for wait time after 2 lJ3 1) The final node: Jt = Long.s ite -90- -AX (uncorrected for wait before TEI) 2 Knowing the inclination and initial and final nodal pos it ions, the velocity increments for LO1 and TEI can be determined from the parametric data in Reference 1. The payload capabllity for a specif Fc site can be defined. This approach is more consistent with the data and assumptions of the existing Abulfeda site analysis and is the technique used for the imrestigations of this report. Any orbital abort requires a small plane change which is assumed to be achieved by utilizing the ZIM rescue mode propellant in the C&SM. The abort will shift the departure nodes and will affect the TEI velocity increment. Depending on the site chosen, the TEI velocity increment will increase or decrease and may require an additional payload pen- alty. This condition can be alleviated by waiting in the lunar orbit for the nominal departure time. Based on the second method, specific payload penalty curves have been determined and are presented in Figure V.2-4. A stay time of three days is assumed and nominal LO1 and TEI velocity incre- ments arutilpzed. The node pos itions allow for the initial wait between LO1 and landing and the final wait between LM pick- up and TEI. Any site along the equator is accessible without pen- alty. As the site latitude is increased north or south, the cost in tern of payload penalty increases. A basic capability of 2800 pounds of scientific payload is assumed for equatorial orbits. MARrtN MARIilTA QORPORArtON DEIdVER DIVISION Figure v.2-4 Site Analysis- Payload MARTIN MARIE-A CORPORATION DENVER DIVISION The, payload penal-ty is subtracted from this value. At Abulfeda a payload penalty of 500 pounds is noted resulting in a payload capability of 2300 pounds. The xaximum payload penalty 02 1680 poundr; ref l(?cts the maximum propulsive capability of the C&SM based on tht? weight as preserlted 211 ?'n:)lt~V. .?-1,. I\~I#.I:, OII 1 s i \It> of the 1b80 pound envelope are. not accessible under these con- strnints, The data are cross-plotted in Figure V.2-5 to indicate t!ls effect of inclination on payload penalty for the must Fwror- able landing site location and least favorable landing site long- itudinal location. It should be noted that the LM pickup tech- tliquc chosen establishes the inclination for a specific site at an angle sl.ightly larger than the value of latitude at the site. Lunar orbit inclination is not freely variable in this case. Should other inclinations be desired the landing site must be changed or the LM pickup technique must be changed resulting in new LOX and TEI values. Specific payload data are presented in Table V.2-1 for an Ahul- feda lai~ding. The data assume a heavy IN weight to account: for the extended mission t-ime (three day surface stay) . This assump- tion requires the C&SM to bring the LM to an 8 n m orbit for de- scent and landing and n.ecessit~tesa C&SM transfer from the ini- tial. 60 n m circular orbit to the 8 n m circular orbit and back to the 60 n m circular orbit. Two sets of numbers are shown to compare MMC and NAR data. The only significant differences are in the LO1 and TdI velocity increments which can be attributed to differences in the accuracy of reading parametric data (MMC) and specific case coinputations (NAR). NASA-MC hzs directed the use of the NAR payload data. Reference : 1. "Lunar Orbital Survey Missions ," Final Report. Volumes I, 11, and 1x1, Lockheed ~issiles-andSpace company, 16 ~anuary 1967. r MAX. P8OPkLLAfL"T 0 10 20 3 0 40 50 60 70 80 90 INCLINATION OEG MARTIN MARIETTA CORPORATION LE\.&F: ;>,, >I021 Table V.2-1 Payload Comparison Abulfeda I M MC NAR I - (DEG) 14.85 1 5.00 ~V(LOT)- F PS 2993- 2871 - AV(TEII -- FPS 362 5 3456 1 I W(IM) - LB 3380 5 33805 I W(CSM) - LB 25000 2 5000 i W(INJ) - LB 99165 99165 ISP - SEC 31 3 31 3 290 290 AV(HVY-LMI- F PS I W(PAYLOAD)-L8 19 00 2300 MARrIN MARIE-A CORPORATION OENLER I):#S'ON PR34-113 Page V-26 The impact of' one-mm CM operations in orbit durlng phase 11.0, 1,unar' . urt ucr2 stay, WPLB inveatigatcd for both the 10.1 day md th~Ilk. j duy n~i:;nion.-. Very limited crew time, only :'-:j hour6 would Lie r~vnl!tib!r durinv the 10.1 day lnissjon for sci~nccrxperi- :i.enti; . '1'11c. ill. i lirty nllsn ion ~lrovideea nominal 7;' hour6 in Lunar ,trbii :111t1 1)c"nltlL:; periodic cyclco !'or !;/C eyutems mana~ernr*nt r~nd iP1 :;~ipl~)t't1 rr~k~I o provide L'-15 hours crew time for idrntil'icd ID, cxi~rirnmnttutiktl . [:fintiIda te LO.; experimento require crrw part- icip~tiolit111~t Invo1vc.ti primarily experiment 8w LLcllin;: Laslts cllld EVA Clttt~r~tt'iev~l at thr end of the mission. Manual control ol' 311 expcrin!etlt on-of!' cycling is not recommended, and automaLed experiment control to assist the crew is necessary. The recommended S/C attitude is "stremliued" (x-axis paralle 1 to velocity vector ) with the CGM -2axis at local vertical. This choice permits crew visibility of the groand track, but requires canted orientation and mounting of Sector I LOCj experiments. The most feasible time for EVA dats. retrieval is during trans-earth coast. The EMU/PTSS, already provided for the Lunar Surface mission, should be reused to minitcize overall misaion weight and stowage requirements. Since primary mission zontingencies ray cause cancellation of EVA, alter- nate data retrieval techniques should be investigated. V.3-1 Crew IVA - The Lunar Orbital Science (~06)mission super- imposes selected experiment management tasks on the reference ApoLlo Miss ion G ( lunar landing) crew operations. Throughout this study the primary obJective of a succe~sfullunar landing mission was considered for new landing sites and experiments. The LOS crew operations must complement the lunar landing mission while satisfying the science experiments requirements. Crew IVA is restricted for the science experiments to one-man CM operations during lunar orbit while the Lfi! and two crewmen are on the lunar surface. During one-man CM operateions in lunar orbit, the CM pilot must manage the C&SM, support the LM, and support the LX>S experiments while continuing the established crew work/rest cycle at Lunar Orbit Insertion (LOI). Critical mission events just before LM- undocking and after the lunar orbital rendezvous period preclude major ad justments to the required SICtasks and crew work/rest cycle to accommodate JX)S operations. Crew experiment tasks are dependent upon the scientific experiment grouping considered. For the LOS mission, the experiment groupings fall into two categories: C&SM mounted, and sub-satellite mounted. After a sub-satellite is placed in lunar orbit (post-LOI), the CM pilot would have no active control over it. Any required sub- PR34-113 Page V.-27 ::ate 11 ite t'un *tiona would be con trolled by K;FN/DSTF commands, 'l'tlc rc l'~)rc, -no ldlno i F, in the study was placed on CWM mounted +-xj-n?r.irn~t~t:,. In order to Insure the greatest experiment grouping 1'1rxi11ilit.y, clrew,'~*xpcrimentoperationc and interfaces were inveo ti.- r:ateil l'or nl 1 #*andid~te IOS experiments. The crcw/exp?pcria~cnt tasks are e~mrnari;~.edit1 Table V. 3-1 for C&SM -mounted cxpcr~mcnts. Table '4.3-2 shows ttlese int.erferces for sub-eate llite-mounted experiments should they he placed in the CUM, Figure V. 3-1, Gross Mission Timeline for Lunar Orbital Period, shows those mission and experiment events that impact the CM pilt)~, The bar graph indicates those critical mission periods when the CM pilot would be fully occupied. Obviously, independent switch functions for 0n/0ff cycling of experiments are excessive for manual control and monitoring by him, even during lass critical mission periods. Those experiments requiring this regular cycling should be provided with on-board timers pre-set by the CM pilot or by ground commands to relieve him of these repetitive operations. Crew participation in 0UCh activities should be limited to primarily backup or select- ive control and periodic updating of automatic sequencing functt ons . The regularity of experiment on-off cycling lends itself to various solutions from relatively simple intervalome ter type controls monitored by the CM pilot to computer-sequenced operations requiring little crew participation. MSFN/DSIF command control of experiment on-off functions may provide a simpler alternative to completely automating the on-board experiment. Experiment sequential timing functions could be integrated iuto the individual experlmenls or made part of selected experiment group sequencing functions. Integrated experiment group control would be preferable if the CM pilot is required to update, set or otherwise perform simultaneous control functions. Controls and displays should be portable utilizing inflight disconnects that could be abandoned in the LM after completion of the Lunar experimentation program. Experiment panels may need to be carried in from the LM and ~ountedtemporarily at the center 'couch area near the main C&D panel to permit simultaneous operations by the CM pi10 t of S/C and experiment operations while maintaining out-tne-window visibility of the Lunar surface. All 7VA for the LOS mission is based on evaluation of m.inim~~m necessary crew interface with the LOS expriments bec~useof the crew tinie commitments to the primary lunar landing miss ion. Increased crew utilization in some areas is desirable including more effective camera film uoage. The metric cameras carry considerably more film than can be usefully exposed over the ground track as now planned (each 4th orbit). Crew options to photograph targets of opportunity on in-between orbits would permit increased usage of the available 1200 frames from 50% to 75$-80$. MAIPrfN MARIPZIA COWPORAIfON DENVER DIVISION TABLE V.3-1. CREW/CM/SM/EXPERLMEK~TASRS TAIiLE V.3-2 - CREW/EXPERIMEET TASKS ALIGN. OR CRElV CRE3 KZ?. DATA ;.TTZT'..E % FXER~ ORIENTATION c.~L~,~,'DSOWTROLS~ -SE]& * E. F~T~~ALP^IT!PZT~EQ '15 DISTWYS G'TS. ME t~oroids Space (Not No Pdr ClN,'OFF Non~ None 1 z;ir--S.%?e:lite 1 ~ritical; .I_ . 7eocher istry Lunar surface No Pvr 31i;cFF I None Sone ;-~:-cstellite ) 53~-ieplcy JI I- . 1R Ivlapping Radiorceter Lunar Surface No Deploy Exper. None None -.,..-,,-se%clll:e, 1 Mass Spectrometer Lunar Surface No Pwr OH,'OFT Nong Rone -,:..-v -saielli-ze ) Seal Cover *ploy IR Spectrometer Lunar Surface No FW ON/OFF None Nosle ( Suk-sate llie ) M'd Radiometer Lunar Surface Nc Pdr ON/'OFF Xo~e None :Su3-saLLellice) RF' Reflectance Ianar Surface No Eonc None Kone .I=.. - Jk-s~tellite : BI-Static Rad~r Lunar Surface No None W one None !'-.Yo-satellite Radar Altimeter Lunar Surface N3 PCC ON/CFF None None {51k-satelLitx) Antenna deploy Electroeagnetic Xezs. Lunar Surfece No Patenna deploy Ncne None '3k-satellite 1 Noise Survey ~/.4 No Pdr ON/OFF fio~ie l?om ' :..-,D-satelli$e) EL Change Ant. Orie~t 1 !& errection i Magnetometer ( DC ) h~ No :done None &one '3k-satellite ) TABLE V. 3-2 - C!lEW/EXPbKIMENT TASKS - (Continued) TLQGZT ALIGN. OR ?E.C VCX. DATA 3TI'TTUT)F 9c EXPERmNT ORIEN'f'ATIOX CAL RE!Q1!2 ,I CONTROLS/ FFQ'F . FI3TIFVAL PO: NTifir; RFB '3. DISPLAYS TG'TS. Magnetometer (42) I Any No None None Scne f5ub-s:i~~Llfze' Plasma Probe .by Yes Pwr ON/'OFF Y~ne Vane ( IUD-aecel' i"~) Data Recor? Der~OY I Electric Field knar 3urlace No ?wr ON/OW None 7i3~e !~l~b-srt41: i te ' Antenna deploy Low Energy Particles N/A No ~wrON/OFF Hone :{one (sub-s~+fllite! Deploy Covers Solar Energetic N/A No Pwr OM/OFF None None (~ub-stitcl?%e ! Particles Deploy Covers MARrIN MaRIE-A CORPORATION DENVER DIVISION Figure V.3-1 CrewlExperiment Timeline for Lunar Orbital Period MARTIN MARIE-A CORPORATION DENVER DIVISION PR34-1,13 Page 32 Expc~ri~llent data 111dy b(1 ollt~arlc~dby increasinh L he crew participation in the Mdcrowave Radiometer experiment. Color maps of the Lunar surf~cevic3wed o:i n color Cathode Ray Tube in real time by the CM Pilot could rPveo1 interesting phenomena that would be selectively ir!ves tiyatr.;l I ri crrc*nter dctai 'l an oucceeding orbits. Tl~roughre y~ti- tivc ol~servut!ons, the c:M Pilot could distinguish small cht~n~es%ti color ( Fndlcatiny; temp-rature chrznges) wd correlate these with inl'ormirtion t'rom optical or radar senaors. An on-hoard color CRT monitor would be required in the CM as part of the carry-on control and display. V. 3-2 Crew Visibility & S acecrcft Orientation - The recommended S/C orientation is '"streamlined-!-7 Local vertical with -I, axis toward nadir). This attitude is choaen from the several alternatives available to provide crew visibility of the ground track and Lunar Landing site through the docking wlndows while minimizing RCS propell- ant usage for attitude control. Rolling the SM Z-axis away from the local vertical (e,g. to allgn Sector I center line with the ground track to simplify experhent mounting) introduces cross-coupling into RCS thruster firings causing more complfcated crew maneuvers and less efficient propellant usage. This S/C orientation and visibility footprint is shown in Figure V. 3-2. As indicated the SM Sector I center- line is aligned 38 degrees off the ground track, and experiments located in Sector I must be mounted to accommodate this offset position in order to point the -% axis toward local vertical. Crew visibility of' the nadir is not possible unless the: C&SM is pitched downward, While this pitched attitude is not recommended for orbital flight, it reniains as a crew option anytime viewing the nadir is necessary. In the "l;treamlined" attitude the near edge of the visibility footprint is 60 NM dowrrange. Pitching maneuvers to bring this visibility footprint nearer to the nadir would interrupt ZOS experhent Lunar Surface ( loca: vertical) pointing and should be performed only as required. The nose down S/C at-titude is not considered acceptable because LOS experiment mounting in SM Sector I would be seriously complicated. The in-between ~ttitude,pitched 45 degrees toward the Lunar Surface, provides the optimum crew visibility, but also causes cross-coupling of the RCS thrusters and spacecraft interference with LOS experiment field of view in Sector I. In Figure V.3-2, the visibility footprint is shown for the right couch docking windm. In one-man lunar orbit, a crewman in either the left or right couch could control the S/C with little change i.n crew visibility. If the FDA1 l light Director Attitude ~ndicator)is required by the crewman, the left seat would be used. MARTIM MARIE7TA CTORPORAI#OW DENVER DIVISION HORIZON % 335 ?TM I i I i i i I HT DOCKING WINDOW OOST 6 ENTRY EPE POSITION I 0 SEC T& T& mc, Figure V.3-2 Crew Visibility MAR'IIN MARIE-A CORPORATfON DENVER DIV'S!r?\ PR34-113 Page V-34 Tn :~:tlkini: this anczlysis, the Yeber couch modification in the CM was assumed to have minimum impact on the eyv positions f'ornlerly specified i'or boost and entry, and docking, To maintain the Iacal vertical, "s.treamlined", orientation within the + lo attitude constrainto required by the IXlS experiments, the spacecraft must be manually controlled using the ORDEAL system with attitude reference from the Stabilization Control System (SCS). Automatic stabilization utilizing IMU reference is not feasible due to the power-down requirement for the lMIJ in orbital flight. Hi~h drift rates for the SCS attitude reference (approximately ZO/hr) require that .the Gyro Display Coupler (GDC) be re-aligned each orbit, This is most simply performed using the Crew 0:tical Align- ment flight (coAS) and the lunar horizon. Manual S/C control is not acceptable for long periods, as required by certain IQS experi- ments, Addition of an automatic pitch programming capability is required for local vertical control. V.3.3 EVA Data Retrieval - As presently conceived, MS experiment payloads mounted in the C&SM require EVA data. retrieval prior to earth return. EVA requires all three crewmen to be aboard the CM, limiting EVA opportunfties to lunar orbit after the surface mi.ssion while the CM and LM are joined or to the trans-earth coast period after the Dl has been jettisoned. Crew safety dictates that no EVA be performed from the CM during lunar orbit except in emergency situations. Uncertainty of equipment functioning after depresauriza- tion/repressuriza-tion and crew fatigue factors after lunar surface stay warrant no other choice. If EVA were performed from the LM prior to LM-jettison, an additional day in lunar orbit and a much more complicated external surface traverse to SM Sector I would be required for data retrieval. In view of these considerations, EVA for LOS data retrieval during lunar orbit is not recommended. No plan to retain the LM during TEI is considered due to limited SH propellant for the earth return maneuver. The EMU/PISS already carried for the lunar surface mission is recom- mended for the MS data retrieval EVA to minimize overall mission weight and volume stowage requirements. Even though long, high pres- sure oxygen and water umbilicals offer certain advantages, the required CM mcdifications toaccommodatethe umbilicals and water cooling, and the additional boost stowage of the umbilicals and Pressure Control Units (PCU) cause the concept to be discarded. PLSS refilled from the IM ECS are required for both EVA crewman so that the backup crewman may egress the CM, if necessary, to assist the primary EVA crewman. 6WAWT#FW BBQAWIE77"A UCDRP88PA~I9N DENVER DIVISION PR34-113 Page V-35 Figure V,3-3 shovs the general configuration of EVA required. The CM wou'ld remain depressurized throughout the EVA. The EVA crewman would proceed over the SM surface to Sector I, backed up by a 2nd standby crewman in the CM crew hatch. The 3rd crewman would monitor the CM systems. Access to Sector I from the CM would be provided by a deployed hand.. rail mounted on the SM with a short telescoping section to reach to the CM main crew hatch. Fixed handholds at the access openings and fixed handholds on transfer packages are recommended rather than use of portable handholds. Since the film cassettes and solar wind foil requiring EVA retrieval are in the 5 to 65 lb. mass range, they should not require more complicated transfer aids. In the study of EVA concepts for the Apollo Experiment Pallet (PEP), Martin Marietta evaluated, through 6 degree of freedom simulation, various concepts for retrieving data cassettes from SM Sector I. Results of these studies showed the feasibility of EVA maneuvering from the CM main crew hatch to Sector I and return via a deployed handrail and handholds. Evaluation of the transfer of slmilar size data cassettes resulted in selection of these deployed handrail and handholds. More complex transfer aids are considered unnecessary for the EVA operations required. Since LOS experiment data retrieval is very similar to that evaluated in previous AEP simulation, results are applicable to the planned LQS EVA operations. Subsequent to work in mP, the results of Gemini EVA'S on flights 9 through 12 provided substantial neutral bouyancy and actual space flight data on the utility and design of handrails, work site tethers and restraint devices. In the mainline Apollo and AAP programs, continued neutral buoyancy and KC-135 simulation hao refined the criteria for design and effective use of handrails, restraints and tethers during surface translation and data retrieval. Detailed analysis of the specific handrail, tether, and other EVA equipnent devices would follow the preliminary vehicle/experiment configuration. Figure v.3-4 shows a typical timeline for EVA during trans-earth coast. Although EVA data retrieval during trans-earth coast is feasible, the value of the data to be retrieved seems marginal interms of possible jeopardy to the completion of an alreridy successful lunar landing miss ion. It is strongly recommended that alternative methods not requiring EVA retrieval be considered. Film chutes may be designed to return exposed film to a cassette attached to the CM or even MARrRlU MARIETTA CORPORArlON DENVER DIVISION STANDBY CREWMAN -- - Figure V.3-3 C&SM-S&~O~ I Configuration During EVA MABZTIM MWRIEWA (CQRPCVAIANON DENVER OlVlSlON CDR CMP LMP CDR CMP LMP (PH - PERGONAL HYGIENE) MISSION POSTURG: C&SF IN TRANS-EARTH COAST. THIS T/L STARTS 'r7ITR THE FTET SLEP ?i%IOD AFER THE FIRST MID-COURSE CORRECTION AT SPHERE OF INFLUZECE BO'JmkSY Figure V.3-4 Typical Timeline EVA MAR7IN MARIE-A CORPORATJON DENVER nlVISION PR34-113 Page V-38 into the CM itself, Alternately, the weight penalty should be investigated Cor n separate reentry film pod that could be injected into earth entry from the SM prior to SM jettison upon earth return. This concept would take tu!vantage of the techniques for film return presently in uae from earth orbiting sstt.llites uhcre t'ilm pads are snagged by aircraft after reentry, The! solnr wind 1'011 could be rn~nuallydeployed and recovered on the lunar surf'nce during LM surface operations obviatin~any requirement l'or F:VA retrieval, V. 3.11 Clr" BoostlRcentry .;towage Requirements - F:xpcrirnent and rnissio~lpeculi~r crew equiwent required to be etowcd in the CM is shown in Table V. 3-3 CM Stowage Requirementts. These stowage requirements are based on the recommended concept for EVA data retrieval. MARTIN EWACCll57F' CQRPOlcaATIOAI DENVER DIVISION Table V.3-3 CM Stowage Requirement Equipment No. Est. Gt. Size (in.) V~lume most Entry 1. Portable Life Support 2 Charged-86.1 26 x 17.8 x 10.5 4,850 in3 X System (PLSS) Unchgd -66.0 2. Safety tether 2 10.0 1.6 x 0.75 x 30 ft 450 in3 X X 3. Metric camera film I 21.0 7 x 8 dia 56 in3 X cassette I 4. Stellar camera film . 2 6.0 4 x 7 dia 28 in3 X cassette 5, Panoramic camera 1 65.0 6 x 20 dia 120 in3 X 6. Solar wind foil I. 0.25 36 x 1.5 dh 60 in3 X 7. Film canister 1 5.0 9x5~3 135 in3 X (radar imager) * *-Experiment may not be on (=M/SM c + I 4 I I-' WW \O MARIIN MARIEITA CORPORA7IbW DENVER DIVISION PR34-113 Page VI- 1 VI S JBSYSTEM DESIGN KEQUIREMXNTS VI.1 MTA AND COMMUNICATION SYS'!XMS The data syrtam capability of four possible carriers (CMM, Lunar Orbiter, ATMP, and Pioneer) was compared to the experiment raquire- ments derived from res pons ible scientis ts and indus try sources. The a~nlyalsbelow is based on a representativa sat of groupings of experiments presented to North American Rockwell in August, 19i8, For tliesc groupings, the ger~rraltype and level of modification requlred to encli of the four carriers was determined. Tllr intPgrat ion of the cxperimenta into a carrier in almost every cauc rexu l ted in a modi f Fcntiun, (ranging from a 40% redesign of tl~eLunar Orbiter system to a complete redesign of the AIMP) requiring addition or replacement of some portion of the data system. A summary of these modifications is presented in Table VI.1-1, based on the experiments of Table VL.1-2, This cable fur- ther verifies that the change8 required are dependent upon the misoion experiment grouping. The mejor components affected are the data storage system, the modulation system, the PCM multi- plexing system and the wiring and distribution of commands. The prime areas of concern are the multiplexing of digital data and the storage of high bit rates data (up to 60 kbps) for up to two hours. Further definition of hardware required to accomplish these tasks should be s~udiedin future programs. v1.1-1 Stable Satellite Analysis - Lunar Orbiter - The basic Lunar Orbiter satellite contains a 'talemetry (T/M) system capable of trans- mitting 50 bits per sacond PC24 and 250 kilo hertz video data plus a 38 KHz pilot tone over an S-band carrier frequency. The satellite also has a receiver which is used for receipt of commands and coherent ranging when used in conjunction wich the T/M transmitter. The 50 bps data are used for monitoring analog and bilevel measurement of satel- lite performance, command verification, flight programmer data and micrometeoroid detectors. Removal of the micrometeoroid detectors, the photographic system, and redundant measurements will provide ca- pacity for approximately 300 bits of data at very low sample rates. Removal of the photographic system also provided the RF link for wide- band inputs. The Lunar Orbiter system is incapable of handling serial digital bits, and with its low bit rate, the capability of handling experiment data without modification to the data system is limited. ------. ------To establish the level and scope of modifications required to the Lunar Orbiter several experiment groupings wers reviewed and their impact on the system a.scertained. Figure VI. 1-1 is a system which results from the four sample groupings of Table V1.1-2. The experi- ments considered and theit. data rates are enumerated on the figure. This system will accommodate all the experhent data. The basic - - -.. MARrfN MAIFFILh-A 001gPQlPAIIQN DENVER DIVISION TB3LE VI .1- 1 SUMXARY OF DATA SYSTEM MODIFICATIONS REQUIRED TO IMPLEMENT EXPERIMENT GROUPINGS PRESENTED STABLE SATELLITE SP NNING SAT :LLITE cxfm PIONEER Am GROUP I G30U-P I1 PROPOSED PROPOSED WEIGHT MAJOR COMPONEIqTS MClD EOD POmm MO Antennas & Ant Ckts No Ckg. 14.95 New Kew Traveling Wave Tube Am No Cng. New New NO Chg. ------I Transmitter No Chg. New New Modulator New 4. O* New New No Cng. Hew in SM Command Receiver No Chg. Minor New No Chg. Nc Chg. I Mod Command Decoder No Chg. Minor New No Chg. No Chg. I Mod Command Decoder Driver No Chg. Minor New No Chg. No Chg. I Mc3d Transpondei Circuits No Chg. 12.18 Maj or New No Chg. Bc Chg. Mod Data Storage System 20. o* New New Hew in 2-New SM INew PM Encoder/Digital Pioneer 18.46 New New New in 10.00* New in I 10.W Multiplexer or Equiv. SM Signal Conditioning New New 3ew in 1.5" l New 1 lS5* SM Timing (Time of Day) New D New New No Chg. ------Wiring & Controls Ckts Major Mod 20* Major New New in 20* New in 20f;' Mod SM SM Command Distr. System Major Mod Major New New in 3" Mod SM *Estimated Included in Tran: 1 Command Decoder NOTE: See Table VI.l-2 for experiments included above. MARTIN MARIE-A CORPORATION DENVE.? DIVISION TABLE VI .1-2 SAttrlPLE EXPERIMENT GROTJPING - INPUT TO NORTH AMERICAN ROCKFELL AUGUST 1968 Command and Service Module S,lbsar 2llite Group I 1 Group I1 Stable I Spinrling - r I Metric & Stellar Camera IR Mapping Radiometer Geochemistry Expts. ' DC 3agne:sncter I Panoramic Camera IR Spectrometer I Mass Spectrometer I' .4C mgnetcmeter Photoelectric PhotometerI Microwave Radiometer RadarILaser Altimeter Plasma Probe Geochemistry Metric & Stellar Camera RF Transponder/Laser Electric Field Experiments Re troref lec tor Laser Altimeter RF Reflectance Electromagnetic Low Energy Particles Measurements RF Transponder I Bi-Static Radar Radio Noise Survey Solar Energetic Particles Solar Wind Fcil Radar Altimeter i Gravity Gradiometer Meteoroids RF Transponder Meteoroids i RF TransponderILaser Retroref lector I Radar Imager I I UV Spectrometer MARTIN MARIE-A CORPORATION DENVER DIVISION I------1 r------Comand I 4 Decoder I-; ' Low Gain L-, ,, , ,, J -- - I Alt 1 Trrnoponder It Cravi ty I 1 Grnziometer Yaltiplener, I i Parawtars L------J Hodu la t ion Radar Altinctcr bss Spcctrorccter h'oise Survey Elcctr-;netic Ikasurc~cnt ihrdware Figure VI.l-1 Stable Satellite - Lunar Orbiter - Data Sysrec: MARTIN MARIE-A CORPORATION DENVER DIVISION PM4-113 Page VI-5 data, command, ra~lgingand transponding systems are left intact. The analog data from the experiments would be encoded to six bit resolution and sampled at one sample per sixteen or sixty-four frames. The encoder would provide synchronization and clock signals to the experiments to provide the correct cycling of data inco the PCM main frame. The data storage system would be capable of recording data for two hours with a playback speed of 10:l. Three data tracks would be required and the recording speed would be 1-7/8 ips or less since the data rates are very low. The modulator and modulation select logic system would contain the circuitry required to place the incoming data on separate subcarriers, and provide mixing and signal conditioning of the proper signal for input to the transmitter. The configuration changes required to implement this grouping are noted in Table VI. 1-1. VI. 1-2 Spinning Satellite Analysis - Pioneer - The basic Pioneer satellite data system has the capability of transmitting real time and stored data at bit rates of 512, 256, 64, 16 and 8 bits per second. It has four PCM main frame formats for ana- log and digital data plus two subframes for ,lousekeeping data which are multiplexed into the selected main frames. The system output is bi-phase modulated onto a 2048 hertz (HZ) subcarrier which then modulates the S-band carrier of the transmitter. The satellite also contains a command and transponder system for controlling operation of onboard systems and doppler tracking of the vehicle (Ranging is not available). Each main frame contains 26 words that are available for experiment data. each word is seven bits long (six data plus one parity bit). Using the maximum bit rate of 512 bps provides 355 bps for experiment use. The Pioneer data system as described does not have the capacity to transmit realtime and stored data simultaneously, nor does it contain a ranging system. Adaptation of this sytem to a particular experiment grouping presents the same type of problems as noted for the Lunar Orbiter. Figure VI.l-2 presents a data systtm configured to manage the data from the experiments presented in Table VI.1-2. The total bit rate to be handled (3,150 bps) is much greater then the 355 bps available; there- fore the system requires the addition of new hardware. Table VI.1-1 summarizes the changes required to satisfy the experiments. The basic Pioneer System transmitted 2048 Hz signals over its S-band link. Therefore the bandwidth of the transmitter, TWT's and antennas, if optimized to this signal, will not contain sufficient bandwidth to handle the playback data. Modification MA#?T'#PEPIWAWbEg"lfrEP CO~POIWATIB~ DENVER DIVISION PR34-113 Page n-6 or replacement of these components would be required. The tape storage system would require 3-314 ips recording speed, with a 10:l playback ratio. The PCM encoder would have data capacity as noted in Figure VI.1-2 and the modulation system would perform the same function noted for the Lunar Orbiter modulator, A major modification of tlie tran~pondercircuits would be required to provide capability for ranging, to support tracking and experi- ment Eunctions. The modifications required on the Pioneer are to a greater level than those of the Lunar Orbiter but the basic approach to handling the data is the same. VI. 1-3 Spinning Satellite Analysis - Anchored Interplaneta -Monitoring Platform CAIMP) - The basic AIMP satellite telemeter utilizes pulse frequency modu1ati.cn and was designed to provide telemetered data at distances up to 600,000 KM. This system utilizes VHF frequencies (136 & 142 MHz) for ranging, commands, and transmission of telemetered data. The system is capable of encoding analog data and multiplexing digital data. Both data are then pulsed frequency modulated onto the carrier utilizing frequency bursts which have an equivalent binary value of four bits. The TIM output is composed of frequency bursts that are equiva- lent to a sequence of 1320 digital data bits, 974 analog data bits and 128 snychronization bits which require 81.92 seconds to transmit, resulting in an average bit rate of 29.6 bits per second. The data experiment requirements for the AMP and the Pioneer are the same (see Table VI. 1-2). The AIMP system is totally inadequate to handle the dsta; therefore, a new system is required. Since the experiment requirtimerlts are the same as for the Pioneer, the system shown in Figure VI.1-2 can represent the AIMP system with all new hardware. VI. 1-4 Command and Service Modu1.e (C&SM) Analysis - The C&SM data and ranging system has as its prime function the support of the Apollo ~ission. Excess data capacity and time sharing are therefore the only areas that. may be considered for support of the experiments. The data requirements for the sample experiment groupings for the C&SM (Table VI.l-2) are indicated in the data and system shown on Figures VI.1-3 and VI.l-4 respectively. The major obstacles to accommodation of all the experiment data directly into the Apollo f?rmat for either of the two C&SM experiment groups of Table VI.l-2 are: MAmrIN MAmIErrA OQRPORARON DENVER DIVISION COPrnQS TO ALL SYSTEXS DIGITAL DATA HG&TI?LEXER PW) DULATION Figure VI.l-2 Pio~ieer- Spinning Satellite Data System IS.A RTfN MA RIE-A CORPORA78ON DEN JER DIVISION Laser Video to PSotocleetrfc Signal 1- Photometer Cond . . * 6KHZFM Link *Group I Exparimento per Table VI.1-2 Figure VI.l-3 C&SM - Group I Experiments* - Data System MARrIW MARIEWA CORPORAZION DENVER DIVISION Radar Altbeter &l tiplaxed r Sflcctiince Video to Encoder Xztric -\a, WiC BPS 6 HHZ Link Czwra a 1@33-1/3 Bist~tic 1 @ 66-1/3 Redat Radar 30 at 1/8 SPS Imger Modulator 200 KIPS & Recorder #2 Modulation Select Logic W Spectroaeter Pecoxder 91 Iu Yapping 10 KDPS bdioneter I IB 2.85 KDPS Spectrozieter 4 a *Group XI Exporfiz=nts per Table VI.1-2 Figure ~1.1-4 C&SM - Group 11 Experiments* - Data System MARTIN MARIE-AI CORPORAtIOlV OENVER OlViSlOU PR34-113 Page VI-10 (a) Limitad digital data capacity, (b) Insufficient data storage (digital or FM),insuffi- cient storage time (duration) and insufficient playback to record ratio (1:l) of the recorder (in the high bit rate mode); (c) Time sharing of the encoding system and the FM links whicl~coulcl impair the Apollo data accluisition or the experiment data acquisition or both, due to the high data rates required by the experiments. The adaptation of the groupings to the C&SM system could be accomplished in several ways. Figures VI.1-3 and VI.l-4 present simplified methods of accommodating the data. Figure VI.1-3 represents a system that would utilize the biomedical subcarriers and the 6 MHz wideband FM carrier identified by North American Rockwell for transmission of realtime and playback data. Updata link, manual commands, and ranging would be accomplished by the existing C&SM onboard systems. The tape recorder used for this grouping would be operated at 1-7/8 ips and playback at a 10:l ratio for a two hour recording time. Therefore, the use of the 125 KHz subcarrier would be limited to about 12 minutes per orbit. This data system has a minimum impact upon the C&SM system. The experiment daca system shown on Figure VI.l-4 is more complex and places greater requirements upon the h~sicC&SM system. The groupirg contains experiments with high data rates and require- ments for continuous operation; therefore, realtime and playback capability is required to support them. The system shown would require a PCM Encoder/multiplexer, two recorders and a modulator located in the SM. The three subcarriers of the C&SM (85 KHz, 125 KtEz, and 165 KHz) and the six MHz wideband FM channel would also be required to support this mission. If the six MHz channel is not available, a separate S-band transmitter would be required. Total transmission time of the wideband FM would be limited to approximately twelve minutes per orbit. VI. 1-5 Ground Checkout and Launch Requirements (a) Subsatellite Data Systems - The 1oc.ztion of the subsatellite dTring prelaunch checkout will be in the Sector I of the SM. In this position, al.1 the deployable antennas will be stowed and open loop verification will not be possible. Therefore, all checkout after satellite installation will be accomplished closed-loop or by using test outputs that will be designed into the system. The interface after installation will be manual-disconnect-umbilicals with terminations inside MARrfN MARf/1TA 00RPORArfON DENVER DIVISION PR34-113 Page VI-11 Sector I therefore not accessible after cover installation. Critical parameters that require monitoring after Sector I is sealed and during the interval of time prior to deployment could utilize spare channels in the C&SM data system. Th~refore,they could be monitored as part of the normal C&SM data flow throughout prelaunch and pre-deployment pnases of flight. The data system checkout requirements under the constraints noted previously would be limited to verifying the correct operation of prime components. Parameters such as output power and voltage levcl~and the verification of subcarrier and carrier frequencie~l, the bit rates, command and command responses and othar prime param- eters required wil.1 be monitored to establish the integrity of the system. Due to the limited number of measurements and the non-standard bit rates of each of the experiments, complex auto- matic checkout equipment will probably not be used to verify systems operation. The ground checkout equipment will utilize standard ground station equipment with the possible exception of the PCM decommutator. The deconunutator will require the flexi- bility of handling non-standard bit rates. (b) C&SM Data System - The experiments studied thus far for flight in the C&SM utilize the RF loops and command system provided by the C&SM. Therefore, the checkout of any portion of the data system located in Sector I can be performed using the normal C&SM checkout methods. Self calibrate or simulators may be required to support checkout but verification of the parameters defining experiment integrity will be required in any case. Since the recommended systems do utilize the six MHz wideband FM channel, automatic checkout equipment can be by-passed. This allows use of ground station facilities tailored to meet the Sector I data re- duction requirements with reduced complexity in equipment and checkout methods. '71.1-6 Peculiar Subsystem Support Requirements - It is antici- pated that any data subsystem used in the C&SM or subsatellites will be designed to meet the expected environments without special support. Protection will be required for extreme environments such as micrometeorites high radiation sources, temperatures below O'F and above 1600 F.' MARrlN IWARRE'17A UORPORAllON DENVER DIVISION PR34-113 Page VI-12 'Illc power 8ystc.m requirements necessary to support the selcctad experiment groupings on each of four carriers, the C&SM, tl~c? Lunar Orbiter (small stable satellite), the Pioneer mall spinning satellite) and the Anchored Interplaqe~aryMonitoring Platform (small spinning satellite) were analyzed. The latter three carriers are subsatellites deployed from the C&SM Sector I. This section presents the existing vehicle Fower config- uration, reflects the power required to support the sample groupings, and suggests the revisions to the existin& vehicle power systems to provide the required power. Based on the experiment requiraments and eroupings of Section 111, the required power f~rexperiments exceeded the existin capacity of all four carriers. The capacity of the three st,- satellites with power output approximately doubled is summarized as follows: Lunar Orbiter Solar array - double size - 766 watts total output Battery - add 2 for a total of 3 252 watts Pioneer Solar array - double size - total 170 watts output Battery - add Lunar Orbiter type 84 watts SolarArray-doublesize- total 130watts output Battery - add Lunar Orbiter Type 84 watts However, even the increases shown above are inadequate to meet the grouped experiment requirements of Table 111.1-2. The power system revisions necessary to support the experiment groupings in the CGSM Sector I were analyzed for a three-day mission. Since the CSSM does not have a solar array, all power required must be provided by the addition of batteries to the system. The revisions to the power systems suggested herein reflect the rnaximun~ feasible increases based on using multiples of the exist- ing solar panels and batteries. However, further studies may reveal that it would be possible to provide the required power by complete redesign of the solar array and use of a battery with an increased output capacity. It should also be noted that such MARrDN M'LeRDR-A OORPORArDON DENVER DIVISION perturbations as changing the orbit from circular to alLipticaJ. could improve the day to night: ratio and reduce tlla battery power required, thus reducing the battery charging power rt~quiredof the solar array. Experiment operating timeline rcvioions can alsu be used to reduce the total power required as opposed to the continuous operation now requirrd of most experiments. The experiment requirement that the experimant axis remain aligned to the lunar local vertical for the small stable satellite experiment groups, together with the power system requirement that the solar array be aligned to the sun, necessitates a gimbal system for either the solar array, the experiments or both. However, since the power output varies as the cosine of the illumination angle a ,sixty degree change will still provide one half the total power. The proposed battery dissipation during experiment opera- tion is only 25% and the battery should provide full power for 6000 cycles or 500 days, VI.2-1 Subsatellite Power - Each of the three subsatellite vehicles, the Lunar Orbiter, the Pioneer and the AW, has a solar array/battery power system. Only the Lunar Orbiter, however, has a system capable of recharging the battery to provide power for vehicle operation during sun occult for a circular lunar orbit. The use of these vehicles as subsatellites for future Apollo missions will require a system to provide power during both orbital day and orbital night operation, The power subsystem requirements are based on a sixty nautical mile, circular lunar orbit which provides a near 1:l day to night ratio. Since all power dissipated by the battery during orbital night must be replaced by the solar array via tke charge controller during orbital day and since the present timelines call for continu- ous operation of most experiments, the day and ntght experiment power requirements are almost identical; therefore, the solar array output capability must be approximately double the battery output. Based on a thirty day prime mission and a two hour orbital period the battery would require 360 cycles. VI.2-2 Small Stable Satellite Power - The present Lunar Orbiter power system is a solar array/battery system. The solar array is a four flat panel n-p solar cell power source composed of 10816 2x2 cm cells connected in a series-parallel arrangement to provide a voltage of 26.6 to 31,O WC with a maximum power of 383 watts. The total output power of the solar array may be increased by increasing the number of solar panels composing the solar array system. The maximum number cf panels feasible con- sidering the structural and weight constraints imposed on the vehicle is probably about eight. The existing system and a system with double the power are shown below: MARTUN ICBAffJdrTA bORPOffATION DENVER DIVISION PR.34-113 Page VI-14 Eight Panel Exia t ing Des ign Des ign Power 383 Watts Power 766 Watts Voltage 28.- Voltage 28 VDC Weight 71 1bs Weight 142 Ibs The Lunar Orbiter battery is a 30-pound 12-ampere hour nickel- cadmium power sotlrce to provi.de vehicle power for sun occult operation. The major design cons ideration for establishing the number of batteries required however, is the fact that the orbital 1:l daylnight ratio the current dissipated during the sun occult operation must be replaced by the solar array via the charge controller during the illuminated portion of the flight. Therefore, the maximum cha: ge rate of three amps at 28 VDC limits the battery output to 84 watts. Based on the weight and space constraints for the Lunar Orbiter, the maxi- mum number of batteries to be added would be two for a total of three batteries. The thiee battery system will provide a total continuous power during sun occult of 252 watts at 28 VDC. The weight of the three battery system Is 90 pounds. The total vehicle power to support the proposed missions includes subsystem power as well as the power to operate the experiments. The subsystem power requirements are given by Table VI.2-1. The required experiment power and total vehicle power for the small stable satellite experiment groupings follow in Table VT.2-2. The total maximum power available from the eight panel solar array and the three battery system is not sufficient to support the Group A and C experiments as noted in Table VI.2-2. There- fore, a redesign of the solar array/battery system would be required to further 5.ncreasc: the output capacity to support such a mission. The Group B experiment totals, however, reveal that the eight panel, three battery system would be capable of support- ing the mission as the nightti~uediscrepancy is sufficiently small and should he easily solvable. MARTIN MARIE7TA CORPORAYION DENVER DIVISION PR34-113 Page VI-15 TABLE VI. 2-1 STABLE SATELLITE SUBSYSTEM POWER REQUIREMENTS ORBITAL DAYTIME ORBITAL NIGHTTIME WATTS WATTS Communi.cations Subsystem - Total (142.5) Timing (Time of Day) Iransponder System 17.0 Command Receiver Transmit tcr TWT 70.0 Data Storage System 35 .O Signal Conditioning 0.5 PCM Encoder/Multiplexer (2) 14.0 Modulator 2.0 Command Decoder 4.0 Command Decoder Driver Power Subsystem - Total ( 16.0) Battery Charge (Varies with total power - See groupings) Shunt Regulator 2.0 2 .O Charge Cont::o 1ler 14.0 Attitude Control Subsystem - Total (125.0) Thermal Control Subsystem - Total ( 43.0) Total Subsys tem Power (without active thermal control) 283.5 (with active thermal con- trol) 326.5 Note: Thermal control is based on the stable satellite experiment Group C on Table VI,.2-2. MAWTIM MARIETZA CORPORATION DENVER DIVISION TABLE VI.2-2 STABLE SATELLITE POWER SUMMARY GROUP A GROUP B I GROUP C Power Power Power Experiment Watts Experiment Watts Experiment Watts Geochemistry Ceochemistry 75 IR Mapping Radio- Radar Altimeter Mass Spectrometer 12 rcetefi 150 Gravity Gradiometer Meteoroids 5 IR Spectrometer 15 Electrcmagnetic Meas. Hd Radiometer 40 Mass Spectrometer RF Reflectance 70 Meteoroids Bi-Static Radars 2 Noise Survey Radar Altimeter 70 Electromagnet Meas. 50 - - Noise Survey 1-2 Total 1 222 watts 92 watts 402 watts Daytime Operation: Experiment Total 222.0 watts 92.0 watts 'M252.0 watts i Total Subsystem Power 283.5 watts 283.5 watts 326.5 watts Battery Chrg. Power 500.0 watts 375.0 watts 800.0 watts Total Daytime Power 1005.5 watts 750.5 watts 1378.5 watts Max Power Available 766.0 watts 766.0 watts 766.0 watts from 8 panel solar array I - Power margin from 8 1 Panel Solar Array - 239.5 watts + 15.5 watts - 612.5 watts Nighttime Operation: Experiment Total 222.0 watts 92.0 watts 402.0 watts Total Subsystem Power 173.5 watts 173.5 watts 216.5 watts % w Total Nighttime Power 265.5 watts 618.5 watts & i2 395.5 watts 1- 0 C I Power Available from 3 z 3 batteries 252.0 watts 252.0 watts 252.0 watts I W - F Power Msrgin from 3 3 batteries - 143.5 watts 1 - 13.5 watts - 366.5 watts * Operates every other orbit, nigh time * Operates every orbit nearside ** Daytime power is total power min s the IR mapping radiometer PR34-113 Page VI-17 The three groups discussed here provide a total power range sufficiently broad to allow any grouping from the Section 111 roadmaps to be analyzed. Assuming the subsyrltem power is con- stant, the total experiment power for any such grouping may be compared with the total experiment power of the sample groupings given in Table VI.2-2 and the feasibility determined. VI.2-3 Small Spinning Satellite Power - Two existing satellites, the Picneer and the Anchored Interplanetary Monitoring Platform (AIW), were considered for use as spinning subsatellites. Both liave solar arraylbattery power system and the main difference is in structural design configuration. The Pioneer solar array system consists of approximately 10,000 n-p type solar cells 2x2 cm spread over the cylindrical spacecraft surface. The cells are connected in a series-parallel combination to provI.de 24 to 33 VDC with a maximum power of 85 watts. From a 2ower point of view, spinning satellites are normally used when the total power required from the solar array is 100 watts or less. When the power required is in excess of lOG the small stable sate'llite with the sun oriented solar array is selected. For the purpose of this study, doubling the size of the solar array was assumed to be feasible, thus doubling the power output capacity. The existing system and the double capacit) configura- tion are shown below. Doub 1e Exis ting Output Design Des ign Power 85Watts Power 170 Watts Voltage 28 VDC Voltage 28 VDC Weight: 17 lbs Weight 34 lbs MARFPPIJW MARIETTA CORPORATION DENVER DIVISION PR34-113 Page VI-18 The AIMP solar array is composed of four flat panels with 7680 2 x 2 cm solar cells. The cells are connected in a series - parallel combination to provide 19.8 VDC with a maximum power of 66 watts. The maximum number of solar panels feasible is probably about eight. The eight panel configuration will double the solar array output capacity. The number of cells in series will be increased to provide the 28 VDC required by the experi- ments. The existing system and eight panel configuration to double the power output are shown below. gting Design Eight Panel Design PI Power 66 watts Power 130watts Voltage 19.8 VDC Voltage 28 VDC Weight 21 Ibs Weight 42 Ibs Neither the Pioneer two ampere-hour silver-zinc battery nor the AIMP eleven ampere-hour silver cadmium battery have the capacity to provide the continuous power required for sun occult opera- tion. To support the proposed groupings one Lunar Orbiter type thirty pound, twelve ampere hour, nickel cadmium battery replaces the existing battery for either vehicle. The maximum charge rate of 3 amps at 28 VDC limits the total dissipation to 84 watts per orbit. The total vehicle power to support the proposed missions includes sutsystem power as well as the power required to operate the experiments, Table VI.2-3 gives the subsystem power requirements. The required total vehicle/experiment power for the small spinning satellite is given in Table VI.2-4. The total maximum power available from the solar arrays as shown in Table VI.2-4 is not sufficient to support either Group A or Group B. Therefore, a redesign of the solar array system would be required to further increase the total output capacity. The Lunar Orbiter type battery added to either vehicle would provide sufficient power for sun occult operation as shown in Table VI.2-4, MARTIN MARPfE-Al COI;TPLPWA71ON DENVER DIVISION PR34- 113 Page VI-19 TABLE VI. 2-3 SPINNING SATELLITE SUBSYSTEM POWER REQUIREMENTS ORBITAL ORBITAL DAYTIMIS NXGllTTIME WATTS WATTS TWT Transmitter Modulator Command Receiver Cornmand Decoder Command Decoder Driver Data Storage System PCM EncoderIDigital Multiplexer Signal Conditioning Time (Time of Day) Power ( 8.0 ) ( 4*5 Shunt Regulator 1.O 1.0 Charge Controller 7 .O 2,5 Battery Charge (Varies with total power - see groupings) Attitude Control ( 5.0 ) ( 5-0 ) Therrnal Control ( 5.0 ( 5.0 ) Total Subsystem Power 93.75 39.75 IWARWN MARRkrrA C0,QPORAIRON DENVER DIVISION PR34-113 Page VI-20 TABLE VI.2-4 SPLNNXNG SAITELLITE POWER SUMMARY GROUP A EXPERIMENTS GKOUl? B EXPERIMXNTS POIJER POIJER EXPERIMENT WATTS EXPERIMENT WATTS k2agnetometer - DC 1.5 Magnetometer - DC 1.5 Magnetometer - AC 2.0 Magnetometer - AC 2.0 Plasma Probe 3.5 Plasma Probe 3.5 Electric Field 3.0 Electric Field 3 .O Low Energy Particles 5 .O Low Energy Particles 5 .O Solar Energetic Particle 5.0 Solar Energetic Particles 5.0 Meteoroids -5 .O - Exper iment Total 25.0 watts 20.0 watts Daytime Operation: Experiment totals 25 .OO watts 20.00 watts Subsystem Totals 93.75 watts 93.75 watts Battery Charging 100.00 watts 100.00 watts Total Daytime Power 218.75 watts 213.75 watts Max, Power Avail from Double Capacity Picneer170.00 watts 170.00 watts Max. Power Avail from 8 Panel ALMP 130.00 watts 130.00 watts Pioneer Power Margin - 48.75 watts - 43.75 watts ATW Power Margin - 88.75 watts - 88.75 watts Nighttime Operat ion: Experiment totals 25.0 watts 20.00 watts Subsystem totals 39.75 watts 39.75 watts Total Nighttime Power 64.75 watts 59.75 watts Power Avail from Lunar Orbitor Type Battery on Both AIMP & Pioneer 84.00 watts 84.00 watts Power Margin Pioneer & AIMP 4-19.25 watts 3-24.25 watts MAR7fM MARfh-A CORPORATION DENVER DIVISION PR.34-113 Page VI-21 The feasibility of using the small spinning satellite to support other experiment groupings rasulting.from the roadblocks of Sec- tion 111 can be analyzed by comparing the total experiment power requirad for such a grouping with the experiment power of Group A and B of Tabla VX.2-4, assuming that the subsystem power is con- e tant. 4. CGSM Power - Experiments hardmounted in Sector I of the ervice M- arls supplied 10 KWH power from the C&SM power ystem. All additional power required will come from batteries added for that purpose. The total power requirements are based on a 60 nautical mile, circular lunar orbit and reflect a three day miss ion. Table VI.2-5 reflects the power required by subsystems added to the C&SM to support the proposed experiments. They interface with the appropriate C&SM sys",ms to complete the desired func- tions where applicable. Two groupings of experiments, Group t: and Group B, are identified by Table VI.2-6 and VZ.2-7. The total power for each experiment is computed on the basis of total operational time for the desig- nated missions. Ten KWH power is reserved for experiments from the basic C&SM power syetem. The batteries added provide for the addi- tional power required. VI.2-5 Power System Ground Checkout and Lzunch Requirements (a) Subsatellite Power Systems - Ground checkout of the subsatellite power system will be completed prior to the installation of the subsatellite in Sector I of the CWM. Com- plete checkout will include battery load tests, solar panel output verification and continuity checks of the electrical power and distribution system. No further testing will be re- quired after the subsatellite is installed in Sector I. The subsatellite power syetem is not operational until the sub- satellite is deployed from the CMM. (b) CGSM Experiment Power Syst$ - The powitr system required f-o support the experiments hardmounted In Sector I of the CMM will be checked out using the CUM checkout umbilical and will be part: of the normal CMM checkout sequence, MARPbW MARRE-A OQHPOWArROIY DENVER DIVISION TABLE VT.2-5 ChSM SUBSYSTEM POWER REQUIREMENTS -WATTS Communications (42 .O) Experiment Group A (See Table VI.2-6) PCM Encodcr/Multiplcxer 7 ,O Data Storage System 35 .O Communications (75 .O) Experiment Group B (See Table VI.2-7) PCM EncoderlMultiplexer 7 "0 Data Storage System 67 .O Modulator 1.9 Power ( 5.0) Altitude Control (Provided by C&SM) Thermal Concrol ( 5.0) Experiment Group A Thermal Control (43.0) Ex~erimentGroup R Total Subsystem Power 5 2 Experiment Group A Total Subsys tern Power 123 Expe r iment Group B MARIIN MAISfETTA CQRPORArfON DENVER DIVIS!ON TABLE VI. 2-6 C&SM POWER SUMMARY - GKOUP A EXPERIMENTS TOTAL EXPERIMENTS EBRk! OPERATION POWER Metric & Stellar 80 .O Every 4th Orbit .7 Camera Day1 ight Panoramic Camera 500.0 1 Hr 1st Day 1.1 1.2 Hr 2nd day 1 Hr 3rd Day Photoelectric 4.0 Every Other .1 Photometer Orbit Daylight Geochemistry 75 -0 Experiment Continuous Laser Altimeter 100.0 Continuous 7.2 Solar Wind Foil 0 Every Orbit 0 Daylight (30 Hrs required) Total 759 .O Total 14.5 KWH TOTAL EXPERIMENT SUPPORT POWER REQUIREMENTS Experiment Total 759 Watts 14.5 KWH Communications Subsystem 42 Watts 3.0 KWH Power Subsystem 5 Watts .4 KGnr Attitude Control Subsystem C&SM PROVIDED - Thermal Control Subsystem 5 Watts - .4 KWH Total Experiment Loads 811 Watts 18.3 KWH VEHICLE POWER SYSTEM REVISION Power Available from C & SM 10 KWH Add 1 Silver-Zinc Battery 3 Day Mission 12 KWH 130 LBS. EHAIPTfN MARUE-A eORPQRATUQM DENVER DIVISION PR34-X13 Page VI-24 TABLE VI. 2-7 C&SM POWER SUMMARY - GROUP B EXPERIMENTS POWER TOTAL EXPER'LMENTS WATTS OPERATION POWER IR Mapping Everyother Orbit Radiome tcr 150 .O Uarkside 5.4 IR Spectrometer 15 .O Continuous 1.1 MW Radiometer 40.0 Continuous 2.9 Metric and Every 4th Orbit Stellar Camera 80.0 Daylight 0.7 RF Reflectance 10.0 Continuous 0.7 Every Orbit Bi-Static Radar 2.0 Near Side 0.1 Radar Altimeter 70 .O Continuous 5 .O Every 4th Orbit Radar Imager 340.3 Day & Night 6.1 Every other W Imager 24.0 Orbit, Day 0.4 Total 731.0 Total 22.4 KWH TOTAL EXPERIMENT SUPPORT POWER REQUIREiiiENTS Experiments 731 Watts 22.4 KWH Communications Subsystem 75 Watts 5.4 KWH Power Subsystem 5 Watts .4 KWH Attitude Control Subsystem C&SM Provided - Thermal Control Subsystem 43 Watts --3.1 KWH 854 Watts 31.3 KWH VEHICLE POWER SYSTEM REVISION Power Available from C&SM 10 KWH Add 2 Silver-Zinc Batteries 3 Day Mission 24 KWH 260 LBS MARrtN MARIErYA CORPORATfON DENVER DIVISION VI. 1 GUIDANCE & CONTROL Tliis sect Lori describes the attitude control subi ys tem support for lunar orbital scLpnce experiments. The experimcnL~require evn luation of tliree ticparate control concepts: Continuous orientation of SM Sector I mounted experiments along the lunar local vertical. Continuous long term orientation of experiments on a sub- satellite along the lunar local vertical. Spin stabilization of a subsatellite. Subsystem study effort was limited to an investigation of the feasibility of meeting experiment requirements and it was con- cluded that provision of the necessary attitude control to meet these requirements is feasible. No unreasonable problems were identified, although not all hardware is available "off the shelft' at the present time. In all cases, requirements may be met by alternate methods. Principal identified alternatives are described, but mode studies have not been performed and no selection is iniplied. VI.3-1 Command and Service Module - Three systems for achieving local vertical orientation and attitude hold of Sector I mounted experiments are described below:- a. PGNCS and RCS ; b. Orbital Stabilization System (OSS) and RCS with Sector I mounted horizon sensors ; c. Independent auxiliary system in Sector I. A detailed accuracy study has not been performed, however, previous work has shown that a pointing accuracy on the order of 1 deg. is reasonable. The PGNCS could be expectea to do somewhat better than this dep+>ndenton the frequency of Inertial Measurement Unit alig,lments. The PGNCS would provide local vertical pointing througri appropriate computer programming to calculate the direction of iocal vertical from an inertial coordinate system plus further transforinations KO determine the orientation of Sector I relative to the local vertical. On this basis the vehicle could be maneuvered to the proper attitude and maintained there in the automatic command module computer MARTIN MARIEITA CORPORATION DE'qVEH DIVISION PR34-113 Page VI-26 attitude hold mode. This system would require no additional hardware and is obviously lowest ir~component weight. Cost would depend on the extent of computer reprogramming required and availability of memory, Crew time would be required for periodic realignment of the IMU. Power requirements are relatively high and the additional operating time of C&SM components would some- what reduce the probability of mission success. Thc OSS and RCS system wculd obtain a local vertical attitude reference from horizon sensors mounted in Sector I and aligned wit11 tile experiment sensitive axis. Signal processing, including compensation, pulse modulation and thruster command logic, is provided by the MSC developed OSS. The OSS outputs would inter- face with the RCS for control acc~?lerations. However, in this case the difference between the coordinate system which would be defined by the sensors and the coordinate system established by the RCS must be considered. The rel.ationship between the sensor reference (R) system and the body (B) system established by the thrusters is indicated in gigurt VI.3-1 where superscripts denote the coordinate frame and X = X . Since the angle B (on Figure VI.3-1) is about 32 degrees the cross coupling is severe and, in addition, the normal horizon sensor output polarity is reversed. To correct this, the system must include resolver circuitry. Also, since a horizon sensor provides pitch and roll attitude information only, a gyrocompass or equivalent technique is necessary for 3-axis attitude control. The resulting system block diagram is shown in Figure VX.3-1. This system overcomes some of the disadvantages associated with the PGNCS but with a weight and possibly a relative cost penalty. A horizon sensor for lunar application is not presently avail- able hut a suitable instrument has been under development by the Barnes Engineering Co. If a separate pcqel: supply is furnished for the components in Sector I, it would also be necessary to provide isolation between the resolver output and the OSS. Both these systems would utilize the C&SM RCS for control and therefore it is necessary to evaluate RCS propellant usage rates. In the applicable lunar orbit, the principal dynamic disturbance torque would result from gravity gradient effects. However, the following discussion demonstrates that this effect may be neglected for this case and that RCS propellant may be calculated on the basis of an undisturbed limit cycle. Taking p), 8 and 9 as roll, pitch and yaw respectively and moments of inertia as customarily defined for the C&SM, the gravity gra- dient torques will appear essentially in pitch if it is assumed that angular displacements are small and that Iyy r/Izz. MARTIN MARIE-A CORPORATION DENVER DIVISION C&SN View AFT Loo- For :art3 Figure VI.3-la Orbital Stabilization Coordinate System MARIIN MaRIEITLI COIPPORATION OEWVER OILISION Figure VI.3-lb Orbital Stabilization System and RCS Mechanization MARTIN MARIE-A CORPORAZION DENVER DIVISION PR34-113 Page VI-29 The solution to the equation is: 0(t) = 0coshdt + -@o sinhdt 4 where Q and Q0 are the atcitude deadband and the attitude rate 0 produced by a thruster firing respectively. It is possible to plot propel.lant usage, W as a function of P ' -CdQ, and this results in Figure VI.3-2 where a one jet limit Q cycle has been ass med with a minimum impulse propellant usage, of 5.48 x 10-3 lb/pulse. Values obtained directly from the wml figure woul4. be valid only for a symmetric case with 8 the same 0 after each impulse although it appears possible to obtain W £01 P non-symmetric cases by summing the AW for segments of a limit P cycle. For the study, however, it is adequate to note that the astronaut could re-initialize and eliminate particularly bad cases and inspect the figure for the applicable value of which is approximately 1.8 x 10-3 sec -1. Inserting numerical data for Qo of 0.5 deg. aad a go given by: results in: "0 = -0.28 and W Q 0.08 lb/hr -- P @o By contract, the propellant for an undisturbed limit cycle is given by: 2 2 N 3650 1% W- ('rain) P 8 MAR7IN MARIEITA CORPORATION DENVER DIVISION = 5.48 x LB. wMP 8 VI .3-2 Propellant Usage as a Function of d Figure -0 MARTIN MARIErTA CORPORA7fON DENVER DIVISION PK34-113 Page VI-31 where : r = 6.4 ft., moment arm n = no. of thrusters f ired = 0.93 lb-sec, , minimutn impulse bit Imin 4 I = 5.43 x 10 ft-lb-sec2, pitch mount of inertia YY Qd = 8.73 x lom3 red., pitch deadband I 5 170 sec., pulse specific impulse s P from which : The total undisturbed symmetric limit cycle propellant is: A North American Rockwell (NR) report, IL190-320-S & C-68-70, 8 Aug 1968, includes a factor of 1.1 of minimum impulse varia- tions based on simulation results and gives a propellant consumption figure of 3.388 lbjhr. The value computed above closely agrees with the NR value and is considered a good estimate. It is also seen that most of the propellant is consumed by the limit cycle about the roll axis and that the effects of gravity gradient disturbances in the pitch axis are inconsequential for nomina 1 parameter va lues . However, changes such as a reduction of the deadband to 0.1 deg. could substantially increase ?itch axis propellant usage as the nominal operating point would be on the nonlinear portion of the curve. The third alternative, an independent auxiliary system mounted in Sector I, would also utilize a horizon sensor/gyrocompass system as the attitude reference. Electronics similar in function to the OSS would be necessary to process the sensor signals and provide output commands to the cold gas valves. A simple block diagram is shown in Eigure VI.3-3 where gyro- compassing is assumed but resolver circuitry is unnecessary. The cold gas valves could be oriented to avoid any direct im- pingement into the field of view of experiment sensors a.nd would nut produce combustion product contaminants. It is also possible MA67IIIV MALsRIEl7A CORPORATION DENVER DI'/ISION GYRO PACKAGE A c HORIZON COLD GAS w ELECTRONICS VALVES A + I I, ELE CTXICAL POWER CONTROLS & STATUS DISPLAY Figure VI.3-3 Auxiliary Command and Service Module System For Lunar Orbital Experiment Orie~tatio~ MARNN MA6TIZ-A CORPORATION DENVER DIVISION PR34-113 Page VI-33 to substantially reduce the minimum impulse with a cold gas system. With a value of Wm = 0.1 lb-sea., extremely low limit cyc1.e rates result and substitution in the preceding equation for W gives: PT WpT = 0.014 Iblhr. tIowevcr, this value and the associated rates are so low that crew motion disturbances could be expected to be the determining factor in propellant usage rates. The low limit cycle rates would make it possible to reduce the attitude deadbands for more accurate pointing without incurring significant propellant weight penalties. An estimate of subsystem add-on weight and power for both the OSS/RCS and auxiliary system is given in Table VI.3-1. The additional component weight of 25 lb for the auxiliary system compared to the OSS/RCS would be partially offset by lower attitude control propellant requirements. However, the total weight penalty of the auxiliary system would exceed that of the OSS/RCS up to the point where sufficient propellant is included in both systzm weights for 75-100 hrs of operation. VI.3-2 Subsatellite Mounted Experiments Spin Stabilized Sub-Satellite The attitude control study of a spin statilized satellite reduces to the questions of spin axis stability and spin axis orientation accuracy. Two satellite configurations, Pioneer and ATMP, have been briefly investigated and some pertinent characteristics are presented in Table VI.3-2. It is well established that stability of the spin axis in the absence of external torques requires that the spin axis coin- cide with the principle axis of maximum inertia. This is a matter of obtaining the proper mass distribution during design and fabrication and not a question of feasibility. An area requiring further study is the effect of the external torque resulting from gravity gradient dynamics. A q~ar~titative evaluation of this is beyond the scope of the present study but the results cited in a NASA report, NASA CR-433, Effect of Gravity Gradient Torque on the Motion of the Spin Axis of an Asymmetric Vehicle, April 1966, indicate that perturbations could be very small. Further, the magnitude of the torque is a function of the vehicle mass distribution and is, therefore, subject to control. MARTIN IWARIEIIA CORPORAIION DEWVER DIVISION PR.34-113 Page Kt-34 Table VI.3-1 Command and Service Module Attitude Control Add-on Alternates oss /RCS Auxiliary Sys Weight Power Weight Power Horizon Sensor Sys ten1 22 lb 20 W 22 lb 20 W 3-Axis Gyro Sys tern 15 4 0 15 40 Interface Electronics 10 10 13 15 OSS 5 5 Gas, Valves, Plumbing, Tank 27 Total 521b 75W 77 Ib 75 W Table VI.3-2 Subsatellite Characteris tics Attitude Control Lunar Orbiter AlMP Pioneer A Spin Axis Orientation 0.3 Deg Accuracy Spin Rate NA 25+5m 60~~ Impulse Capability 8.6 x I& lb-sec 2.1 x lo4 0 Ib-sec Spacecraft and P/L Weight Attitude Control Propellant A - Estimated B - Dependent on spin-up and separation mechanism MART#N MARIEWA 00RPORATlON DENVER DIVISION PR34-113 Page VI-35 The Pioneer system includas a cotnmand link, electronics and a single gas jet. This system makes it possible to adjust the orientation of the spin vector about either or both of the two axis orthogonel to the spin axis by pulsing the gas jet at the proper fixed point in the spin cycle. Control is furnished by earth commands and the necessary attitude reference is the received signal strength from the vehicle high gain antenna. The relatively narrow antenna pattern leads to an estimate of 0.3 deg. for the acllievable orientation accuracy. This techni- que has been succesnfully used on past: Pioneer flights although ambiguity in the apparent attitude may require some trial and error pulsing of the gas jet to determine attitude from the resulting response. The Pioneer configuration does not include a propulsion system and would not be capable of performing orbital changes. The AWconfiguration does include a propulsion system capable of performing a maneuver along the fixed thrust vector axis. However, the AIMP does not include a system for adjusting the spin vector orientation. Thus, the spin characteristics, including precession and nutation of the spin vector, would be determined by the separation and spin-up mechanisms which establish the initial conditions. A paper presented by Lunar Orbit Mission Analysis for the Improved Delta Launch Vehicle and AIKP-D Spacecraft , AIMA 4th Aerospace Sciences Meeting, June 1966, cites i:hree sigma elevation and azimuth flight path angle errors of 1.1 deg. each during engine burn of the spin stabilized stage. Spin-up is achieved on a spin table prier to separation from the preceding stage. -Stable Subsatellite The Lunar Orbiter Configuration was investigated as an available satellite with an active, non-spinning: attitude stabilization system. This vehicle was devel.oped ta photograph discrete, selected portions of the lunar surface. The attitude control system is of course compatib1.e with this particular set of requirements. A simplified block diagram is given in Figure VI.3-4. In operation, the system is inertially stabilized with solar arrays oriented toward the sun. The inertial hold orientation is established by the vehicle-sun axis detected by the sun sensor system and the vehicle-canopus axis detected by the star tracker. The programmer initiates required functions through time reference sequential stored commands or real time earth uplink commands. Maneuvers are performed by summing rate mode gyro output with a fixed "sleeving" voltage to command the dWAW7#,% MAWIETTA QORPBRArfON UENVER DIVISION iiil! (BODY FlXED) I r N SiJX t Z PROGRuPER SEKSOR ELECTRONICS JETS SYSTEW i A STAR (CALOPUS) PI TUCKER > J * I A 7 SUN V/II PRESEPT D1C AHGLE I Figure VI.3-4 Lunar Orbiter Block Diagram AIaRTICY MARIE-& CORPORANOW DENVER 'JIVISION gas jets. Maneuver attitude excursion is determined by integra- tion o? the rate signal. This maneuver procedure must be followed to orient the cameras to the nadir for a photographic sequence and than reorient to the basic inertial attitude for reacquisition of the sensors. The various orbital reratLonships are depicted schematically in Figure VX.3-5 where angular relationships are greatly simplified for clarity. The plane of the earth's equator is, of course, in- clined at about 23 dagraas to thrl plane of the ecliptic and the moon's orbital plan^ is inclined 5 degrees to earth's equatorial plane. Tha star Canopus is in the region of 50 degrees south latitude. It is adequate for system illustration to consider that the sun (s), earth (E) and moon (M) all lie in the plane of the paper of Figure VX.3-5, Canopus then is roughly along the line to the reader's eye. The low inclination orbit of the Lunar Or- biter mission then also liea roughly in the plane of the paper and is denoted by the circle around M in the position labeled "Launched." Although the sun is occulted by the moon on each orbit, Canopus is continuously visible from the orbit plane snd the plane of the orbit does not deviate radically in its relation to the sua during the annual rotatton of earth about the sun. 1n the mission presently considered, however, the lunar orbit plane would be inclined at 90 degrees. This 3.8 shown in Figure VI.3-5 by the line through M in the "Launch" position representing an edge view of the oxhit plane--assr~med for convenience to be about the terminator at this time. Unber this condition, Caaopus is occulted by the moon on each orbit. Further, in the next posi- tion, "3 months" later, the orbit plane has rotated through 90 degrees relative to the sun. Conti.auing on to "6 months ," etc. it is seen that the orbit plane rotates 360 degrees relative to the sun in a year, correspoding to the earth's one year orbital period. As a result, the inertial reference frame established by the lunar orbital sensors is not suitable for a long duration mission in a polar lunar orbit. Wlrther, the sensors establish an inertial frame rather than the moon centered coordinate fra- reqvired for continuous experiment operation. Even if eech sen- sor were somehow provided with 360 degrees of angular freedom to maintain the line of sight to the reference source, the computa- tions required to determine the direction of lunar local vertical from the inertial frame would require some general purpose dig ital computer capability. The programer in the present system does not have this capability, These considerations consider that it would be necessary to replace the existing Lunar Orbiter attitude reference system. The best alternative would again utilize a horizon sensor sy~temto provide a closed loop lunar surface attitude reference. The resulting system wou1.d be essen- tially as previously shown in Figure VI.3-3, although further MAWrIN MARIEITiB (CIORPBRATCQN DENVER DIVISION POLA3 ORBIT PLAhiE0 / (j of G LAUhTCiI i MOhTliS Figure VI.3-5 Lunar Orbikl Relationships MARTfN MARIE-& CQRPQRAr#ON DENVER DIVISION PR34-113 Page VI-39 studies could be expected to lead to the addLtion of momentum storage device with the cold gas system used for desaturation only. Figure VZ.3-5 also shows a problem with respect to the present, fixed, solar arrays. To provide the mtieipaced electrical energy requirements for a 30-day mission En a polar orbit with modes in the region of the earth-moon line it would be necessary to provide a single degree of freedom gimbal system for the so- lar arrays. The axis of rotation would be normal to the vehicle orbit plane. For long duration missions it would be necessary to add a second degree of freedom with the axis of rotation par- allel to the orbit plane. Due to the elow rate of rotation re- quired about this second axis it could be mechanized with dis- crete changes in response to earth commands. MAIRTIN MARIEWA UORPORAlION DENVER DIVISION PR.34-113 Page VI-40 V1.4 THERMAL CONTROL SYSTEM - Thermal control sys terns for three subsatellite candidates and the WMwere reviewed and compared to experiment requirements for several sample groupings generated during the study. The impact on the thermal control systems of the four carriers is as follows: 1. The first and third CMM groupings shown in Table VI.4-2 can be passively controlled, with the exception of the laser al- timeter which requires its own cooling system (or interface with the carrier). A second C&SM grouping also shown in Table VZ.4-2 represents a far greater heat load and will require an active coolant sys tem,, 2. Three experiment groupings were analyzed for a stable satellite (Lunar Orbiter). Two of the groupings *posed no stringent thermal control requirements on the satellite systems . However, the third grouping required an active coolant system in addition to the passive Lunar Orbiter system. 3. Two spinning satellite experhent groupings were analyzed and compared to the them1 control system of the Pioneer and AIMP satellites. The existing passive thermal control, systems appear adequate for these groupings. The analysis results show a great dependence of the thermal con- trol systems of the carrier vehicles on experisrent groupings selected. VI.4-1 Exieting Thermal Control Configurations - The existing thermal control systems on each of the three subsatellite candi- dates are presented in Table VI.4-1 and briefly discussed in the following paragraphs. Anchored Interplanetary Monitoring Platform (AIMP) - The AIMP satellite utili.zed a passive. thermal control subsys tem (TCS) which consisted of a fiberglas heat shield located between the spacecraft, and the retromotor, multi-layer insulation blankets, and thermal control coatings on the internal components and external surfaces. The multi-layer insulation blanket consists of between 20 and 30 layers of three mil mylar all.uminized on one side as the outer shield and 114 mil mylar aluminized on both sides as the inner shield. The small package6 (less than 1 ft2) within the satellite were protected with 20 to 20 layers of crinkled (NRC-2) with aluminum vapor depo~itedon one side only. Larger components (more than a few ft2) utilizec? 20 to 30 layer mylar shields with either dacron or nylon separators between the MARrlN MARIKWA UORPORArION DENVER DIVISION Table VI.4-I Thermal Control Requirements for Satellite Missions r Satellites Parameters A-IMP Lunar Orbiter Pioneer System Pass ive Passive & Heaters Passive & Active Heat Loads, BTu/HT Orbit 170 1300 to 1400 290 Thermal Barrier Heat shield between space- Silicone oxide on aldnhed Mulei-laper craf tjretromotor (f iberglas) mylar insulat5on Multllayer insulation 3 layers dacron I blankets blankets 3 layers sluniinized mylar Coatings Internal Components painted black External White paint oC= ,363 Silicone base paint with €9 -85 zinc oxide Cat-A-Lac black paintK= .95 Pignaent E= .9, oC= -22 E= .87 Buffed aluminum oC= .15 €= .04 Temperatures, OF Prelaunch 77 -5 Orbit Internal 41 to 122? 35 to 85 30 to 90°F - External 30 to 425 MAmT#U MAR#.CITA QORCOSPAT#OU DENVER DIVISION PR34-113 Page VI-42 either dacron or nylon separators between the aluminized mylar shields . The internal components within the satellite were coated with Cat-A-Lac black paint, while the external surfaces (top and side) wore primarily coated with Dow Corning's white paint, The optical properties of the coatings are respectively oc .95, € = .87, and oC - .363 and E - .85. Post flight data evaluation revealed that the increase in tem- perature of certain components was due primarily to absorptivity changes, which increased by a factor of two. A ma,lor contribuLor to the absorptivity change was contamination of the white painted surfaces by the exhaust products from the retromotor. NASA-Goddard is presently elFluinat ing my use of white painted surfaces and will utilize tefl.on coated with either vapor depos- its of aluminum or silver. Future flights are also (;onsidering Kapton instead of mylar, and the application of heat pipes. The average heating load to be dissipated on these satellites was 170 btu/hr. Lunar Orbiter - The Lunar Orbiter satellite utilized a passive thermal control gyatem and mounted electric heaters on the critical coniponents. A heat shield barrier and multi-layer in- sulation blankets, consisting of 3 layers of aluminized mylar with layers of dacron separator8 were used. The outer mylar shield was coated with silicone oxide. The equipment mount deck sur- faces were also coated with white paint, containing zinc oxide pigment and a silicone base. The optical properties were abaorp- tivity,oC=.22, and emissivity, € m.9. Post flight data evaluation showed that the internal temperatures frequent::ly exceeded the limits of 1080F'. The high temp.eratures, as for those of the AMP, were attributed to degradation of the external coating absorptivity property. To stabilize the temperatures during orbit, the attitude control subsystem was utilized to change the vehicle sun line attitude long enotgh to reduce temperatures. The internal heating loads ranged between 1300 and 1400 btu/hr. --Pioneer - The Pioneer spinning satellites utilized primar- ily a passive. thermal control subsyetem, which consisted of multi- lavnr insulation blanketa. Some of the vehicles had proviaiona for an active system in the platform area, which consisted of louvers fndividually actuated by bimetallic springs, which sensed the temperature of the equipment platform in the region of the louvers. MARGIN MARIE-A 00RPORAlION DENVER DIVISION PR34-113 Page VI-43 VI.4-2 CGSM Tl~ermalControl Sys tem Analysis Three experiment groupin{-8 were investigated. The first two were given to North American Hoclcwell early in the study; the third grouping was generated later. All three groupings should be ~orlsidercdas representative samples of poterltial C&SM experiments aqd are listed in Table ~1~4-2for convenience. Al- lowable tempcraturv limits for a11 experiments are given in Table 111,4, Experiment Requirements Summary. Three typical duty cycler were generated based on the groupings shown in Table VI.4-2. These are shown in Figures VI.4-1, VI.4-2, and VI.4-3 for groups 1, 11, and I11 respectively. The aversage heating load for the first C&SM experiment grou~irig (Table VI.4-2) is 715 btu/hr, with a peak of 2700 btu/k:lr for a duration of 1 hour. The peak load is due primarily to the opera- tion of the Panoramic camera. However, the camera has sufficient thermal capacity to absorb the loac, and maintain its components within their allowable temperature limits. The Jaser altimeter experiment appears critical for continuous operation, and a cooling system is recommended for this experiment only. For the second experiment grouping, the average inter- nal heating load to be dissipated is 1260 btu/hr, and peak loadings between 2330 and 2430 btu/hr will be generated. The average 1.260 btu/hr could be dissipated by the present Apollo Bl.ock I1 C&SM coolant loop and environmental control subsystem (ECS) radiator configuration. The radiator has a maximum capabil- ity of approximately 8425 3tu/hr with an effective total radiator area of 56.0 ft2. The mi~sionorbital requirements for the C&SM are approximately 6500 btu/hr, leaving a margin of 1925 btu/hr. However, to avoid extensive requalification of the C&SM environ- mental control syntem due to experiment interface, use of a separate experiment coolant loop is recommended. A schematic of a TCS cs~ybleof dissipation of the average heating load is shown in Figure VI.4-4. The coolant loop wculd utilize the fluid E-lat a coolant flow rate of 400 lbs/hr. As shown, the system would include a by-pass arrangement, temperature control valve and two of the present Apollo 3lock TI LM pumps. The radiator size required would be 21.0 ft . Cold plates are probably required for six experiment components. These are: (1) IK happing Radi- ometer, (2) MW Radiometer, (3) Metric Camera, (4) RZi Reflectance, (5) Radar Altimeter, and (6) Radar Imager. The total weight of the system would be 169 pounds. IWAIPrlN MARlSrrA UORPORAllON DENVER DIVISION TABLE ~1.4-2 C&SM-EXPEZIPIENT GROUPINGS GROUP I GROUP I1 GIii)OP- I11 Metric Camera Mapping Radiometer Metric & Stellar Camera Stellar Camera IR Spectrometer Panoramic Camera Panoradc Camera Electroaics fhotoelectric Photometer Phot~electricPhotometer MW Rzdiometer Laser Altimeter Laser Altimeter Metric Camera RF Transponder RF Transponder Stellar Camera Solar Wind Foil Solar Wind Foil RF Reflectance Extender Bistatic Radar Foil Radar Altimeter Geochemistry Expts. RF Transponder Gamma-Ray Radar Imger Neutron Albedo Antennas Gamma Particle Data Processor X-Ray Sensors ~rans/Rec. Solar X-Rag Film Record (optional) Cos~cRay U.V. Imager Electronics Imager I Film Recorder I. MIORTJN MAR#-A CORPORANOW DENVER DIVISION I 2800 2700 BTU/HR 2400250; 1 2200- 2mo- I 1800 1600 1400- 12004 1000~ I - 800------_- -__- ______- - - AVG- 715 - -BTU~ - 600, -- 200 L t I I I o 2 4 6 8 io i2 li 1; 18 2b 2i 24 TIME + HOURS Figure VI.4-1 Heating Loads - WE3 GI-p I Experiments MAIP'TIN MAR#-& CORPORA78OW DENVER DIVISION TIME r' HOLqS Figure VI.4-2 Heating Loads - C6SM Group I1 3bpnrhents MARrJN MARIEITA CORPORATJOW DENVER DIVIS!OV 0 I-:1 -7-,-1------r--- 0 2 4 6 8 1Q 12 14 16 18 20 22 24 w '3 i3 8P C- nm r' HOURS UAnw c.1 W Figure lT,4-3 Heating - C&SM Group 111 Ever-nts 4 MARr8N MARIE774 CORPORAT8ON DENVER DlVlSlOU Qout Area CONTROL VALVE COOLANT A E- 1 COOuwT FLOW UTE/ 4008iiB SYSTEN WFIGHT // 169 LBS, Figure VI.4-4 Maximum H9ating Load Condition - C&SM Group I1 Experiment Coolant Sgste31 MARTIN MARIEBA CORPORA7-ION DENVEZ DIVISION PR34-113 Page VI-49 The IR mapping radiometer is seen as a thermal control problerri, :since the experiment requirements apecify that it8 op- tical mirror be maintained at 150°~by radiation t echniquee. This places a requirement that tne experiment radiator viewt deep epacs during experiment operation, The orbital environmc~nt otherwise will not enable this temperatare Level to be maintained. The average heating load for the thirh C&SM experiment group is 545 btu/hr, with a peak load of 2460 btu/hr when a11 experiments are in operation. Operation of the experiments of this group- ing should not require interfdcc with the C&SM thernal control system with the pomible exception of the laser altimeter. Its con- tinuous op-ration will cause a component heating problerr. As in Group I, a separate cocling system sho~ldbe cocsldered for this experiment. All experiment and subsystem components should re- ceive coatings similar to those of the Lunar Orbiter satellite components, with good therm~lconductioc paths to the mounting structure. VI.4.3 Stable Satellite (Lunar Orbiter') Thermal Control Analysis - Three candidate experiment: groups were i.nwstigated for use on a stable subsatellite. The grouping^ are presented in Table VI.4-3. Table VI.4-3 Stable Satellite Experiment Groupings - .-- Group I I Group 11 I Group 111 Geochemistry Geochernis try IR Mapping Radiorne ter Mass Spectrometer I Mass Spectrometer I IR Spectrometer Radar /Las er RF Transponder MW Radiometer A1 timeter Meteoroid RP Reflectance RF Transponder/ Laser Bi-S tatic Radar Retroreflector Radar A1 t heter EM Measurements RF Transponder Noise Survey EM Measurements Gravity Gradient Noise Survey Meteoroid- MARrlN MARIEWA 00RPORA7ION DENVER DIVISION PR34-113 Page VT-50 Hatinr: lori,i:; :'or ( ttv I'i~wt; two r,*:lble :iatt*lli kt. b:r.ou~~irtj::, ur1la ti:; foll.ow::: Hea tixg Lofidn (btu/ hr) Group I I Group I1 Experiments '3 314 Subsystems TOTALS 324 These heat loads are compatible with the Lunar Orbiter passive thermal contro: system. Therefore, with the exception of local insulation, no changes are required. The third atable satellite experiment group presents a greater thermal control problem. Figure VI.4-5 sl-,ows the heating loads for this case. An active coolant loop circuit is required to dissipate these heating loads. A schematic of a typical sys- tem is shown in Figure ~1.4-6. The coolant fluid recommended is E-1 at a flow rate of 600 Ibs/hr. The coolant flow is divided equally to the subsystem and experiment components. The radiator size re uired to dissipate the average 1840 btu/hr to space is 30.0 ft3 . The size of this radiator poses a significant integra- tion problem. Cold plates are probably required on five experiment packages. These are: 1. IR mapping radiometer 2. MV radiometer 3. RIP reflectance 4. Radar altimeter 5. Electromagnetic measurements Also, some subsystem components will use cold plates. The total weight of the TCS is estimated as 192 lbs., including 20 Ibs. of nulti-layer insulation blanket. MAlPllN MARIETTA 0aRPORArION DENVER DlVlSlON PEAKS 2167 BTU/HR 2200 1 -- ?----I- ,..-,---,-- TIP 7.- - 0 2 4 6 8 10 12 14 16 18 20 22 24 TIm HOURS Figure VI.4-5 Heating Loads - Stable Satellite Group 111 Experiments MARrNN MARIE-A CORPORA7NON DENVER DIVISION ,. - -rC 1 I Tin = 73-S0F E-pASS ZERO FLOW - Qout=lW BTU/HR Ag 4 EXPERIMEEJTS COLD PLA'ES AREA = 30 FT 1 Tout = 61.2OF T = 61.2OF *\ TCW COIrnOL VALVE COOLANT 4 E- 1 COOLANT FLOW RATE j 600#/HR SYSTEM WEIGHT / 192 mSo VI Figure VI.4-6 Maximum Heating Load Condition - Stable Satellite Group 111 Experimerc Coolant System N MARTIN MARIETTA CORPORATION DENVER DIVISION PR34-113 Page VI-53 VX.4-4 Spinning Satellite (AIMP! Thermal Control Analysis - Two candidate experiment groups were investigated for the spinning satellite missions. These are 11s ted in Table VI.4-4. Table VI.4-4 Spinning Satellite Experiment Groupings Croup I 1 Group I1 Magnetometer (D.C.) Magnetometer (D.C.) Magns tome ter (A, C. ) Magnetometer (A.C.) Plasma Probe ' Plasma Probe Electric Field Electric Field Low Energy Particles Low Energy Particles Solar Energetic Solar Energetic Particles Particles Meteoroid RF Transponder RF ~ransponder/I,aser Retroref lector The average heating loads for Groups I and TI are respectively 135 and 279 btu/hr. These heating loads are within the capa- bllities of the passive thermal control techniques applied on the AlMP satellites. MARrfN APARIEWA OORPORATfON DENVER DIVISION