"

.. ,"

NASA CONTRACTOR ~..NIA REPORT

STUDY OF TECHNOLOGY REQUIREMENTS FORATMOSPHERE BRAKING TO ORBIT ABOUT MARS AND VENUS

Volume I - Summary

Prepared by NORTH AMERICANROCKWELL CORPORATION Downey, Calif. for Ames Research Center

NATIONALAERONAUTICS AND SPACE ADMINISTRATION WASHINGTON,D. C. SEPTEMBER 1968 w NASA CR-1131

/ /-- /J STUDY OF TECHNOLOGY REQUIREMENTS FOR y a.e ATMOSPHERE BRAKING TO ORBIT BOUT MARS AND VENUS 14 b ' -.

Distribution of this report is provided in the interest of informationexchange. Responsibility for the contents resides in the author or organization that prepared it.

/ SD 67-994-1" 9

,/Prepared under Contract No. NAS 2-4135 by NORTH AMERICAN ROCKWELL CORP- Downey , Calif . for Ames Research Center NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

- " - . - ~ ". .. .. For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield,Virginia 22151 - CFSTI price $3.00

FOREWORD

This volume presents a condensed summary of the study. It is being submitted in accordance with Paragraph 2. 6 of Specification A-12241 contained in Contract NAS2-4135, Study of Technology Require- ments for Atmosphere Braking to Orbit About Mars and Venus, which was issued by the Mission Analysis Division, National Aeronautics and Space Adminis- tration. The data were generated between January and October 1967 and are presented in three volumes:

Volume I - Summary (SD67-994- 1)

Volume I1 - Technical Analyses (SD 67-994-2)

Volume I11 - Appendices (SD 67-994-3) Volume IV - Final Briefing (SD 67-994-4)

CONTENTS

Section Page

1.0 INTRODUCTION . 1

2. 0 STUDYGUIDELINES AND SCOPE . 4

3. 0 AEROBRAKING MISSION-SYSTEM DESCRIPTION 6

4. 0 SIGNIFICANTRESULTS . 9 4. 1 AerobrakingVehicle Requirements . 9 4. 2 ConfigurationAnalysis . 12 4. 3 RecommendedConfigurations . 16 4. 4 SensitivityAnalyses 17 4. 5 ModularApproach to Vehicle Synthesis 20 4. 6 Testand Qualification Considerations . 21

5. 0 RELATIONSHIP TO OTHER NASA PROGRAMS . 21

6. 0 SUGGESTIONS FORFUTURE WORK . 22 6. 1 Planetary-CaptureMission-System Analyses 22 6. 2 Atmosphere-Braking-Vehicle TechnologyStudies 23

-v-

ILLUSTRATIONS

Figure Page

1 MarsAerobraking Maneuver . 1 2 Initial Mass in Earth Orbit - Mars Missions 2 3 Initial Mass in Earth Orbit - Venus Missions 2 4 Study Logic Diagram . 3 5 Flared-Cone Concept . 4 6 Mars and Venus Parking-Orbit Characteristics 6 7 MarsAerobraker Miss'ion Profile . 7 8 Spacecraft Configurations . 7 9 Effect of Orbit Eccentricity on Vehicle Weight 10 10 Lift-to-DragRatio Requirements . 10 11 Effect of m/C+ and Analytical Model on Mars AerobrakerStagnation-Point Heating . 11 12 Effect of Entry Velocity on Heatshield Weight 12 13 Heatshield Weight Breakdowns . 17 14 Modular Approach to Aerobraker Synthesis . 20

TABLES

Table Page

1 Entry Parameters . 4 M ission Trip Times Trip2 Mission . 5 M ars Orbiter Mission VehiclesMission Orbiter 3 Mars 13 M ars Lander Mission Vehicles 13 Vehicles Mission Lander4 Mars V enus Orbiter Mission VehiclesMission 5 Orbiter Venus . 14 M odu le Shape Variations Shape6 Module . 14 7 AerobrakerSpacecraft Weight Sensitivities . 18

- vii -

" . . .. . ~ . 1.0 INTRODUCTION

Manned missions to Mars and rational selection at the earliest Venus, although not a stated national date. goal, are a logical extension of the currentspace program. Overall Previous studies have shown that technical feasibility of such mis- atmosphere braking to orbit about sions appears within the current and Mars and Venus (shown schemat- projected near-term state of the art. ically in Figure 1) offers the poten- However, the eventual accomplish- tial of significantly lower gross ment is constrained by certain key system weights in Earth orbit as issue? which must be resolved in the compared to retropropulsive cap- earlyplanning phases. One key ture. Weightcomparisons for decision which must be made early aerobrakers using space-storable in the program involves the choice propellants for planet-orbit depar- of the mode to effect planet capture ture and all nuclear retrobrakers (i.e. , retrobraking or aerobraking). are shown in Figures 2 and 3 for Each mode obviously must be Mars and Venus mission opportuni- studied intensively to develop the ties from 1980 to 2000. Both orbiter requisite background data for a and lander missions which are shown

1000 KM ALTITUDE

I APO PLANET OF SPACECRAFT SPACECRAFTRETRACTS ALL EXTERNAL EQUIPMENT ll AND ASSUMES ATTITUDE SPIRAL MANEUVER FOR ENTRY INTO MARS -*.. V. .., -._ / SPACECRAFTFIRES /Q."" , D- .-. - - - -

SPACECRAFT HALTS SPIN

Figure 1. MarsAerobraking Maneuver

1 NUCLEAR EARTHDEPARTURE STAGES ASSUMED 800 km (430 n mi) CIRCULAR PARKING ORBIT AEROBRAKERS AT MARS (SPACESTORABLE PROPELLANTS)

1 IO3 Ib n 24w Z r

1980I9821984 1986 1988 19W 19931995 IW9 1980I982 1984 1986 1988 1990 1W3 1995 1999

LAUNCH YEAR

Figure 2. Initial Mass in Earth Orbit-Mars Missions

IO3 kg IO3 Ib 0NUCLEAR RETROBRAKERS Mars, while the Venus data were AEROBRAKERS (SPACE-STORABLE obtained from the recently completed I200 PROPELLANTS) -24M) 45.400 kg P/L Commonalitystudy(2). The three (100.000 Ib) opportunities for Venus represent

8 the minimum, nominal, and maxi- 2 BO 3 mum energy requirements for Venus missions'in the 1980 to 2000 period.

m The use of elliptical parking orbits at Mars and Venus result in 0 - significant weight reductions for both 1988 198890 91 90 91 aerobrakersand retrobrakers. For LAUNCH YEAR example, results obtained for Venus Figure 3. InitialMass in Earth orbiter missions showed that the Orbit-VenusMissions nuclear retrobraker wasapproxi- mately 50 percent heavier for circu- weresynthesized using AV require- lar orbitsand 20 percentheavier for mentsdetermined by Deerwester(1)elliptical orbits than the aerobraker forVenus-swing-by missions for utilizing space -storable propellants

(lberwester, J. M. and S.M. D'Haem."Systematic (2)Codik, A. and R. D. Meston,"Final Report - Comparison of Venus Swingby Mode With Standard Technological Requirements Commonto Manned

Mode of Mars Round Trips. " Journ. of Spacecraft Planetary Missions. " North American Rockwell andRockets. Vol. 4 No. 7(July 1967). pp. 904-912. Corp., Space Division, SD 67-621 (Dec. 1967).

2 for all possible mission years. The the advantages and disadvantages of aerobraking maneuver is potentially this mode could be evaluated. more complex than the retrobraking Included were parametric and con- maneuver, however, and the space- ceptual design studies to define the craft designs contemplated are more spacecraft weights and their sensi" than an order of magnitude larger in tivity to variations in the environ- mass and volume than the largest mental models assumed, crew size, entry vehicles yet considered. mission profile, propellants , etc. , Potential technology problems in the as well as internal packaging fields of gasdynamics ,-thermodynam- arrangements. A secondaryobjec- c_ ics, structuresand materials, and tive was to analyze the requirements guidance and control could possibly for system simulation, testing, and negate some of the indicated weight qualification. savings; additionally, no serious consideration has previously been The study was conductedin accord given to the problems of system test with the logic diagram shown in Fig- and qualification. ure 4. Initially,the aerobraking vehicle requirements, including The primary objective of this entry corridors, heating, loads, and study was to conduct a detailed packaging, were determined para- investigation of integrated aerobrak- metrically for a wide range of mis- ing spacecraft and the associated sion and vehicle parameters. With technology implications so that both these requirements established, a

-ITUDIOBJECllVES

.CONFlGUlUllON

.DtMLOPMlNI REQUIREMENTS

DAlA R~QUIREMfNlS

lEQUlREMENlS

.SCALING UWS .GlOUND ltS1S EUlH AIMOSWEI 6 CIS-LUNAR SPACE

PROfElUNlS I

RfIRO 6 AtR0

Figure 4. Study Logic Diagram

3

I. .. configuration analysis was initiated. selected, and the technology impli- Concurrently, sensitivity analyses cations associated with each were were conducted to establish the delineated.In addition, preliminary effects of variations in mission consideration was given to test and parameters, propellant selection, qualificationrequirements. The and module selection on the vehicle pertinent results are summarized in designs.Over 30 detaileddesigns this volume; more detailed analyses were generated and evaluated on are presented in the volumes comparative performance, size, entitled, Technical Analyses and weight, and technology requirements. Appendices, SD 67-994-2and In the final phase, the four most 67-994-3. attractive configurations were

2.0 STUDY GUIDELINES AND SCOPE

The candidate configurations were Entry trajectories were deter- derived from a family of symetri- mined'for the entry velocities and cal, biconic shapes trimmed to atmosphere models specified in operate at lift-to-drag ratios of from Table 1. Allowableentry corridors 0.5 to 1.0. In addition, a blunt, were determined assuming a 10-km Apollo-type configuration was ana- (32,810-ft) minimum-altitude con- lyzed to determine its applicability straint at Mars and 5- and 10-g tothe aerobraking mission. The undershoot trajectories at Venus. basic biconic configuration is shown An approach corridor guidance- in Figure 5. accuracy capabilityof 20 km ( 10 n.mi.)

Table 1. EntryParameters

I Planet I Entry Velocity 1 AtmosphereModels

(32, 810 ft/sec) VM-8, MSC-3( 4)

12 km/sec MSFCLDM, I Venus I (39, 370 ft/sec) MDM, UDM (4)

(3)Martin,C. D., PhysicalCharacteristics and Atmospheric Data for Mars, NAA/SID 65- 1684 (Nov. 1965). (4)Venus and Mars Nominal Natural Environment for WsionStudies. Figure 5. Flared-ConeConcept NASA SP-3016 (1967).

4 .. .

was assumed. Flow fields and con- NASA/MSC(6); the Earth-return vective and radiative heat-transfer module, and other major subsystems rates were determined; nonadiabatic were obtained from reviously com- and self-absorption effects were pleted NASA studiesr7), (8). In considered. Thermal protection addition to the basic modules and requirements were established using subsystems required forthe mission, Avcoat 5026 (Apollo material) as the a 9100 kg (20,000 lb) probe comple- reference ablator. ment was included in each design. The trip times, which size the con- The entry corridor and heating sumable~and meteoroid protection data were then employed to identify and establish subsystem require- the allowable range of values for the ments, are shown in Table 2. ballistic parameter (m/CDA) and Internal propulsion stages employed L/D for each atmosphere. Packaging for planet departure were sized for studies were conducted concurrently the orbits and hyperbolic excess using these m/C+ and L/D require- velocities shown in Figure 6 and ments to establish initial geometric were equal to or exceeded the relationships and weight distribution requirements for 80 percent of the requirements for a 10 m (33 ft) base swing-by mission opportunities in diameter configuration (to be com- the metonic cycle. patible with the current Saturn V). A multiloop iteration then was per- Table 2. Mission Trip Times formed to establish configurations (Days) which were acceptable from both aerodynamic and packaging consid- I Planet I Outbound I Stay I Homebound 1 erations.Structural analyses, including provisions for radiation 160 30 240 and meteoroid protection, were con- 1 y::ts 1 80 30 210 ducted in support of the packaging studies . Propellant combinations consid- ered include space storables (OF2/ Mission ground rules specified MMH and FLOX/CH4), cryogenics the use of an eight-man crew for (LO2/LH2 and LFz/LH2), and a both Mars and Venus missions, nuclear (HZ) system. The respective having four men available at Mars to land an excursion module. The (6)Definition of Experimental Tests for a Manned mission module considered for the MarsExcursion Module. Contract NAS9-6464 interplanetary journey was basedon (Nov. 1966). the designs developed for the manned ("Technological Requirements Common to Manned Marsfly-by mission(5). The Mars Planetary Missions, Second Interim Report excursion module was obtained from (Contract NAS2-3918) NAA SMD, SID 67-294-1 a concurrent study conducted for (10 Mar. 1967). @)Study-~ of Manned Vehicles for Entering the Earth's ("Manned Planetary Flyby Missions Based on Saturn/ Atmosphere at Hyperbolic Speeds, Final Report ApolloSystems. (Contract NAS8-18025) NAA NAS2-2526,Lockheed, LMSC 4-05-65-12 S&ID, SID 67-549 (1 Aug. 1967). (Nov. 1965).

5 performance capabilities and pack- km/uc 103ft/sec aging problems associated with both 5.0 -- 16.0 bell- and plug-nozzle engines were B tankage the as Inasmuchconsidered. 4.0 - requirements for hydrogen-propelled s! -c configurationsmaysystem nuclear 3.0 - propellants otherexcessive,become c Yz than hydrogen (e.g., LiH, NH3, etc.) 5 n. 2.0 - exploredwere to determine if any VENUS, V, 5.5 km/scc (18, 046 ft/scc) advantage accrues from using - 4.0 1.0 - higher -density propellants to reduce tankage requirements (and sizes).

I I 0 0.3 0.6 0.9 ECCENTRICITY

VENUS

ORBITAL PERIOD hours

PERIAPSIS ALTITUDE, 3M) km

Figure 6. Marsand Venus Parking- Orbit Characteristics

3.0 AEROBRAKINGMISSION-SYSTEM DESCRIPTION

A typical mission profile was logistics spacecraft which, presum- developed so that the various opera- ably, has been developed in support tions and systems involved in a Mars of other Earth orbital programs landing mission could be viewed in (e. g. , space stations). After the perspective.The mission is shown crew is transferred to the aero- schematically in E'igure 7, and an braker, injection stages are mated inboard profile of the aerobraker to the vehicle, and rendezvous with vehicle is shown in Figure 8. An the tankers is effected. outbound Venus swing-by mode was chosen, but the profile developed The fully fueled and checked-out could apply (with minor changes) to aerobraker is shown to be injected either Mars or Venus orbiting or into a transfer orbit to Mars by landing missions . nuclear modules. (Two modules may be used for the first-stage injection, The fully assembled aerobraking a third module being used after the vehicle is launched unmanned to an first-stage modules are jettisoned. ) assemblyorbit. The crew is deliv- If an abort situation should occur, ered to the assembly orbit in a the Mars-orbit-escape propulsion

6 0Launch to @ Earth Orbit @) Earth Earth Orbit OperationsOrbitEarth Departure

Figure 7. MarsAerobraker Mission Profile

Figure €4. SpacecraftConfigurations system is used to place the space- density condition (which is a function craft on an intercept trajectory with of the chosen parking orbit charac- Earth. When Earth is approached, teristics) is reached for initiation of the crew enters the Earth reentry the exit maneuver. The maneuver is module (ERM) and effects a normal initiated by redu.cing the roll angle entry and recovery. so that the vehicle ascends through the atmosphere.The atmospheric After injection, if a requirement exit trajectory is guided and con- for artificial gravity is established, trolled in the same manner as entry the mission module (MM) and MEM to achieve the desired exit velocity may be extended on cables after the and path angle. course correction and the resulting configuration spun up to achieve about After the spacecraft leaves the 1/3 g. Theelectrical power system atmosphere, it coasts up towards the radiator s, communication antenna, apoapsis altitude previously deter- and instruments (e. g., telescope) mined. The heat shield panels are then are extended and adjusted. jettisoned during coast. A powered maneuver at apoapsis delivers the As the spacecraft approaches spacecraft to its orbit, and the orbit Venus along the outbound swingby parameters are then determined. trajectory, the Venus probes are Observations of the surface can be checked out and injected into their made to determine potential MEM propertransfer trajectories. Upon landingsites. The spacecraft crew approaching Mars, the final course performs scientific observations in correction is made, the spacecraft's orbit, including the injection of a attitude is adjusted for entry, and all probe complement, after the MEM external appendages are retracted and its crew descends to the surface. and stowed or jettisoned. The space- When the planned surface staytime craft encounters the atmosphere at has been completed, the MEM an altitude of about 200 km (656 x ascends, rendezvouses with the 103 feet), with a fixed attitude, and spacecraft, and the crew transfers at an entry angle (with respect to the tothe MM. Finalcheckout is initi- local horizontal) near the middle of ated, and the procedure culminates the allowable corridor. As the aero- with ignition of the planet-orbit- dynamic forces increase, decelera- departurestage. After burnout, the tion rates are measured and time trajectory parameters are deter- integrated to sense and generate mined in conjunction with the DSIF, real-time data on the entry flight and the time and attitude for the first path.The trajectory is matched to midcoursecorrection are estab- a preprogrammed reference trajec- lished. Upon approachingthe Earth, tory, and deviations are corrected the ERM is separated from the MM by modulating roll. and propulsion units and is properly oriented for entry into the Earth's After pullup, the r 011 angle is atmosphere. Entry and recovery are adjusted to maintain constant-altitude similar to current Apollo lunar mis- flight until t'he proper velocity- sion procedures.

a 4.0 SIGNIFICANT RESULTS

The study results have shown that The MEM rendezvous propulsion the selected aerobraker configura- requirements prompted an examina- tions exhibit satisfactory perform- tion of an additional mission mode, ance and packaging characteristics. i. e. , a staged-aerobraker lander There is considerable commonality which employs an aerobraking between the spacecraft systems excursion module to establish an required to accomplish the Mars and intermediate circular orbit (at lower Venus missions if the parking orbit altitudes) prior to deorbiting the eccentricity and probe payload lander. The propulsion stage in the parameters are adjusted appropri- intermediate orbit later effects ately. This finding resulted in the rendezvous with the parent spacecraft development of a modular approach in the elliptical parking orbit. A to spacecraft design for Mars and/or laboratory attached to the propulsion Venus missions.These areas and stage could gather low altitude data other pertinent results are discussed and photographs, Use of this mode, in the paragraphs which follow. along with highly elliptical parking orbits (i. e. , e > 0. 6), results in 4.1 AEROBRAKING VEHICLE lowertotal system weights. For REQUIREMENTS eccentricities lower than about 0. 6, however, a weight penalty is incurred. Mission Modes-As can be seen Detailed comparisons of possibly in Figure 6, departure AV require- enhanced operational flexibility versus ments for orbits with an eccentricity the increase in system complexity as of 0.6 are approximately 25 and 50 it effects crew safety and mission percent lower than those required for success were beyond the scope of this departure from Mars and Venus study. circular orbits with corresponding

"periapsis altitudes. I' The lower Entry Performance-Performance energy requirements can be translatedstudies defined the aerobraking to reductions in aerobraker entry entry corridors for Mars and Venus weights of 40 percent at Mars and 55 as a function of L/D, m/CDA, percent at Venus for orbital missions.atmosphere model, and undershoot Introduction of a manned excursion criteria (i.e. , minimum altitude or module for a landing at Mars reduces maximum deceleration).Assump- the effect of the elliptical orbit inas- tions included an approach navigation much as the excursion module ascent corridor capability of 20-km propellant requirements increase (-10 n. mi.), a 10-km (33,000-ft) withorbital altitude and/or minimum pullup altitude at Mars , a eccentricity. 10-g undershoot limit at Venus, and

9 a 10-m (33-ft) spacecraft base diam- eter. The L/D requirement at Mars increases monotonically with the 1.6 - ballistic parameter (m/CDA) for the several atmospheres investigated. 1.4 - Figure 9 shows this variation for the "worst case" atmosphere (VM-8) for 1.2 - a range of velocity while Figure 10 - shows typical values of L/D as a 1.0 function of CDA and angle-of -attack m 0.8 - for a given set of vehicle geometric parameters. It should be .noted a 0.6 - given L/D ratio can be achieved by flying at both low and high angles of ANGLE G ATTACK 0.4 - attack. These data essentially define IIY- the biconic configuration require- 0.2 - ments, i. e. , packing density and

geometry. For example , inasmuch 0- . as preliminary estimates of Mars 0 vehicle weights on the order of 150,000 to 200,000 kg (330,000 to Figure 10. Lift-to-DragRatio 440,000 lb) were indicated, the need Requirements for a high angle of attack is immedi- ately apparent to satisfythe m/CDA - In the dense Venus atmospheres, L/D relationship indicated by Fig- the effect of m/C+ on required L/D ure 9 for the low-density VM-8 is eliminated inasmuch as m/C+ atmosphere. only affects the position of the \. corridor (i.e., height) in the atrnos- -\ phere. ,ln the future, an altitude limit may be placed on Venus entries 1.2- MODEL ATMOSPHERE VM-8 to assure pullup prior to cloud pene- IO-h MINIMUM ALTITUDE 20-h CORRIDa DEPTH tration-Such a limit wasnot 1 .o- considered at this time because of the lack of suitable estimates of cloud-layerheights. Inasmuch as m/C& is not a problem, the Venus- 0.6 VE - 10 h/mc entry vehicles can be operated at low angles of attack. 0.4 0.5 1.o 1.5 2.0 10' lb/it2 0 Gasdynamic Heating - The 02 4 8 10 IO3 te/m2 atmospheric density profile and com- BALLISTIC PARAMETER position, entry conditions and corridor boundary, and vehicle characteris tics Figure 9. Effect of Orbit (e. g., L/D,m/CDA, geometry) all Eccentricity on significantly affect aerobraker gas- VehicleWeight dynamic heating rates and loads.

10 Heating-rate differences induced by entryconditions and geometry. the density variation across the Ablator thicknesses were obtained assumed 20-km entry corridor were at different points on the vehicle, greater than those due to uncertain- and thickness distributions were ties in the atmospheric densityprofile developed.The ablator weight orcomposition. Consideration of contributes about 2/3 of the total nonadiabatic and self-absorption heatshield weight; the other 1/3 is phenomena in determining radiation support structure and insulation. heating further reduces the effect of The total heatshield weight fraction variations in atmosphere composition. as a function of entry velocity is Figure 11 shows the effect of m/CDA shown in Figure 12 for both Mars and analytical model on radiative andVenus vehicles. The bands heating which predominates at the shown represent a range of ballistic stagnation point. Nose bluntness and parameters and show that the heat- forebody cone-angle are the dominant shield weight fractions range from geometric parameters with regards 5 to 20 percent of the aerobraker to heating at other locations; the entry weight. If these ratios are afterbody half-cone angle has a interpreted as mass fractions, an negligible effect. equivalent heatshieldIsp for Mars is approximately 7000 seconds The heat loads developedwere used(AV-4 kilometers per second), while to determine heatshield requirements for Venus (AV-3 kilometers per for the many possible variations in second) it approaches -2000 seconds.

JPL VM-7 ATMOSPHERE JPLVM-7 ATMOSPHERE vb 1.0 I VE - lOkm/YC : 1i0WYC I @OPTICALLY THIN, ADIABATIC @OPTICALLY THIN, ADIABATIC

@OPTICALLY THIN, NONADIABATIC N @ OPITCALLY THIN, NON ADIABATIC N5 J @SELF ABSORBED, ADIABATIC d @SELF ABSORBED, ADIABATIC 2 ';. @SEI€ ABSORBED, NONADIABATIC 3 3 @SELFABSORBED. NONADIABATIC Irr, ta I2 - 10

4-

Figure 11. Effect of m/CDA and Analytical Model on Mars Aerobraker Stagnation-Point Heating

11 based on performance, storability, and engine cooling characteristics; 0.20 this combination was selected for the final designs. 0.16 k In addition to the propulsion z 5 0. I2 analyses, structural concepts for the

j? thermal-structuralsystem, load-

0.08 carrying structure, and tank and module supports also were investi- 0.04 1 gated. A skin-stringerconcept was i 10 12 14 I6 selected to accommodate the high ENTRY VELOCITY anticipated in-plane loads; the skin- Figure 12. Effect of EntryVelocity stringer unit weightsof 15 to 20kg/m2 on Heatshield Weight (3 to 4 lb/ft2) were approximately 10 percent lighter than equivalent 4.2 CONFIGURATION ANALYSIS honeycomb or truss-core concepts.

Mars aerobrakers for orbiter and A study of spinning-configuration lander missions and Venus designs dynamics was conducted for typical for orbiter missions (i. e., no aerobraker designs, in the event that manned excursion module) were artificialgravity is required.The developed for space-storable, results indicated no significant prob- cryogenic, and nuclear propulsion lems with this mode and that the planet-orbit-departure systems for propellant requirements for spin and the several parking orbitsof interest. despin are on the order of 3 to 5 per- Indicated lengths for the nuclear cent of the vehicle weight. (All of vehicles were in excess of 40 m the designs considered are applicable (132 ft) if the base diameter is con- to either the spinning or nonspinning strained to 10 m (33 ft); furthermore, mode. ) the resulting configurations did not provide satisfactory aerodynamic Matrices of thedesigns consid- characteristics.Difficulty was ered are presented in Tables 3 encountered in achieving reasonable through 6. Theconfigurations shown stability margins with cryogenic were chosen to compare the effects propellants (i. e. , LO2 /LHz). of circular and eccentric parking orbits; storable (OF2/MMH or Of the propellant combinations FLOX/CHq), cryogenic (LO2 /LHz), considered,the space-storables and nuclear-LH2 propulsion systems; (e. g., OF2/MMH or FLOX/CH4) and module selection on the aero- showthe most promise for braker packaging arrangement. aerobrakers. The Mars orbiters shownin Further analyses indicated that Table 3 are grouped by propellant FLOX/CH4 is the most desirable type,the heaviest vehicles being

12 Table 3. Mars Orbiter Mission Vehicles

MM: 8-man. 430 -day

Periapsis altitude 300 km

Base diameter 10 m (33 ft)

Eccentricity 0 0.6 0.9 0.6 0.9 0.6 Propellants S S S C C N Length m 19.7 18.8 18.8 23.4 21.5 29.6 ft 65 61 61 77 71 97 Entry weight lo3 kg 265 186 165 188 161 164 lo3 lb 584 410 364 41 3 353 362 Injected weight lo3 kg 279 195 173 197 168 173 lo3 lb 614 430 382 434 370 380

-~~ ~~ ~

Table 4. Mars Lander Mission Vehicles " ~ ~ ~~ . ~~ ~~ ERM:Apollo

MEM: direct, Apollo

MM: 8-man, 430-day

Periapsis altitude 300 km

Base diameter 10 m (33 ft), except as noted D =11.7 m (39 ft)

"~ ~ - I I Eccentricity "9.6 0.6 0.9 0.6 0.6 Propellants S C C C N Length m 1 26.6 26. 6 40.0 34.6 30.0 40.6 ft 87 87 127 113 98.5 138

247 244 227 242 220 543 536 500 533 484 Injected weight IO3 kg 260 256 239 254 231 lb 685 lo3 lb 571 563 526 560 509

13 Table 5. Venus Orbiter Mission Vehicles

ERM: Apollo

MM: 8-Ip.n, 380-d8y "I- Periaprir altitude 300 km

B8re di8me ter 10 rn (33 ft)

hylord weight IO3 kg 45.5 IO3 lb 100

Eccentricity 0.9

Propellantr S

Longth m 23.8 25.0 ft 78 61 1 18.8 82 Entry weight IO3 kg 276 135 174 lo3 lb 60 7 297 3 82

Injected weight IO3 kg 290 I42 185 lo3 lb -638 -312 40 8 Table 6. Module Shape Variations 1

Poriapmir altitude 300 km

B8re diameter 10 rn (33 ft) I Eccentricity 0.6 0.9 0.6 Propelhntr s S C Mimmion Mar. orbiter M8rr rbiter Mar. rbitor ERM Apollo(Conic reg Apollo Apollo Apollo Zonic re Apollo zonic ..I MEM - - - - Apollo Ltfting b-JY Langth -I- m 18.8 21.0 23.4 21.9 26.6 28.7 ft 61 69 77 72 87 94 Entry roiaht IO3 kg 165 165 187 I86 297 298 IO3 lb 364 364 413 410 652 652 Injoctod roiaht IO3 ka 174 174 197 196 310 310 IO3 lb 382 382 434 43 I 685 685

'I14 located to the left of the chart. It is of the propulsion systems considered: readilyapparent that, while' the i. e., storable (OF2/MMH, FLOX/ cryogenic and nuclear propellant CHq), cryogenic (L02/LH2), and aerobrakers are lighter for a given nuclear(LH2). The comparison also mission, they are also much longer. is made between the several propul- The three smallest orbiters use sion systems for a given orbit (e. g., storablepropellants for planet e = 0. 6). This group of landers is departure; all are approximately the arranged similarly to the orbiter same size, with length-to-diameter family, the heaviest vehicle shown ratios of about two. This allows a to the left. fairly large aft-cone angle, resulting in a center of pressure located aft of Mars vehicles using nuclear (LHz) thecone intersection. In order to propulsion are the lightest of the achievean acceptable center of total Mars lander spectrum: their gravity, the POE propellant is stored weights run about 10 percent less in the annulus around the three-floor than the cryogenic landers and 11 mission module. percent less than the storable landers for a givenorbit. The 10-m (33-ft) Based upon the investigations base diameter configuration sized conducted during the study, a recom- for the 0.6 e orbit approaches a mended Mars aerobraker orbiter, length-to-diameter ratio of about indicated by the shading in Table 3, five, because the LH2 located for- wasdeveloped. This configuration ward of the MM requires such a is powered by a single stage for large volume that only a small aft- planet departure, using storable coneangle is allowed.The MEM FLOX/CH4 propellants and an aero- and probes preclude the propellant spikeengine. The mission module from being in the aft end, with its chosen has two floors, curved bulk- larger volume, because of their own heads, and a diameter of 8.2 m size and position requirements. (27 ft). It incorporates an equipment These lengths and packaging require- bay on the forward end in which ments create difficult stability antennas,telescopes, solar panels, problems and preclude the use of and scientific instruments are stored, nuclear (LH2) propulsion for aero- Theaft end of the MM provides braker vehicles. storage for life-support gases, outbound-course-correction propellant An integrated lander design, tanks, and a scientific laboratory indicated by the shading in Table 4, area.The probe compartment is wasdeveloped. This configuration located at the aft end of the vehicle. achieves a proper balance between cg and cp and makes use of FLOX/ Aerobraker vehicles for the Mars CH4 propellants in an aerospike landing missions are shown in engine.The FLOX/CH4 combination Table 4. Theconfigurations afford appears to present less long-term a comparison of the effects of park- storage problems than the other l I ingorbit on vehicledesign for each storables considered (OFz/MMH). I

15 For the Venus orbiter family No appreciable change in spacecraft shown in Table 5, aerobrakers gross weight is noted for these shape carrying an Apollo-shaped ERM and variations. either a light or a heavy probe com- plement were developed to compare Figure 12 presents a summary of the effects of parking orbits and pro- the vehicle injected weight variation pulsionsystems. All the spacecraft with parking orbit eccentricity for shown are similar in arrangement both Mars and Venus missions. Also and carry the single planet-departure shown are the weights for direct and propulsion stage in the nose, followed staged Mars excursion modules. by theERM, MM, andprobes. A recommended design, indicated by 4.3RECOMMENDED the shading in Table 5 was generated. CONFIGURATIONS For purposes of commonality, the exterior shape and basic arrange- Recommended aerobraker designs ment are identical to the equivalent include Mars and Venus orbiters and Mars aerobraker, even though the Mars lander, Venus orbiter vehicles Venus mission requires smaller sized for a planet departure AV of POE tanks. If the POE tanks were 3.6 km/sec (11,800 ft/sec), equiv- filled, an orbit with an eccentricity alent to an eccentricity of 0. 6 at as low as 0.2 could be achieved at Marsand 0.2 at Venus. These Venus. designs used FLOX/CH4 for the departurepropulsion and carry Table 6 illustrates a number of Apollo-shapedEarth-reentry Mars aerobrakers (both orbiters and modules.The Mars lander also landers) configured to compare the carries an Apollo-shaped excursion effect of the conic segment ERM and module.Schematic profiles of the thelifting body MEM. Also, two four recommended designs and Mars orbiter designs, sized for pertinent weight data are shown highly eccentric orbit (i. e. , e = 0.9) shaded in Tables 3 through 6. and using storable propellants, were developed to compare the effect of an Heatshield weight breakdowns for external probe compartment on the the four selected configurations are aerobraker length from base plane to shownin Figure 13. Theweights nose. ranged from 7 to 15 percent of the vehicle entry weight; the Venus Little or no difference was found orbiter required the highest heat- in the length or packaging arrange- shield design was considered forboth ments between the Apollo-shaped and Mars and Venus entry (assuming the conicsegment ERM, Storing all the spacecraft could be trimmed at the probes externally aft of the bulkhead proper angle of attack). The weight resulted in a 5-percent shorter Mars penalties imposedon the recommended orbiter.The lifting body MEM designs for a common heatshield can requires about the same length com- approach 50 percent of the heatshield partment as the Apollo-shaped MEM. weight.This penalty can be reduced

16 r

increase in the number of man-days.

The POE AV requirement is a function of the parking orbit selected and the departure V,. Changingthe parking orbit from 0.6 eccentricity to a 300-krn circular orbit increases the POE AV 20 percent, resulting in a 20- to 40-percent increase in Earth departure weight.

ABLATOR Performance - Entry corridor sensitivities to the environments, flight modes, mission parameters, and guidance concepts were estab - lished to define the approach naviga- HONEYCOM tion and maneuvering requirements. INSULATION SUPPORTS A vehicle with an L/D of 1.0 and an MARS MARSMARS VENUS LARGE ORBITER LANDER ORBITER PAYLOAD ORBITER m/CDA 5 8500 kg/m2 (1750 lb/ft2) could accomplish the aerobraking Figure 13. HeatshieldWeight orbital capture maneuver at Mars in Breakdowns VM-8, the most tenuous atmosphere postulated. At Venus,the maneuver by judicious selection of the mission can be accomplished in the atmos- and vehicle parameters (e. g., Ve, phere models considered with an L/D m/CDA, cy, corridordepth). of 0.5 to 1.0 and m/CDA 5 20,000 kg/m2 (4000 lb/ft2). 4.4 SENSITIVITYANALYSES Flight-mode performance was Sensitivity analyses were con- evaluated in terms of exit velocity ducted to determine the effects of and flight-path-angle accuracies changes in assumptions, mission required to achieve a prescribed parameters,crew size, propulsion parking orbit versus the applied AV requirements,and packaging correction. An almostone-to-one arrangements on the selected config- relationship was found between the urations, and the data are summarized deviation in exit velocities from the inTable 7. TheEarth-departure prescribed value and the applied AV weights are most sensitive to changes required.Flight modes considered in the planet-orbit escape (POE) AV included constant flight path angle requirement, and in the mission and constant altitude trajectories; (MM) weight. The MM weight is a the choice of mode produced a function of crew size, mission negligible effect on exit velocity duration, and type of life support accuracy.Required roll rates for system, For the’ subsystems both modes were estimated to be on considered, the “weight increases the order of 3 degrees/second. by 10 percent for a 25 percent

17 Table 7. Aerobraker Spacecraft Weight Sensitivities

Percent Change in Injected Weight t Mars IVenus Orbiter Module Change Lander lrbitel Large Payload Small Payload

A Mission module 8 men - 430 days *4. 1 f5. 1 8 men - 380 days f3.3 f4. 1 Earth reentry module 8 men - Apollo shape 110% fl. 1 fl. 4 fl. 0 *l. 3 n module Probes veight 9., oao ~g (20, ooo 1b) fO. 5 fO. 7 *o. 7 +o. 9I 36,320 Rg (80, 000 lb) *2.4 Tars excursion module 4 men - 30 days ?rl.9 Apollo shape - ~. . ~. -. .. - - c Propulsion Svstem rnlsec ftlsec htbound course 76 250 :207'0 AV fO. 4 0.4 fO: :orrection 152 500 I2070 av +o. a fo. a htbound spin 132 440 12 cycles:: fl. 0 fl. 0 md despin 66 220 11 cycle fO. 5 *o. 5 Jlanet orbit 152 500 120% av *o. a fo. a .ttainment and 76 250 :20qo av fO. 4 *o. 4 naintenance

'lanet orbit spin 66 220 :1 cycle *o. 2 *o. 2 fO. 3 fO: 4 nd despin 'lanet orbit 3, 600 1, aoo 110%AV t10.7 -13. a scape - 7. 6 9. a 2,700 a, 200 :lOqo AV t4.3 *5. a -3.4 -4. 7 .eturn spin and 66 220 11 cycle *o. 3 *o. 4 fO. 4 f0. 5 e spin

.eturn course 76 250 I2070 av !Lo. 3 *o. 4 *o. 3 f0. 3 2rrection

J ~~

18 Heating and Heat Protection - compromises in the mission and Convective and radiative heating vehicle par.ameters. rates for the selected vehicles were computed using methods represent- Structure andshielding -Integrated ative of the current state of the art. structure unit weights of 15 to 20 kg/ It was found that the choice of m2 (3to 4 lb/ft2) were found for the analytical model used to obtain the range of configurations studied. radiative heating (i. e. , adiabatic, These weights include the load- nonadiabatic with self-absorption, carrying structure, as well as etc.) was more significant than the meteoroid protection requirements variations in the atmospheric compo- to assure a 99-percent probability sition. At Mars, there was a of no penetration in 430 days, given decrease in maximum heating rate the nominal meteoroid flux model.(9) of 15 percent in going from VM-7 Theunit weights derived were tothe MSG-3 atmosphere. Changing more sensitive to the analytical the analytical model from the simple method applied than to the mission adiabatic case to the sophisticated assumptions. nonadiabatic,self-absorption approach results in a stagnation Design Integration - Sensitivities heating rate almost 50 percent lower, of theconfiguration packaging while peak heating at the other arrangements to changes in the locations is reduced by as much as orbital parameters and candidate a factor of five. A 20-percent propulsion system for Mars lander increase in entry velocity increased vehicles were illustrated earlier in the peak stagnation heating rate by Table 3 through 6. The Earth reentry a factor greater than two;increasing module (ERM) shapewas found the angle of attack from 8 degrees to have no effect on the vehicle length to 28 degrees increasedthe radiative orarrangement. An Apollo-shaped heating rate on the conical surface Mars excursion module (MEM) was by more than an order of magnitude. selected; use of the 16-m (35-ft) long lifting body MEM does not severely The effects of these uncertainties penalize the packaging arrangement. on the heat protection system weights The effect of 4- to 12-man crews on were reflected in heatshield weight the packaging arrangements is not variations of almost 100 percent, or clear because crew free-volume vehicle entry weight variations of requirementsare poorly defined. , from 7 to 15 percent. The heat- For example, a mission module shield design for the Mars mission designed for8 men witha free volume was 35 percent lighter than the Venus per man of 20 m3 (700 ft3) could be,: heatshield. If a common heatshield used for 12 men at 13 m3 (470 ft3)' ' design were used for both missions, free volume per man; recommended the weight-in-Earth-orbit penalty for the Mars mission would be 3 to 8 percent. This penalty can effec- (')Manned Planetary Flyby Missions Basid on Saturn/ tively be eliminated by suitable Apollo Systems, op. cit.

19 free volumes range from 10 to 20 m3 The vehicle is comprised of three perman, (10) major modules; i. e., a propulsion (or nose module), a mission module, 4.5 MODULAR APPROACHTO and a probe or probe-and-lander VEHICLE SYNTHESIS module.Initially, the mission and probe modules for each configuration In the latter phases of the config- were developed as a unit to satisfy uration analysis, it became apparent the cp-cg relationships, volumetric that a series of common modules requirements, and the base-diameter could be used to synthesize the constraints.After the configurations aerobraker.The approach, were analyzed, it appeared possible illustrated in Figure 14, is attractive to treat these modules as discrete because only four basic segments components in order to establish a are needed for all the orbiting and mission module shape common to both landing missions of interest. the orbiter and lander configurations.

Each spacecraft heatshield initially was treated as a unique design; if ('"Davenport, E. W., Congdon, S.P., and Pierce, 6. F. "TheMinimum Volumetric common elements were to be specified for the propulsion, mission, and Requirements of Man in Space, " AIAA Paper No. 63-250, presented at AIAA Summer Meeting probe modules, a common heatshield in Los Angeles, California (17-20 June 1963). capable of satisfying multimission

Figure 14. ModularApproach to Aerobraker Synthesis

20 I

and spacecraft requirements might found that the aerobraking maneuver be expected to suffer some weight is of the same orderof complexity as disadvantage. It was found that return and entry along a maximum although the Mars heatshield weights range trajectory on current Apollo increased 33 to 36 percent to achieve lunar missions where the Apollo commonality, the gross spacecraft exits the atmosphere after being weight increase would be less than slowed to near orbital velocity and 3 percent; the Venus heatshield eventually enters again to achieve weight would increase 2 to 6 percent the desired range. This entry profile and the spacecraft gross weight would is, in almost all respects, a valid increase less than 1 percent. These simulation of the aerobraking concept. increases could be reduced to negli- gible values by altering the mission- Additional aspectsof the vehicle's system parameters slightly (i. e., design and performance could be velocityand L/D). Although some tested in a combination of ground, penalty would be incurred in the earth-atmosphere, and cislunar injection stage, the obvious advantage spaceenvironments. No planetary- of a common heatshield appears to based tests would be required beyond make this a worthwhile trade-off. the currently planned unmanned data-gathering probes. If the mod- 4.6TEST AND QUALIFICATION ular approach to the aerobraker CONSIDERATIONS design is adopted, parts of the vehicle could be used as Earth- Acceptance of the aerobraker orbital space stations while being mode will require sufficient testing tested for the more demanding to establish a high level of confi- planetarymissions. Although this dence in the aerobraking mission. approach is most promising, it has The requirements for these tests not been studied in detail. were considered briefly. It was

5.0 RELATIONSHIPTO OTHER NASA PROGRAMS

The current study was performed The baseline aerobraker configura- for the Mission Analysis Division of tion used in the present investigation NASA's Office of Advanced Research was developed in 1964 under Contract and Technology and is the fifth in a NAS9 - 1748, "Study of Subsys terns related series of manned interplane- Required for a Manned Mars Mission tary mission requirements studies Modulell; a sensitivity analysis of the conducted by North American Rock- baselineaerobraker vehicle was well. The series started in 1964 with conductedin 1965 underContract Contract NAS2-1408, "Manned Mars NAS2 - 247 7, Study of Unmanned Landing and Return Mission Study. I' Systems to Evalute the Martian

21 Environment.The results of a con- More comprehensive data have current study of the Mars excursion been generated in this investigation module,conducted under Contract of the aerobraking mode to effect NASS - 6464. entitled "Definition of planet-orbitalcapture. The size- Experimental Tests for a Manned and weight-scaling relationships Mars Excursion Module,(I were used developed can readilybe extrapolated directlyin this study. Two other to assess effects of mission date, parallel studies which furnished val- duration,and flight mode. The uable data in some areas were Con- designs developed illustrate possible tract NAS8-18025, "Study of Manned packaging arrangements and volume Planetary Flyby Missions Based on utilization and will serve as the base-

Saturn/Apollo Systems, fI and Con- lines for the next series of manned tract NAS2-3918, "Study of Techno- planetary capture and landing logical Requirements Common to studies. Manned Planetary Missions.

6.0 SUGGESTIONS FOR FUTURE WORK

Two broad areas of future studies braking appears feasible in the cur- are recommended as a result of the rently postulated family of Mars .and workaccomplished. The first is Venus atmospheric models; a seg- described as "Planetary Capture mented modular vehicle concept Mission-System Analyses" and the which affords multiplanet, multimis- second as "Atmosphere Braking sion capability and has obvious eco- Vehicle Technology Studies. nomic attractions has been defined. Consequently, detailed comparative 6.1 PLANETARY-CAPTURE analyses of aerobraking and retro- MISSION-SYSTEM ANALYSES braking should now be performed for a wide range of mission opportuni- The current study provides data ties. Theanalyses should consider which indicates that the aerobraking developmental problems, cost, mode has a potentially significant reliability, and schedule risk, as advantage over the retrobraking well as the system requirements. mode when the two are compared on the basis of weight in Earth orbit. The availability of an acceptable Furthermore, the technology devel- test and qualification plan for an opment requirements in propulsion aerobraker spacecraft system, pref- (i.e., nuclear systems) are more erably one which could be conducted demanding for retrobraking than for in the near-Earth environment, aerobraking, inasmuch as the latter would go far in promoting acceptance can be accomplished with a chemical of this mode. This program should planet-0rbit;departurestage. Aero- encompass both scaled models and

22 the modularized segments to man characteristics of the modular rate the total system; detailed sizing, spacecraft.Evaluate candidate tri- scaling, and simulation would be a conic configurations to assure ade- prerequisiteanalysis. The modular quate stability and lift capability. approach to vehicle synthesis sug- Study entry dynamics in sufficient - gested by thisinvestigation should de tail to derive the guidance and con- be extended to both Earth-orbital and trol system requirements, including planetary-flyby missions, and the the reaction control system. Parking associated compromises (and penal- orbit characteristics for selected ties) in the mission and spacecraft mission opportunities must be requirements and capabilities should examined in more detail so that orbit be identified. precession is taken into account as well as the proper orientation of the 6.2 ATMOSPHERE-BRAKING- approach and departure V, vectors. VEHICLETECHNOLOGY STUDIES Heating and Heat Protection - Refine the computation of the radia- Suggested technology studies of tive heating environment to include the atmosphere-braking vehicle are the contribution of atomic lines presented briefly in these concluding considering a self -absorbed, nonadi- paragraphs. abatic flow field.The effects of ablation-product radiation in the flow Crew Systems and Functions - field and wake should be examined in Establish crew timelines, functional de tail. Investigate the properties of operations, and associated crew high-density carbon-based ablative system requirements (e. g. , free materials in the regions of high heat volume per man, displays and con- flux. Experimentalverification of trols, living quarter arrangements, ablation rates and material reactions etc. ). These data are needed for in chemical models of the candidate the design and planning of ground and atmospheres must beobtained. The Earth-orbital tests. heatshield design currently is envi- sioned to be composed of a large Vehicle Design - Determine the number of panels which must cover a design requirements for the modular surface area of approximately 750 m2 approach to spacecraft synthesis, (8000 ft2). m-e-se panels are jetti- including the arrangements of the soned after -exit from the'atmosphere. major modules (i.e., ERM, MEM, Determine the performance of such a probes) within the spacecraft, heat- segmented heatshield design and shield joints, separation planes, investigate other candidate heatshield external appendages (such as radia- concepts.Determine entry-velocity tors and antennas), and manufactur- and angle -of-attack requirements for ing considerations. a common heatshield for both planets.

Aer odynamics and Entry Performance - Establish the aero- dynamic coefficients and stability

CR-1131 NASA-Langley, 1968 - 30 23