Electric Propulsion for Geostationary Orbit Insertion

NASA Technical Memorandum 106942

Electric Propulsion For Geostationary Orbit Insertion

Steven R. Oleson NYMA, Inc. Engineering Services Division Brook Park, Ohio

Francis M. Curran National Aeronautics and Space Adminstration Lewis Research Center Cleveland, Ohio

Roger M. Myers NYMA, Inc. Engineering Services Division Brook Park, Ohio

Prepared for the 30th Intersociety Energy Conversion Engineering Conference cosponsored by ASME, IEEE, AIChE, ANS, ACS, and AIAA Orlando, Florida, July 31--August 4, 1995

N95-27727 PROPULSION FOR GEOSTATIONARY ORBIT •,_ (NASA-TM-106942)INSERTION (NASA. ELECTRICLe_is Research Center) _ p Unclas National Aeronautics and Space Administration G3/20 0050100

ELECTRIC PROPULSION FOR GEOSTATIONARY ORBIT INSERTION

Steven R. Oleson Nyma Inc. NASA Lewis Research Center Brookpark, OH 44142

Francis M. Curran NASA Lewis Research Center Cleveland, OH 44135

Roger M. Myers Nyma Ir_ NASA Lewis Research Center Brookpark, OH 44142

the mission analyses, propulsion options, and the results for ABSTRACT four geostadonary insertion options. Solar Electric Propubion (SEP) technology is already being used for geostationexy satellite stafionkeeping to increase payload mass. l By using this same technology to perform part OPTIONS AND of the orbit Iran.ffer additional increases in payload mass can MISSION ANALYSIS, ASSUMPTIONS be achievecL Advanced chemical and N2H 4 arcjet systems are used to increase the payload mass by pc_orming Mission Analysis statienkeepin8 and part of the orbit transfor. Four mission olxiom are analyzed which show the impect of either _aring The approach is to consider various sub-GTO orbits the orldt Inmsfe_ between chemical and SEP systems or (including GTO) _s starting points for the SEP raising and having eithor complete the transfer alone. Results show that plane changing. It is assumed that the launch vehicle, in this for an Atlas IIAS payload increases in net mass (geostatinnary case the Atlas KAS, places the payload satellite (including satellite mass less wet propulsion system mass) of up to 100 the necemaxy on board propulsion to achieve geostationary kg can be achieved using advanced chemical for the transfe_ orbit) into some elliptical transfor orbiL Elliptical transfer and advanced N2H 4 mejets for stationkeeping. An orbits which have their apogee at geostationary altitude additional 100 kg can be added using advanced N2H 4 arcjets (36,000 lan) shown in Fig. 1 me called geostationary Iransfer for pen of a 40 day orbit tmmfor. orbits ((}TO) and _ regularly used today. The transfer ellipse apogee can also be lowered (trainedhereas sub-GTO) INTRODUCTION to increase the SEP starting mass. Solar Electric Propulsion (SEP) is already being used for stafionkeeping of goostafionaxy satelfites, most notably the initial orbit is GTO or mb-GTO, the orbit AT&T's Telstar 4. The next step is to use these types of Uamfor can be "split" between the on-bom'd chemical and tlmmors to contribute to placing the spacecraft into SEP systems. While the SEP system can complete the geoslationary orbit. For a given launch vehicle the fuel mass mission alone, the tranffor time is around I00 days (see savings can then be directly used to increase the payload Results section) and significant portions of this time are spent (e._., number of communication transponders). Even a mall in the Van Allen belts. Avoiding the dense parts of the Van increase in mass (100 to 300 kg) might have large revenue Allen belts Ceelow --10,0(30 kln) and large radiation doses,is earning impacts, This study evaluated the mass impact of of iximaty conecm Consequently, Imrig_s me used which replacing some portion of a geostationary speccor_s are above these portions of the radiation belts. chemical apogee propulsion system with an N2H 4 arcjet Mission ODtlOnS and Assumptions system with the N2H 4 arcjet system also performing fifteen years of slationkeeping No attempt was made to optimize the The top portion of Table I. shows the various missions used missinn in this worlc All the inlmts as well as the remits can inthispeps. While by no means exhaustive,they illustrate be found in the text or Table 1. Each section of the paper the potentialsystem peffommace. The initialorbitdcfmcs descn'bes a different portion of the Tablel This paper describes wherethespacecraftstartsafl_separationfromtheAtlasHAS Cmtm_. 2 The on-hom_ ¢hemicM sylm thm _ the typically minor,me not co_ here. mtellite to the inmmed/ate od_t where the SEP system takes over and completes the 8eostationary orbit insertion. Gemmfionmy orbit is assumed to be a 35786 Ion circular SYSTEM ASSUMPTIONS AND MODELING orbit at & inclinmion. _tstlon Keeolno and Final Orbit Delivery The first ce_s, dmoted as baseline, admnced baseline and admm_+ ba_line, show the misskm xonafio of the launch Forthisp_= o_ mte-of-mandadvffi_d _2_ arcjot_ vehkle plaaing the ntellite imo GTO end the on-hoard Ke considm_! for the SEP system. Fiflem years of propui_ system _ the poe.gee to eqwl the north/south _ sts_o_-ping is pm'om_! by four Wosee (c_n) and dumgJ_ the plane to 0°. This one pair placed on the north face8rid the other on scenmio is commonly used today. Admnced and advanced+ thesouth face (soe Fig. l).Thase thmsterpa_s m canted 17° denote the use of 574s and 622s arcjets, respectively, as fromthe verticalthe minimizeplume interactionwith the eWlataed the the sya_m assumptions section. In addition, may. To perform the north/south smtionkeeping either the dl _ ¢X_p( the baseline me 8dvsoced che_cal muth or north pair is fired about the approgiate orbit node technology. The n=a two cases, caUed advanced b_gh on the order of minutes. Four _ IspS ere assumed: e///pt/c md adganced+ h/gA elliptic, again use a OTO orbit 478s (_te-of-a_), 574s (advanced), and 622s (advanced+), but this time the on-board chemical lm_pulsion system only all of which have bern adjusted for the 17e thruster cant raises the perigee to 11,000 ion and the changes the plane to cosine los. The input powers for the power processing 9". As shown in Table 2, this division 'of plane change was units (PPUs) are 1.8 kW (state of the art) and 2.39 kW found to mushly _ the delivered mass in GEO using (advanced).Four PPUs supportthefour NSSK thrusters. the on-hoard chemical end efcjet system_ The SEP arcjet system then delivers the spazect_ to geostationary orbit. Four additionalN2H 4 arcjetsare added for pe_orm/_ the perigee raise and plane change mission. Thesethrustersare The nextcases,termedarcjetto GEO and arcjet+to GEO, assmned identical to the NSSK _ exceptthey ere have a missionscemrio whae the SEP system perfonusthe placedaboutthe chemicalthrusteron the aftportionof the wholeper_eermeandplanechimse. No on-board cbemiml spaceerafl(seeFis.I)_Theselhru.qeNsharetheNSSK PPUs. mmmvers ererequiredforthme ¢ase_ The inputpower forthesethrustmuis2.39kW and theIsps ere either574 s or 622 s. The combined efcjetand PPU The final two cases use a sub.GTO orbit and me called efficiency is 32% During pensee raisin8 and plane changing admn¢_ high ¢ircwlar told _+ iu'gh ¢ircwlar. The four N2H 4 arcjots are ruinS. Each thrust_ unit includes on.bonrd chemical system circularizes tl_ orbit at 15,000 ion structure sad _ller and weighs 1.86 ks. Each PPU unit ned c_ the plane down to 9°. (This choice of a 9 ° incha_ cabling and thmnal system and weighs 6.08 kg. intonnodime piano is made breed on the dam of Table 2.) The N2H 4 afcjet _ life is ]000 hours (41days). S_P s_ thm completes the transfer. _hamlcal On-Board Prooulslon F'u'tem yam of stationkeeping ere _ for every The assumed system is a rote-of-art 314.5 s bipropellant spacmm_ While the yearly AV varies with satellite station system 7 for the baseline and an advanced 328 s bipropellant longitude, 50 m/s is chosen as relmumltstive. 3 The system 7 for the other cases in Table 1. Both systems have a additional cosine losses mcomtered by not completing the dry mass of 23 k8 and a tankage fraction of 0.08. This whole tram at the orbit node (impulsive bum) are small and chemical system is deleted from the spacemafl for those neglected. missiom where the arcjets complete the perisee raise and plane change entirely. The mission AVs (vdocity or energy clmnge required for orbit Powsr System mmfer) for the on-bored cbcmkal systan are assumed to be impu_ve. Tho Imn_ermission AVs for the SEP system differ The GaAs solar arrays which provide payload power in fium impulsive due to constant _g4 and e_e obtained geostafionary orbit ere asmma:d to provide the 9.56 kWs for using the SECKSPOT 5 numerical optimization program four thruster operalion during the SEP orbit transfer since the along with various analytical spreadsheets. For the SEP payload is inactive dining this phase. The battew system is portion of the mission the effects of shading, power ammned to be capable of providing the 3.6 or 4.78 kW for dqpldation, and oblateness me considered. The SECKSPOT two NSSK thruster operation while the payload uses direct FOlFam .t.,.._.._ op/inal steenn8 for a min_um time solar anay power as suggested by Free. 8 The arrays m'e trajeeteo/. The impaas of non-opt_®al steering and assumed to be shielded with 12 mils front and 12 mils back. guidance, navigation, and altitude control limitations, while

2 RESULTS The figm_ ofmer_ oftl_advaaced pmp,]sion sy_ms in this Thedegmdation to the amrysis _ to bearound 3% mxly is m mmd_w=_L Net _ _=s to the useable for the high elliptic and hish ci_zular mission scenarios, and melli_¢ ,-,,*. oace Ibe wet propulsion 'sys_m is removed. 15 % for the arcjet GTO to OEO tnmsfer. No mdiat/on dose Th_ added net mass cam be used for additional equipment calculatm on the payload has been made. While not iacludin8 _ tnmsponde_ rims providing considered ;, Table I, some pert of this loss of may power add/tiomd chmue_ and increased revenue. Table I presents might be charged to the propulsion sysmn. This loss in net the results nmnmically while Figures 3 and 4 graph the net mass fraction should be less flum 20 k8 for the 3% mms md EPUip_ d_sdation f,a,qcs.

Table I. ¢ontlim the resultsof thismalysis Note the CONCLUSIONS bascllnemissionwhere theorbittmmfer iscompletedsolely Additional net mass cm be delivered to geostafionary orbit by the owbomd chemicalsystem end the stationkeepin8 to iaereme the payload mass md communication satellite perfmnedbythe_jet revenues by using advanced on-lx_rd chemical and elece_l propulsion. Up to 100 kg additional net mass is achievable The next two cases, advanced baseline and advanced+ with advanced chemical propulsion for orbit tnmfer and &z_//ne, show the _q_ct of upgrading the propulsion advanced mcjets for NSSK. To achieve 200 k8 net gain, a systemswhile_thebaselinemission scemrio. Up portion of the transfer must be compleaxl by the higher to 100 kg of additional payload is achievable for a state of the performing arcjet system using available payload power. ett missio_ While lra_en l_formed by arc_ alone provide even more payload, the trm_er time is eripled md much more radiation is Toe remaining cases show the impact of using different encountered. mission scem_ios as discussed in the mission options sectio_ The high elliptic cases split the transfer mission Aeknowledomente between the chemical end arcjet systems using a high Resemch for this paper was done at NASA Lewis Research elliptical orbit which avoids most of the:radiation belts. For Center's Advanced Space Analysis Offs:e (Contract NAS3- ,,- aplxoximme 40 day Uip time, 70 to ]00 _ additional 27186). We are indebted to Timothy Wickenheiser and John payload is delivered compared to eq.ivalmt tech-ology Riehl for their lxofouad insights and 8uidm_7,e. using the nomal raison scenmio. Tbecost of this 40 day _st'er is not ¢omdsr_ here. Referencee I. Oleson. S_. "Influence of Power SysU_ Technology of The arc.jetG'TO to GEO and ar_et+ GTO to GEO cases Eleclric Propulsion Missions", CR-195419, NASA Lewis • ow a missionsc_mniowhere thearcjetsystemtnmsfen the Research Center, Janumy 1995. spac_cra_without a cltmnk_ on-boardprolx_ion system. 2. Mission Planner's Guide for the Atlas IAuu_h Vehicle While thistechniqueum add over I00 k8 over thenormal Family, i-GSB910585-68, Rev. 4, July 1993. missionscem_, the triptime ap_oeches fourmonths and 3. Agmwsl, B.N. Design of Geos_nchronous Smcecra_ is additional limespent in the radiation belts. Prentice-Hall, Inc. Englewood Cliffs, NJ. I 4. Edalbmnn, T.N. "Prop-i-_,_n Requirements for Controllable The last two h//_ ¢/rcw/ar cases show mis_ons us/ug a high Satellite_, _ Journal. 31: 1079-1089. August 1961. circular orbit which _ avoids the worn portions of the 5. Seckett, LL., et aL, "Solar Electric Geoc_lric Transfer radiation belts and completes the transfer in just over 40 days with Attitude Cons_a_nts: Analysis", NASA CR-134927, Unfortunately, no addit/onal net mass is provided, since the August, 1975. additional slatting mass is more than overcome by the higher 6. GL. Bennett, et al, "An Overview of NASA's Electric chemical and EP AVs. Propulsion Program',IEPC-93-006, September 1993. 7. Myms, _ et aL, "Small Satellite Propulsion Options", With a mimion life of 41 days of constant thrusting, only the NASA TM-106701, AIAA-94-2997, June 1994. ucjet GTO to GEO cases would require _ thrusters. 8. Free, B,A. "North-South Stationkeeping with Electric While this additional mass is neglected in Table l, the impact Propulsion Using Onboard Battery Power", COMSAT would only be four extra thrusters (not PPUs) and would be Laboratories, 1980 less than lOk8 total.

3 Table 1. Mlmdon Advanced+ Advm_ Advanced+ Ar_jet Al_+ Advmxd Advanced4- Optlone and Baseliae BaseJ_ High Iegh GTOto GTOt_ High Rooult8 mr GEO GEO Circular Chr.ular labild edit G'rO: 167 167 x GTO: 167 x GTO: 167x GTO:. 167x GTO:. 167 G'I_: 167 Sub GTO:. Sub GI_. x 35786 35786 kn_ 35786 x 35786 x 35786 167 x 167 x (_ hAS deu_e_ 35786 kn_ 35786 kn_ 27* 27" 27* 270 kn_ 27" kin@27" 15000 15000 od_) kin@27" kn_ 270 km@ 27° mgh I'nsh High High Circul_. Circular: Inmmed_: odit nolle none none Elliptic: Elliptic: none none II000 x 11000 x 15000km 1500O lun 35786km 9 ° 35786km 9° 9" 9°

T F'malOd_ GIn GIn GK) _Le GR3 (In

Cbm l"ram_ System (_Gal to S_op AdvSipep AdvBilxup AdvSipW none lnmmedi_ orbit) Mimkm AV 1806 m/s 1806 m/s 968 m/s 968 m/s 0m/s 0m/s 1762 m/s 1762 m/s 315 s 328 s 328 s 328 s 328 s - 328 s 328 s Fuel Mm 1588 ks 1539 ks 1539 931 ks 931 ks - 1965 ks 1965 ks System Wa mass 1738 ks 1685 ks 1685 kg 1029 ks I029 ks oks oks 2145 kg 2145 ks F..PTms_ Sys_ (lnusmedi_ to F._ None Nol_ N2H4 AJ N2H4 AJ N2114AJ N2H4 AJ N2H4 AJ N2H4 AJ orbit) Adv. Adv.+ Adv. Adv.+ Adv. Adv.+ Mission &V Om/s 0m/s 0m/s 1369 m/s 1369 m/s 3074 m/s 3070 m/s 1491 m/s 1492 m/s 0days 38days 42days 119 days 42 days 46 days lq, (_05 650s 600s 6.505 600| 650 s 55o ks 512 ks 1458 ks 1369 ks 603ks 562 ks SystemWetmm 596 ks 556 ks 1567 kg 1473 ks 653 ks 609 ks NSSK EP SYSTEM N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ N2H4 AJ SOA Adv. Adv.+ Adv. Adv.+ Adv. Adv.+ Adv. Adv .+ NSSK aV for 15 750 m/. 750 m/s 750 m/s 750 m/s 750 m/s 750 m/s 750 m/s 750 m/s 750 m/s NSSK_ (@ 17",_) 478 s 574 s 622 s 574 s 622 s 574 s 622 s 574 s 622 s 1.80 kWe 2.39 kWe 2.39kWe 2.39 kWe 2.39 kWe 2.39 kWe 2.39 kWe 2.39 kWe 2.39 kWe EPFudM_ 295kF 255 ks 236k_ 265ks 256ks 261 ks 247 ks EP P/Im Wa m_ 367 ks 338 ks 319 ks 330 ks 349 ks 340ks 345ks 330 ks

M_s inhdtislOd_t 3583ks 3583 ks 3583 ks 3583 ks 3583 ks 3583 ks 3583 ks 4660 ks 4660 kg 1996ks 2o44 ks 2o44 ks 2653ks 2653 ks 3583 kg 3583 ks 2695 ks 2695 kg Orbit

BOt,Mm_ GEO 1996 ks 20_kS 2o44 ks 2102 ks 2140 ks 2126 kg 2214 kS 2092 kg 2133 kg 1701 kS 1789 ks 1808 ks 1840 ks 1893 ks 1861 ks 1958 ks 1831 ks 1886 kg Combined NSSK & Transfa Propulsi_ 21o5 ks 2024 ks 2O04ks 1971 ks 1915 ks 1917 ks 1812 kg 3143 ks 3084 kg Systems Wot Mass

F'mal Net Mass 1479 I_ 1579 _q 1612 i_ 1669 kg 1_67 k_ 1771 i_g 1517 kg 1576kg Added Ne_ Mass 0kg +81 kg +100 kg +133 kg +190 kg +188 kg +292 kg +38 kg +97kg ov_ Baseline

4 GTO

GEO

i FIGURE 2. SPACECRAFT THRUS'F_.R CONFIGURATION

------Key:

0 Chemical Thruster

• _,= Arcjet Thruster

------Orbit Plane Aft (zenith) view

TABLE 2. MISSION PLANE CHANGES Initial Mass in Delivcrecl Mass in GEO Trip Tune Inclination In_ Orbit with 600 s AJ

27° 2751 kg 1789 kg 68 d i 21 ° 2725kg 1913 kg 57 d 12° 2608 kg 2020 kg 41 d 9' 2554 kg 2025 37 d if, 2497 kg 2011 kg 34d 0u 2374kg 1940kg 30d Figure 3. Net Mass Delivered (EOL Mass in GEO less Propulsion System) vs Mission / Propulsion Options

]mo _ T ]w] kg

1750 kg _ 1669 1667k m ]7oov_t v_ g P'/////A '650 kg Jl" 1612 kg _] _ !:::+++o:+im m+ RIg.m.N.N.m,m B m,,,,,+,+ J,

1400_ + ++++]+-+ ++++++++.+o<++o++.+

Figure 4. EP Transfer Trip Time vs Mission / Propulsion Option 119 days

I0012080 days I 108days _

60 days 40 days 20 days

0 days Ad_mccd A_mce_ Arc._ Ate+ Advanced Advanced+ mgh High _ to G'roto mgh mgh m_k _ (mo _.o c_cu_ Circular

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Electric Propulsion for Geostationary Orbit Insertion

WU-564-09-20 6. AUTHOR(S)

Steven R. Oleson, Francis M. Curran, and Roger M. Myers

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER National Aeronautics and Space Administration Lewis Research Center E-9671 Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER

National Aeronautics and Space Administration Washington, D.C. 20546-0001 NASA TM- 106942

11. SUPPLEMENTARYNOTES Prepared for the 30th Intersociety Energy Conversion Engineering Conference cosponsored by ASME, IEEE, AIChE, ANS, ACS, and AIAA, Orlando, Florida, July 31---August 4, 1995. Steven R. Oleson and Roger M. Myers, NYMA, Inc., Engineering Services Division, 2001 Aerospace Parkway, Brook Park, Ohio 44142 (work funded by NASA Contract NAS3-27186), and Francis M. Curran, NASA Lewis Research Center. Responsible person, Francis M. Curran, organization code 5330, (216) 977-7424. 12a. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified - Unlimited Subject Categories 13, 15, 16, and 20

This publication is available from the NASA Center for Aerospace Information, (301) 621-0390. 13. ABSTRACT (Maximum 200 words) Solar Electric Propulsion (SEP) technology is already being used for geostationary satellite stationkeeping to increase payload mass. By using this same technology to perform part of the orbit transfer additional increases in payload mass can be achieved. Advanced chemical and N2H 4 arcjet systems are used to increase the payload mass by performing stationkeeping and part of the orbit transfer. Four mission options are analyzed which show the impact of either sharing the orbit transfer between chemical and SEP systems or having either complete the transfer alone. Results show that for an Atlas IIAS payload increases in net mass (geostationary satellite mass less wet propulsion system mass) of up to 100 kg can be achieved using advanced chemical for the transfer and advanced N2H 4 arcjets for stationkeeping. An additional 100 kg can be added using advanced N2H 4 arcjets for part of a 40 day orbit transfer.

14. SUBJECT TERMS 15. NUMBER OF PAGES 8 GEO; Insertion; Electric propulsion; Stationkeeping; On-board chemical propulsion 16. PRICE CODE A02 17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified

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