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Report No. T-29 MARS SURFACE SAMPLE RETURN MISSIONS VIA ASTRO SCIENCES Report No. T-29 MARS SURFACE SAMPLE RETURN MISSIONS VIA SOLAR ELECTRIC PROPULSION IIT RESEARCH INSTITUTE 10 West 35 Street Chicago, Illinois 60616 Report No. T-29 MARS SURFACE SAMPLE RETURN MISSIONS VIA SOLAR ELECTRIC PROPULSION by D. J. Spadoni A. L. Friedlander Astro Sciences IIT Research Institute Chicago, Illinois 60616 This work was performed for the Jet Propulsion Laboratory, California Institute of Technology, as sponsored by the National Aeronautics and Space Administration under Contract NAS7-100, Task Order No. RD-26 APPROVED BY: D. L. Roberts, Manager Astro Sciences August 1971 NT RESEARCH INSTITUTE FOREWORD This Technical Report is the final documentation on all data and information required by Task 7: Mars Surface Sample Return Missions. The work herein represents one phase of the study, Support Analysis for Solar Electric Propulsion Data Summary and Mission Applications, conducted by IIT Research Institute for the Jet Propulsion Laboratory, California Institute of Technology, under JPL Contract No. 952701. Tasks 9 and 10 of this study will be reported separately. /IT RESEARCH INSTITUTE ii SUMMARY AND CONCLUSIONS This report describes the characteristics and capa- bilities of solar electric propulsion (SEP) for performing Mars Surface Sample Return (MSSR) missions. The scope of the study emphasizes trajectory/payload analysis and the comparison of mission/system tradeoff options. Questions concerning mission science objectives, instrumentation, operations and spacecraft design are not treated herein. Subsystem weights and scaling relationships used in the present study are based on previous independent studies. The MSSR mission is examined only for the 1981-82 launch opportunity. This opportunity seems to be realistic in light of current schedules for Mars exploration and SEP technology develop- ment. Several other study constraints which bear directly on the results obtained are: (1) return samples in the range 5-25 kg, (2) use of lifting (offset C.G.) atmospheric entry at Mars which allows a low ratio (1.25) of entry weight to landed weight, and (3) rendezvous and docking in Mars orbit. Major results of the study are presented as performance curves of Earth departure mass versus sample size for a number of different mission/system options. These options represent a spectrum of trip time, launch vehicle"capability, combinations of low-thrust and ballistic maneuvers, chemical retro type, and Earth recovery mode. Six mission concepts or baseline examples are selected from the parametric data. Table S-l summarizes the pertinent aspects of these baseline examples. All assume the direct entry option for the Mars lander vehicle, the Earth orbit capture mode for sample capsule recovery (555 x 9000 km altitude orbit), and the solid propulsion system for retro maneuvers. __ LIT_R.E S.E A R C,H_ IN S.TJ J_UJ1E_..__. iii O 0 cr c_> 05 05 ^* >-H £-4 ^^ 05 H 0 H ^ O Ol uo 01 < C r- o CO 05 If s """ r-l 5 00 IO o t_l -- 1 H CO o r-1 1 01 H 10 r-1 CO a a- \ r-i H h- 1 E • hH O 03 O o CO CO CO o o o 05 .05 r^-4 H O H i 01 oc IO v^ .' 1 g O' o O. 1 rt C CO IO ' H- 1 W H r-1 f- o r^ Q Q r-1 01 01 >— * )—4 a: H-4 LU 0^3 O O CO co i o - o o 0 O 05 05 10 UJ ^ 01 H-( E-t H ex I-H 05 fr] y l__ a: >— 1 01 H 1 • -co . o^ f- o 0 CO 05 05 ' PL, 2: k^ <^ o rH (_f U-' UJ H 01 o"j 00 •* CO CO LU CO 01 0 Q Q § h-l i— t ce W £ 1:3 O o cc £j 0^3 O O co CO CO QC H Q_ 2 C_> ^^ . --^ 05 1— Q IO O §_ CO CO o DI a. oo \.co C.) S Q UJ 5' r-1 r-1 C - ci CO CO CO C/J OQ iO Oi 'IO a: W O co o 0 H O t-H 3: 'H W UJ i \ _l h- t? UJ cr 05 CO 1 <£ ID- Q O IO 10 05 UJ i — i 01 00 io r-l O CO " 01 H-1 " rl Ol 6 ? 1 — 1 CO CO n CO CO CO r-i i— i 1 Ol <5 O o H CO co CO H \ v Q 05 V-H ce 10 U3 00 10 • in 'f •r-l co r-i H 0 H H ob rH ••3 - s 'oi 00 CO co r-( CO CO CO BS lO rH «^ W H U EH a CO < 1/5 en 0 pH 05 N 05 CO W — ' *-tJ W W P fa-, LU •?: CO CO CO -> 0 z Crt o W z"- — ~ kH W W. -a; o LU H | ol W 05 O < o co E .H o o CH 05 H E CO W CO Q- UJ co 33 CO M Q H W >—i 05 2 3 CO CO W H S 3 a. UJ CO Q 0 0 CAPTUR i ESCAP W fc, w z 1— n 33 C£ § o co CO CO . co W LU _i LU CO 05 PU H LU ™"D H-1 «-t^ 1 «T *? 3 <3 & CO •<t P ^? W >~j *-. & o: •^ * * ^1 W CO CO *C^O I— O IV Examples 1 through 4 are distinguished by the use of Titan class launch vehicles, a mission duration of 2.5 to 3 years, and SEP being used for most mission phases. Examples 5 through 6 require Intermediate-20 class vehicles, have a shorter trip time of 1.5 to 2 years, and use SEP only for the return interplanetary transfer. It is possible to return a 10 kg sample using the Titan IIID/Centaur single launch mode provided that SEP is employed for both Mars capture and escape maneuvers (Example 1). The capture spiral time is 98 days; this is approximately the time lag between lander separation and the rendezvous/docking maneuver. The stay time of 34 days refers to the time spent in a 1000 km Mars orbit by the orbiter bus. Example 3 is similar except that a Titan IIID(7)/Centaur is required and the mission duration is 200 days shorter. A hybrid option (Example 2) also employs the 7-segment Titan/Centaur but uses a chemical retro for Mars capture. This would alleviate the problem of orbiter bus/lander communications and the time lag between lander separation and rendezvous. The SEP power requirement for the first three mission concepts is about 20 kw and the propulsion on-time is 60-707o of the mission duration. The dual-launch mode (Example 4) uses a small (4 kw) SEP stage only for the return transfer to Earth. This type of mission could be performed ballistically with two Titan IIID/Centaur vehicles; the flight time is only 100 days longer. The shorter mission examples (Examples 5 and 6) require a relatively high energy Earth-Mars transfer. SEP is not recommended for this phase of the mission since the power require- ments are prohibitively high for large Earth departure mass. Even when SEP is used only for the return transfer the power requirement is at least 19 kw. Example 6 is a 600-day mission which will return a 25 kg sample. This mission uses a Venus I IJL-R ES.EARC.H. I.N STJJ-UXE swingby with the SEP system operating for- only 157 days on the Mars-Venus leg. The required launch vehicle is the Inter- • mediate-20/Centaur; the margin of launch vehicle capability is about 4000 kg. In conclusion, the study has shown that solar electric propulsion can be used effectively to accomplish the MSSR mission. Performance advantages over all-ballistic (chemical propulsion) systems are either a smaller launch vehicle requirement for comparable trip time and sample size, or a significant reduction in trip time for comparable launch vehicles and sample size. The latter advantage is not generally available when a Venus swingby opportunity is employed. I IT RESEARCH INSTITUTE vi TABLE OF CONTENTS SUMMARY . iii 1. INTRODUCTION _ . 1 1.1 Study Background 1 1.2 Study Objectives and Approach 1 1.3 Mission Phase Options 3 2. SYSTEM SCALING ASSUMPTIONS 11 2.1 Launch Vehicle Data 11 2.2 Stage Mass Data ' 11 2.3 Mission Velocity Data 15 3. SOLAR ELECTRIC TRAJECTORY REQUIREMENTS 21 3.1 Interplanetary Transfer , 21 3.2 Mars Spiral Capture and Escape 23 4. MISSION PERFORMANCE CHARACTERISTICS 31 5. BASELINE MISSION EXAMPLES 53 REFERENCES 66 I IT RES.EARC.H INSTITUTE. vii LIST OF FIGURES Figure Page 1 Option Array Set 4 1A Option Selection Example 8 2 Launch Vehicle Performance Curves 12 3 Mass Scaling for Lander/Ascent Probe 16 4 Mars Sample Capsule and Ascent Vehicle Payload 17 5 Solar Electric Trajectory Requirements for Mars Sample Return Mission, 1981-82 Opportunity 22 6 Net Mass Fraction Versus Solar Array Power 24 7 Mars Low-Thrust Capture Spiral Requirements (Spiral Time) 26 8 Mars Low-Thrust Capture Spiral Requirements (Mass Ratio) 27 9 Mars Low-Thrust Escape Spiral Requirements (Spiral Time) . 28 10 Mars Low^Thrust Escape Spiral Requirements (Mass Ratio) 29 11 Trajectory Profile for 1155d MSSR Mission 33 12 Single Launch Solar Electric Performance for 1155-Day MSSR Mission, SEP not Staged 35 13 Trajectory Profile for 1055d MSSR Mission 36 14 Single Launch Solar Electric Performance for 1055-Day MSSR Mission, SEP not Staged (Space Storable Retro) 37 NT RESEA.RCH INSTITUTE VIII Figure Page 15 Single Launch Solar Electric Performance for 1055-Day MSSR Mission, SEP not Staged (Solid Retro) 38 16 Trajectory Profile for 950d MSSR Mission 39 17 Single Launch Sdlar Electric Performance for 950-Day.MSSR Mission, SEP not .staged 40 18 Single Launch Solar Electric Performance for 950-Day MSSR Mission with SEP Staging 42 19 Trajectory Profile for 960d MSSR Mission 43 20 Trajectory Profile for 860d MSSR Mission 44 21 Trajectory Profile for 680d MSSR Mission 45 22 Dual Launch Solar Electric Performance for MSSR Missions 46 23 Single Launch Solar Electric Performance for 680-Day MSSR Mission 48 24 Single Launch/Two Probe Solar Electric Performance for 680-Day MSSR Mission 49 25 Trajectory Profile for 600d MSSR Mission With Inbound Venus Swingby 50 26 Single Launch Solar Electric Performance for 600-Day MSSR Mission With Inbound Venus Swingby 52 27 All-Ballistic Performance for 1040-Day MSSR Mission 62 28 All-Ballistic Performance for 625-Day MSSR Mission With Inbound Venus Swingby 63 .I-I-T .RESEARCH, i N.SJJJJJIE ix LIST OF TABLES Page S-l MSSR Mission Selection Summary iv 1 System Mass Scaling Relationships 13 2 Mission Velocity Data 18 3 Mission Phase Option Selection Guide 32 4 Baseline Mission 1 - Summary 54 5 Baseline Mission 2 - Summary 55 6 Baseline Mission 3 - Summary 56 7 Baseline Mission 4 - Summary 58 8 ' Baseline Mission 5 - Summary 59 9 Baseline Mission 6 - Summary 60 10 MSSR Mission Selection Summary 61 NT RESEARCH INSTITUTE X IIT RESEARCH INSTITUTE xi MARS SURFACE SAMPLE RETURN MISSIONS VIA SOLAR ELECTRIC PROPULSION 1.
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