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ASTROSCIENCES

Report No. T-29

MARS SURFACE SAMPLE RETURN MISSIONS

VIA SOLAR ELECTRIC PROPULSION

IIT RESEARCH INSTITUTE

10 West 35 StreetChicago, Illinois 60616

https://ntrs.nasa.gov/search.jsp?R=19720005139 2020-04-06T06:46:28+00:00Z

Report No. T-29



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, ManagerAstro 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.

/ I T R E S E A R C H I N S T I T U T E

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_..__.

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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 R E S E A R C H I N S T I T U T E

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 MarsSample 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 for1155-Day MSSR Mission, SEP not Staged 35


14 Single Launch Solar Electric Performance for1055-Day MSSR Mission, SEP not Staged(Space Storable Retro) 37

NT R E S E A . R C H I N S T I T U T E

VIII

Figure Page

15 Single Launch Solar Electric Performance for1055-Day MSSR Mission, SEP not Staged(Solid Retro) 38


17 Single Launch Sdlar Electric Performance for950-Day.MSSR Mission, SEP not .staged 40

18 Single Launch Solar Electric Performance for950-Day MSSR Mission with SEP Staging 42




22 Dual Launch Solar Electric Performance forMSSR Missions 46

23 Single Launch Solar Electric Performance for680-Day MSSR Mission 48

24 Single Launch/Two Probe Solar ElectricPerformance for 680-Day MSSR Mission 49

25 Trajectory Profile for 600d MSSR Mission WithInbound Venus Swingby 50

26 Single Launch Solar Electric Performance for600-Day MSSR Mission With Inbound Venus Swingby 52

27 All-Ballistic Performance for 1040-DayMSSR Mission 62

28 All-Ballistic Performance for 625-DayMSSR 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




8 ' Baseline Mission 5 - Summary 59


10 MSSR Mission Selection Summary 61


X

I I T R E S E A R C H I N S T I T U T E

xi



1. INTRODUCTION . . ,

1.1 Study Background

A logical follow-up of the Viking project would be a

mission to return samples of the Martian surface to Earth. The

recent success of the Soviet Luna 16 mission has demonstrated

that automated sample return is a technically feasible concept.

Thus, there is renewed interest within NASA in automated Mars

surface sample return (MSSR) missions.

Previous studies by Niehoff (1967) and Odom (Feb. 1970)

have dealt with MSSR missions in the mid to late '70's. These

studies were concerned primarily with ballistic-type missions

using the Saturn V class launch vehicle. A follow-on study by

Odom (Nov. 1970) included the use of solar electric propulsion

(SEP) and smaller classes of launch vehicle, emphasizing mission

opportunities in the mid to late "70's. The present study

described in this report is based, in part, on unpublished work

initiated in November 1970 for the Planetary Programs Office,

OSSA.

1.2 Study Objectives and Approach

The objectives of this study are the following:

© Determine the capability of solar electric

propulsion for performing Mars surface

sample return (MSSR) missions.

. .. . J..I.T .R.ES.E.AR.CH. INSTITUTE

1

9 Identify various mission and system design

options and show their performance tradeoffs.

© Match mission requirements with candidate

launch vehicles of the Titan III and

Intermediate-20 class.

• Present results in terms of such significant

parameters as sample size, flight time, SEP

power required, and propulsion on-time.

© Define the most promising application of

solar electric propulsion for reducing

mission duration and/or launch vehicle

requirements.

The set of guidelines and constraints used throughout

the study are the following:

o 1981-82 launch opportunity.

o Solar electric stage used for at least one

propulsive phase of the mission. Assume

3500 sec I for all SEP stages.

© Return samples in the range 5-25 kg.

9 Mars orbit rendezvous mode.

• Mars orbiter and lander science is

secondary to primary objective of sample

return.


2

® Use of lifting entry (offset C.G.) at Mars..

© Use of Northrop mass scaling assumptions

for lander/ascent vehicle.

© Earth storable propellants for Mars ascent

stage (e.g., N20^/Aerozine-50, I =310 sec).

® Limit Earth reentry speed to 40,000 ft/sec.

© Utilize existing trajectory date where

possible.

1.3 Mission Phase Options

The MSSR mission profile was separated into the following

distinct phases:

Earth launch,

Earth-Mars transfer,

Mars capture,

Mars landing,

Mars escape,

Mars-Earth transfer, and .

Earth recovery.

Figure 1 depicts, in flowchart form, the options which

can.be associated with each phase of the mission profile. The

selection of various options for each phase was made keeping in

mind the study constraints listed above. It will be seen in

Sections 4 and 5 that not all possible combinations of options

suggested in Figure 1 were considered in this study.

-1 IT -RES EAR C H I N ST.I.T.UJT E

FIGURE I. OPTION ARRAY SET

MISSION PHASE OPTION SELECTION FLOW CHART

EARTH LAUNCH

EARTH-MARSTRANSFER

MARS CAPTURE

MARS LANDING

MARS ESCAPE

MARS-EARTH

TRANSFER

EARTH RECOVERY

T3D/CENT hf- T 3D(7)/CENT I NT-2O

DUAL LAUNCH MODE

SOLAR ELECTRIC

INT-20/CENT

I OR 2LANDERS

BALLISTIC

SOLAR ELECTRICSPIRAL

CHEMICAL RETRO

DIRECT ENTRY

XX

OUT-OF-ORB IT

ASCENT TO ORBIT & RENDEZVOU


vousl

CHEMICAL RETRO

SOLAR ELECTRICDIRECT

SOLAR ELECTRICVENUS SWINGBY

DIRECT REENTRY ORBIT CAPTURE(SOLID RETRO)

XX

X ASCENT PROPULSION-EARTH STORABLE PROPELLANT

XX CHEMICAL RETRO OPTIONSI.) SOLID RETRO, 2.) SPACE STORABLE PROPELLANT

4

Earth Launch

The options considered for the Earth launch phase were

the Titan IIID/Centaur, Titan IIID(7)/Centaur, Intermediate-20

and Intermediate-20/Centaur launch vehicles using a 100 N.M.

parking orbit.

A "sub-option" which was considered for the Titan III

option is that of the dual launch. With this concept, the Mars

lander and Earth return stage are launched in separate launch

vehicles. This concept will be discussed in more detail in

Section 4.

Earth-Mars Transfer

Two types of interplanetary transfers were considered

for this phase; solar electric low-thrust and ballistic. Each

of these options can be classified by either of two types of

transfers: the so-called direct and indirect solar electric•^ ' •

transfers, and the opposition and conjunction type ballistic

transfers. The main difference between direct and indirect, and"

also opposition and conjunction, is that the former type transfer

can be characterized as having relatively higher Earth escape

and Mars approach velocities, and shorter flight times than the

latter type transfers. Also, indirect SEP transfers are charac-

terized by a heliocentric travel angle greater than 360° (i.e.,

more than one revolution about the Sun). •

Note that for Earth-Mars transfers using the Inter-

media te-20 class vehicle, SEP was not needed to achieve the

desired outbound payload.

IJJ_ RfS_EARCH _ [_N_STTT U_T_E_

5

Mars Capture

All systems options require a Mars orbiting bus for the

return phase of the missions. :

The two options considered for capture into Mars orbit

were a SEP low-thrust spiral maneuver, and a chemical high-thrust

retro maneuver. For the chemical retro case, both solid pro-

pellant and space storable liquid propellant stages were

considered. Capture velocity requirements for the assumed orbit

are discussed in Section 2. ,

Mars Landing

Two options were considered for Mars entry. For direct

entry, the lander enters the Martian atmosphere directly from

the hyperbolic approach trajectory, having separated from the

orbiter bus before it maneuvers into Mars orbit. The second

option is to have the lander enter Mars orbit with the orbiter

and then descend from orbit to the Martian surface.

As compared to the orbit capture option, direct entry

would have more critical approach and entry guidance and control

requirements and no landing site selection from orbit, but a

lower orbit capture stage requirement. It was decided that

savings in capture stage weight far outweighed critical guidance

and lack of site selection. Therefore, the direct entry option

was used almost exclusively throughout this study.

System scaling assumptions for the entry vehicle are

discussed in Section 2.

I l l R E S E A R C H I N S T I T U T E

6

Mars Escape

The same options were considered for escape from Mars

orbit as for capture into orbit, i.e., SEP low-thrust spiral or

chemical high-thrust retro. Escape velocity requirements are

discussed in Section 2.

Mars-Earth Transfer

The return-to-Earth transfers considered in this study

are essentially of the direct-type SEP low-thrust. A Venusswingby mode was also examined in which a SEP Mars-Venus transfer

*was matched to a Venus-Earth free return trajectory .

Earth Recovery

Two methods for recovery of the sample container were

considered. The first, direct entry, assumes that the sample

container enters the Earth's atmosphere directly from the hyper-

bolic approach trajectory. No consideration was given as towhether the capsule should be air snatched or surface recovered.

The other available option is to .have the capsule put into a "loose Earth orbit via a chemical retro stage, and then recovered

from orbit, perhaps by a manned vehicle. Only a solid propellant

stage was considered for performing, this maneuver.

Option Selection Example

Figure 1A presents an example of how the options may be

selected for the various phases of a mission. The particular

example shown uses a Titan HID(7)/Centaur single launch with an

SEP Earth-Mars transfer.. Mars orbit is via chemical retro and

/vSearches for Venus-Earth SEP transfers were not made due tolimited time available for the study.

. - _ IIT R - E S E A R C H INSTITUTE-

7

FIGURE IA. OPTION SELECTION EXAMPLE

MISSION PHASE OPTION SELECTION FLOW CHART

EARTH LAUNCH

EARTH-MARSTRANSFER

MARS CAPTURE

MARS LANDING

MARS ESCAPE

MARS-EARTHTRANSFER

EARTH RECOVERY

T3D(7)/CENT

SOLAR ELECTRIC

ICHEMICAL RETRO

**

< IN-

ENTRY

1ASCENT TO ORBIT a RENDEZVOUS


SOLAR ELECTRICDIRECT

ORBIT CAPTURE(SOLID RETRO)

* ASCENT PROPULSION-EARTH STORABLE PROPELLANT

** CHEMICAL RETRO OPTIONS

I.) SOLID RETRO, 2.) SPACE STORABLE PROPELLANT

8

the landing is by direct entry. Escape from Mars orbit is by

SEP spiral and the Mars-Earth transfer is SEP. Finally, orbit

capture of the sample container is selected for the Earth

recovery phase. .

Report Organization

The remaining sections will discuss in detail the

analysis of MSSR missions. Section 2 presents system scaling

assumptions and mission velocity requirements used throughout

the study. Section 3, will show characteristics of the solar

electric low-thrust transfers and maneuvers that apply to MSSR

missions. Section 4 presents a set of conceptual mission charac-

teristics in parametric data form. And Section 5 contains

design-point mission examples using the data from Section 4.

I I-T—RESEARCH -I NSTITUTE

9

Page Intentionally Left Blank

2. SYSTEM SCALING ASSUMPTIONS

2.1 Launch Vehicle Data

Figure 2 presents curves of maximum injected mass as a

function of hyperbolic launch velocity for the four launch

vehicles used in this study. The data for the Titan III class

vehicles was provided by the Jet Propulsion Laboratory, and the

data for the Intermediate-20 class vehicles was taken from the

1971 Launch Vehicle Estimating Factors Handbook.

2.2 Stage Mass Data

Table 1 presents data which was assumed for scaling of

various systems for MSSR missions.

The following sketch presents a possible system configu-

ration concept for a MSSR mission. The system shown would employ

SEP for both outbound and inbound interplanetary transfers, direct

entry of the Mars lander/ascent probe, chemical propulsion for

both capture and escape at Mars, and Earth oVbit recovery of the

sample container. The schematic is taken from Odom (Nov. 1970).

LANDER/RETURN(ASCENT)PROBE

EARTH

,JUC

WAV DOCKING PORT

PROBE ADAPTER

BRAKING/ DEPARTURE STAGE

ORBITER/BUS MODULE

ELECTRIC THRUSTER MODULE

11

28

26

24

22

20

i 82

6

INTERMEDIATE-20/CENTAUR

INTERMEDIATE -20

TITAN mD(7)/CENTAUR

TITAN 3ED/CENTAUR

HYPERBOLIC LAUNCH VELOCITY, KM/SEC

FIGURE 2. LAUNCH VEHICLE PERFORMANCE CURVES

12

TABLE 1

SYSTEM MASS SCALING RELATIONSHIPS

SOLAR ELECTRIC PROPULSION

Specific Mass

Tankage Factor

Specific Impulse

Overall Efficiency

CHEMICAL RETRO STAGES

Solid Propellant

Space Storable Propellant

MARS ASCENT VEHICLE

Earth Storable Propellant

LANDER/PROBE SUPPORT MASSES

Sterilization Canister

Probe Mounting Structure

MARS ENTRY/DBSCENT MASS RATIO

Lifting (Offset C.G.) Entry

SPACECRAFT EQUIPMENT MODULE

Interplanetary Cruise andOrbiter/Bus

30 kg/kw

37o of Propellant Loading

3500 sec

66%

SPECIFIC IMPULSE INERT FRACTION

300 sec

400 sec

0.11

0.20

310 sec 0,20 (1st stage)

0.25 (2nd stage)

12% of Entry Vehicle Mass

(25% for Two Landers)

Entry MassGross Landed Mass

453 kg (outbound)

340 kg (inbound)

= 1.25

-IJ-T- -R-E.S-EAR C-H.—I-N.SJ..I-T..U.T-E-

13

The propulsion stage data, both SEP and chemical, is

representative of current to mid-1970's technology. The

specific impulse of 3500 seconds for SEP is a constraint of the°fc • '

study „ As mentioned previously, both space storable and solid

propellant stages were considered for.the retro capture and

escape maneuvers at Mars. As will be seen in Section 4 space

storable propulsion systems provide better performance, but based

on current technology, are more costly to develop than solid

propulsion systems. • Thus a tradeoff based on cost-effective

performance would have to be made prior to selection of a particu-

lar system. The use of an Earth storable two-stage system for

Mars ascent is based on results of a previous study by, Niehoff

(1967),

The sterilization canister, or bioshield, provides con-

tamination protection to the lander/ascent vehicle from steri-

lization at Earth to Mars arrival. The probe mounting structure

provides the mechanical interface between the lander/ascent

vehicle and the orbiter bus. It also supports the sterilization

canister. The combined mass of the two systems is taken as 12%

of a single lander/ascent probe's mass.

The entry technology assumed in this study was that

derived from the entry analysis performed by Northrop.(Odom,

Nov. 1970). The deceleration system employs a blunt cone aero-

shell utilizing lifting entry, an attached inflatable decelerator,

and a terminal liquid propulsion system. The entry weight to

landed weight mass ratio is assumed to be 1.25; this low mass

ratio is a critical factor in allowing the use of Titan class

launch vehicles,

<JL>

"A 3500 sec specific impulse is representative of current ionthruster development. This value may not be optimum for theMSSR mission; the effect of changing I should be studied.sp


14

The 453 kg spacecraft equipment module (outbound) allows

for such items as structure, telecommunications, navigation and

attitude control, and a certain amount of orbiter science. For

the inbound transfer, 113 kgs is discarded prior to leaving Mars

orbit. This would include such items as the now-unnecessary

docking mechanism and structure, and the orbiter science instru-

ments and associated equipment.

Figures 3 and 4 present the masses of various stages of

the Mars lander vehicle as functions of surface sample mass. The

scaling of all stages was assumed to be linear with sample size.

Note that the Earth recovery systems, i.e., the aerobraking

system for direct reentry, or the solid propulsion stage for

orbit capture mode, remain with the orbital bus in Mars orbit.

The sample container is then transferred to the recovery system

upon rendezvous of the Mars ascent vehicle with the orbiter bus.

As an example of sizing the various systems of the lander/

ascent probe, consider a sample size of 10 kgs. From Figure 3,

the total entry vehicle mass is 2803 kgs and the gross landed

mass is 2345 kgs. Thus, the Mars aerobraking system and descent

propellants total 558 kgs. The ascent vehicle mass is 1330 kgs,

which then allows 915 kgs for the lander. Some of the lander

subsystems and their approximate masses (Odom, Nov. 1970) are:

structure and landing gear, 300 kg; guidance, control and communi-

cation, 140 kg; power, 100 kg; terminal descent propulsion hard-

ware, 75 kg. A portion of the lander's mass may toe allocated for' " • ' • ' -

in situ science instruments and perhaps a small surface rover.

2.3 Mission Velocity Data

Table 2 presents the velocity data which were used in

this study for the scaling of various systems. The data for the

Earth-Mars SEP transfers were obtained by scanning the Earth-

Mars transfer data from Horsewood (1970) for transfers with

—I IT—RESEARCH I NSTI-TU-T-E—

15

3500

3000

2500

o

V)V)

2000

1500

10000

.GROSS LANDED WEIGHT

ASCENT VEHICLE

10 15

SURFACE SAMPLE,KG

20 25

FIGURE 3. MASS SCALING FOR LANDER/ASCENT PROBEMARS SAMPLE RETURN MISSION

16

2501-

200

150

e>

100

50 -

ou

X

XX

0

ASCENT VEHICLEPAYLOAD

XX

X

XX

X

SAMPLE CAPSULE(DIRECT EARTH REENTRY)

AEROBRAKING SYSTEMIN MARS ORBITAL BUS

J_

SAMPLE CAPSULE(EARTH ORBIT CAPTURE)

_L10 15 20

SURFACE SAMPLE,KG

25

FIGURE 4. MARS SAMPLE CAPSULE AND ASCENT VEHICLEPAVLOAD

17

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N T R E S E A R C H I N S T I T U T E

18

appropriate trip time and velocities. The SEP trajectories will

be discussed in more detail in Section 3.

The data for the ballistic Earth-Mars transfers were

taken from the Northrop study as representative of ballistic

transfers for the launch opportunity.

By comparing the velocity data between the ballistic and

SEP transfers, it can be seen that the SEP mode has, in general,

lower velocity requirements. In particular, this is most evi-

dent with the indirect SEP transfer. This effect is mainly due

to the longer flight times of the SEP transfers considered for

this study, and also to the general nature of SEP trajectories.

The characteristic AV of 4.26 km/sec for ascent to a

1000 km altitude circular orbit from the Martian surface is the

result of a numerically integrated trajectory solution. A

circular orbit is desirable for the orbiter bus because of the

requirement for automated rendezvous with the ascent probe. The

1000 km altitude was chosen both as a rough tradeoff point

between capture stage and ascent stage requirements and because

of a sterilizable propellant constraint below 1000 km.

For the Mars-Earth transfers, VHL at Mars was arbi-

trarily set to 0 km/sec for direct SEP transfers, and 2 km/sec

for the SEP/Venus swingby transfer mode. The VHP at Earth was

set to 5 km/sec to correspond with an Earth reentry speed of

40,000 ft/sec.

Finally the capture orbit at Earth of 555 km x 9000 km

altitude is similar to that used in the Northrop study. The orbit

selection was based on the use of an orbit-launched, fully loaded

Apollo GSM, or a system such as the proposed Earth Orbital Space

Tug (if it is operational by the early '80's), for retrieval of

the sample container.

_lj:.T-—R-E.S-E A-R.C H- .J..N.ST-.I-T-U-T-E-

19

3. SOLAR ELECTRIC TRAJECTORY REQUIREMENTS

3.1 Interplanetary Transfers

The analysis of round-trip missions requires a survey of

compatible outbound and return trajectories. SEP trajectory

data was generated to satisfy the 1981-82 launch opportunity as

specified in the list of study constraints. The CHEBYTOP com-

puter program was employed for this purpose (Hahn, et.al. 1969).

A convenient way of presenting the trajectory energy

requirements is shown in Figure 5. The energy measure used is

"J" which is given by the time-integral of a /G(R), where a(t)

is the thrust acceleration magnitude and G(R) is the normalized

solar power (relative to R = 1 a.u.) available to the thrust

subsystem. The parameter J is related to the propellant expendi-

ture; suffice it to say that the lower the J value the lower the

propellant expenditure. .Figure 5 shows constant J contours

plotted in a grid of Earth launch and arrival dates (abscissa)

and Mars arrival and departure dates (ordinate). The outbound

transfers are of the direct type with the exception of the 550-

day indirect transfer point shown. Return transfers to the right

of the slanted broken line are direct while those to the left are

indirect. This type of data map is convenient for determining

suitable launch and arrival dates and the effect of varying trip

time and stay time at Mars. A 950-day mission is shown as an

example, departing Earth on Julian date 2444950 (Dec. 11, 1981),

arriving Mars 2445300 (Nov. 26, 1982), staying 240 days, depart-

ing Mars 24445540 (July 28, 1983), and returning to Earth on

2445900 (July 18, 1984). It will be noted that both the outbound

and return legs are near-minimum energy direct transfers.

Furthermore, the steep-ridge characteristic of the J contours

indicates that an attempt to reduce trip time below 950 days will

-ill- R E S E A R c H I_N STJJJU TE

21

COT) 3iva HONnvi savw

•z.cc.oUJcrUJ_ja.

— tototr

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22

meet with a rapidly increasing energy requirement. It will be

shown subsequently that the shorter missions require a much

larger spacecraft mass at Earth departure than would be possible

using Titan/Centaur launch vehicles„

Upon examining the J contour map, a set of several

outbound and return transfers were selected for further analysis

of sample return capability. These are shown in Figure 6 where

net mass fraction is plotted as a function of normalized power/

mass ratio.. For the return transfers, m is the initial mass of

the return vehicle after Mars escape, but PQ is still the SEP

power referred to a distance of 1 a.u. It is desirable to

choose a design point providing a maximum value of net mass

fraction. As seen from Figure 6 this generally occurs near the

minimum value of P /m , below which the thrust acceleration iso o'insufficient to accomplish the trajectory in the given time of

flight. Design points to the right of this cut off would have

decreasing propulsion on-time requirements as P /m increases.

In the case of the return transfers the design point may not be

chosen arbitrarily. For example, if the same SEP system is

utilized for both the outbound and return legs, then P is fixed

and the ratio P /m is determined by the resulting mass at Mars

departure. In such cases P /m is typically well to the right

of the minimum acceleration cut off. While this may not be

detrimental to the mission objectives, it does raise the possi-

bility of considering a staged SEP design. In Section 4 of this

report several combinations of the selected outbound (SEP and

ballistic) and return transfers will be described as to their

sample return capability.

3.2 Mars Spiral Capture and Escape

The introductory remarks on mission phase options

mentioned both chemical retro and SEP modes for orbit capture

TL ?E?LEARcH INsTijryTE

23

LAUNCHUD.) FLIGHT TIME Voo(LAUNCH) Veo(ARRIVAL)

©©©0©©(7)

| FA " I ' ** ' ^ ' »

EARTH- MARS

EARTH-MARS

MARS-EARTH

MARS-VENUS

MARS-EARTH

MARS-EARTH

MARS- EARTH

244-4950

4845

5540

5200

5240

5240

5640

350 DAYS

550

360

190

560

380

360

3 KM/SEC

0.91

0

2

0

0

0

1 KM/SEC

0-J7

5

11.08

5

5

5

0.8-

0.6-

CO

2I-LU

0.4 -

0.2

NET MASS=INITIAL MASS -( PROPULSION SYSTEM -I-PROPELLANT + TANKAGE)

INSUFFICIENT THRUST ACCELERATIONBELOW THIS POWER LEVEL FORIsp = 3500 SEC

2 4 6 8 10 12 14 16 18 20 22 24

NORMALIZED POWER AT IAU (Po/Mo), KW/IOOOKG

FIGURE 6. NET MASS FRACTION VERSUS SOLAR ARRAY POWER

24

and escape maneuvers. In the SEP mode these maneuvers are

characterized by multi-revolution spirals about Mars due to the

low-thrust acceleration levels. The spiral requirements are

shown in Figures 7 to 10, are based on analytical solution formu-

las (Ragsac 1967). Spiral maneuver time and final/initial mass

ratio are given as a function of thrust acceleration. For the

capture spiral a, is the initial acceleration available on the

hyperbolic approach asymptote. For the escape spiral a is the

initial acceleration available upon leaving the circular orbit

about Mars. Three values of orbit altitude are shown for compari-

son purposes, but only the 1000 km orbit is used in the subsequent

mission analysis. It should be noted that the acceleration value

used must take into account the actual value (P G(R)) of solar

power available at Mars'distance (approximately 1.5 a.u. but

variable as a function of date).

A typical thrust acceleration a, would be 3 x 10 m/sec .

The capture time is then 130 days and the final mass fraction is

0.902 (or, 0,098 propellant fraction). As an example, suppose/ o

a is somewhat higher at 4 x 10~ m/sec due to a reduction in

mass. The escape time is then 80 days and the final mass fraction

is 0.921 (or, 0.079 propellant fraction). Because the SEP system

operates with a specific impulse about an order of magnitude

higher than the chemical retro systems, the resulting propellant

fraction is very significantly lower. The penalty incurred is a

long and somewhat complex (solar array pointing) maneuver.

Another disadvantage is the long time that the Mars lander must

wait on the surface or in orbit before the SEP return stage

reaches the rendezvous altitude of 1000 km. Nevertheless, if

these operational difficulties can be tolerated, the SEP capture

and escape modes can be expected to yield a large performance

advantage over chemical retros; this is particularly important

when Titan/Centaur launch vehicles are employed. This point will

be shown in the following section.

CJH LNSTJTJJTE ^

25

<Q

LJ2

a:D.

LUor

CL<u

1000

800^

600

HYPERBOLIC APPROACH VELOCITY = IKM/SECSEP SPECIFIC IMPULSE =3500 SEC

CIRCULAR ORBIT ALTITUDE

IOOOKM

10,000 KM

20,000 KM

200.5 2 , 4 , 6

THRUST ACCELERATION 0H , IO"4 M/SEC2

FIGURE 7. MARS LOW THRUST CAPTURE SPIRAL REQUIREMENTS(SPIRAL TIME)

26

HYPERBOLIC APPROACH VELOCITY = IKM/SECSEP SPECIFIC IMPULSE =3500 SEC.

0.96


0.892 3 4 5

THRUST ACCELERATION Ou , 10* M/SEC2PI

FIGURE 8. MARS LOW THRUST CAPTURE SPIRAL REQUIREMENTS(MASS RATIO)

27

ESCAPE VELOCITY VHL =0ISp = 3500 SEC


10,000 KM

20,000 KM

THRUST ACCELERATION dc , (6* M/SEC2

FIGURE 9. MARS LOW THRUST ESCAPE SPIRAL REQUIREMENTS(SPIRAL TIME)

28

0.97

0.96

0.95

cc 0.94COCO

_J

pz 0.93

0.92

0.910

ESCAPE VELOCITY VH|_ = 0

SEP SPECIFIC IMPULSE 3500 SEC


20,000 KM

1000 KM

THRUST ACCELERATION O c , l 6 4 M/SEC 2

FIGURE 10. MARS LOW THRUST ESCAPE SPIRAL REQUIREMENTS(MASS RATIO)

29

4. MISSION PERFORMANCE CHARACTERISTICS -

This section describes a set of possible MSSR missions,

arrived at using the guidelines, trajectory and system data from

the previous sections. The approach taken in this study was to

selectively match outbound and inbound transfers to create a set

of mission trajectory profiles, From the set of mission phase

options (Figure 1), combinations of selected options were con-

sidered for each mission profile. The mass fraction and system

scaling data were then used to size the total system requirements,

depending on phase option^ for each mission profile.

The results of this approach are shown in Figures 11

through 26. The set of mission profiles are depicted by polar

heliocentric trajectory plots; associated launch and arrival

dates are indicated on each diagram. The figure(s) following

each trajectory plot presents system mass data dependent on the

selected phase options for the mission profile. This data is in

parametric form; Earth departure mass as a function of desired

sample size. On each payload curve, the range of power required

at 1 a.u. for the SEP stage(s) is indicated. Also, the launch

vehicle capability at the particular Earth departure VHL is

shown.

As an aid in determining the combinations of phase options

for the missions considered in this study,. Table 3 presents which

combination of options relates to which payload figure(s).

Whether the chemical retro option, where used, is solid or space

storable propellant, and whether Earth recovery is by direct

reentry or orbit capture, is indicated on each of the figures.

Figure 11 shows the trajectory profile for an 1155-day

MSSR mission. The outbound and inbound transfers are both SEP

low-thrust, the outbound leg being of the indirect mode. A

I IT R E S E A R C H INSTITUTE

31

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32

245d ORBIT WAITAND SPIRALMANEUVERS

\ MARSARRIVAL

/EARTH ATMARS ARRIVAL

EARTHDEPARTURE

EARTH DEPARTUREMARS ARRIVAL

MARS DEPARTUREEARTH ARRIVAL

AUG. 29, 1981MAR. 2, 1983

NOV. 2 , 1983OCT. 27, 1984

FIGURE TRAJECTORY PROFILE FOR H55dMSSR MISSION

33

245-day stop-over time occurs at Mars which can be used either

totally, for orbit wait or for SEP spiral maneuvers plus orbitwait. Figure 12 shows the injected payload data for this missionif SEP spiral maneuvers are used for both Mars capture andescape. All systems are launched by a single vehicle, and the

same SEP stage is used for the outbound and inbound transfersand the spiral maneuvers. The launch date for this mission isAugust 29, 1981. It will be noted that the Titan IIID/Centaurlaunch vehicle is capable of returning a 10 kg sample in eitherEarth recovery mode, and the SEP power requirement is about 18 lew.

^

Figure 13 shows a 1055-day mission which uses the sameEarth-Mars transfer shown in Figure 11. For this mission, only145 days are available at Mars. This amount of stop-over timedid not allow for both SEP spiral capture and escape. Figure 14presents payload data for two mission concepts using thisprofile. One assumes a space storable chemical retro escape,while the other uses a SEP spiral escape. Both use space stor-able retro capture. Again, this is for a single launch withoutSEP staging.. Figure 15 shows the same concept using solidchemical retro stages. The Titan IIID(7)/Centaur launch vehicleand the spiral escape mode are both necessary to return anominal sample of at least 5 kg. A typical SEP power require-ment would be 20-25 kw.

Figure.16 shows a 950-day mission. Both outbound andinbound transfers are of the direct SEP type. A stop-over timeof 240 days are allowed at Mars, giving time to use spiral maneu-vers for capture and escape. The launch date is December 11, 1981,Figure 17 shows payload data for concepts which use either asolid retro capture or SEP spiral capture; both concepts usespiral escape. The data is for a single launch without SEPstaging. The Titan HID(7)/Centaur and spiral escape provide a

return sample of at least 5 kg.

I.IT RESEARCH INSTITUTE

34

7j-TITAN JHD(7)/CENT.

• DIRECT MARS ENTRY• SEP SPIRAL CAPTURE

AND ESCAPE• EARTH RECOVERY MODE

ORBIT CAPTUREDIRECT REENTRY

o

20.9 KW

20.4KW

o0o

6COCO<t2

UJ(T

DE

PA

RT

X>

K *< 5Ul

TITAN DID/CENT

17.2 K W X 1

or

=I6.8KW

I

10 15

SURFACE SAMPLE , KG

20 25

FIGURE 12. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 1155-DAY MSSR MISSION, SEP NOT STAGED

35

MARSDEPARTURE

!45dORBIT WAITAND/OR SPIRALMANEUVERS

TO EARTHtSEP)

EARTH ATMARS ARRIVAL

\\\\

MARS\ ARRIVAL

\EARTHDEPARTURE^

//i

_J ^,

/ My)/

y

EARTH DEPARTUREMARS ARRIVALMARS DEPARTURE

EARTH ARRIVAL

AUG. 29, 1981MAR. 2, 1983JULY 24,1983

JULY 18,1984

FIGURE 13. TRAJECTORY' PROFILE FOR 1055dMSSR MISSION

36

oooo

V)V)<

LJCC^>Ho:<0.UJo

a:<UJ

8

• DIRECT MARS ENTRY

• SPACE STORABLE CAPTURE

• EARTH RECOVERY MODE

ORBIT CAPTUREDIRECT REENTRY

29.4 KW

- TITAN IE D(7)CENT.«

20KW,O

TITAN m D/ ffENT

S/S

^CHEMICALESCAPE

.11 SEP'i-SPIRAL

-l' ESCAPE

Po = 23.2KW

10 15

SURFACE SAMPLE, KG

20 25

FIGURE 14. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCE

FOR 1055-DAY MSSR MISSION.SEP NOT STAGED(SPACE STORABLE RETRO)

37

e>xooo

V)V)

UJor

UJQ

X(-crUJ

1 1

5 -

• DIRECT MARS ENTRY

• SOLID CHEMICAL CAPTURE

• EARTH RECOVERY MODEORBIT CAPTUREDIRECT REENTRY

TITAN mD(7)/CENT

•') SOLID[•CHEMICALJ ESCAPE

1SEP

fSPIRALJJ ESCAPE

Po =20.2KW

TITAN ffl D/CENT

J_10 15

SURFACE SAMPLE,KG

20 25

FIGURE 15. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 1055-DAY MSSR MISSION.SEP NOT STAGED(SOLID RETRO)

38

MARSDEPARTURE

240d ORBIT WAITAND/OR SPIRALMANEUVERS

EARTH ATMARS ARRIVAL \

MARSARRIVAL

EARTH DEPARTUREMARS ARRIVAL .

MARS DEPARTURE

EARTH ARRIVAL

DEC. II, 1981NOV. 26, 1982

JULY 24, 1983

JULY 18, 1984

FIGURE 16. TRAJECTORY PROFILE FOR 950° MSSR MISSION

39

OOO

V)

LUo:

a:2UJQ

or

ui

7.5

7.0

6.5

6.0

5.5

5.0

4.5

4.0

• DIRECT MARS ENTRY• SEP SPIRAL MARS ESCAPE

• SOLID RETRO

• EARTH RECOVERY MODE

ORBIT CAPTUREDIRECT REENTRY 29.IKW

CHEMICAL

CAPTURE

SEPSPIRALCAPTURE

X0^

/ 22.5KW

XPo =I8.6KW

TITAN TED/CENT

_L J_O 10 15

SURFACE SAMPLE, KG

20 25

FIGURE 17. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 950-DAY MSSR MISSION, SEP NOT STAGED

40

Figure 18 shows data for concepts which stage the SEP

module. Both chemical (solid retro) and SEP capture options areconsidered, but the Mars escape maneuver is via chemical retro.

SEP escape was not examined because the relatively low powerrating of the second SEP stage did not allow sufficient acceler-

ation to perform the spiral maneuver in a reasonable amount of

time.

Assuming the all-chemical retro option it is seen that

the staging concept yields about the same performance as theconcept discussed previously where the SEP is not staged and aspiral escape mode is utilized (see Figure 17). TheTitan IIID(7)/Centaur is marginal in either case. However, if

the SEP capture mode is allowed the Titan HID(7) /Centaur is

capable of a 10 kg sample returned to Earth orbit. The two SEPstage power ratings are about 22.6 kw and 2.9 kw, respectively.

Figures 19, 20 and 21 present mission profiles which usethe same conjunction-type Earth-Mars ballistic transfer, with alaunch date of December 1, 1981. The 960-day mission profile inFigure 19 uses a direct type Mars-Earth SEP transfer and allows310 days at Mars. Figure 20 shows an 860-day mission profile witha 10-day orbit wait at Mars and a direct SEP inbound transfer.Figure 21 is for a 680-day mission with a 10-day orbit wait, anda fast, indirect-type Mars-Earth SEP transfer.

Figure 22 presents payload data for dual launch mission

concepts using the three mission profiles that have just beendescribed. The first launch injects the Mars lander vehicle ontothe trans-Mars trajectory, while the second vehicle injects theorbiter bus/SEP return stage. The two vehicles arrive at Mars at

the same time; the planetary vehicle makes a direct entry, whilethe orbiter/return stage enters orbit. The concepts shown

consider only chemical retro capture and escape maneuvers for the


41

ooo

CO

LJOr

a:

LUQ

Xh-o:LJ

s.Or

• DIRECT MARS ENTRY• SOLID, RETRO ESCAPE

• EARTH RECOVERY MODEORBIT CAPTUREDIRECT REENTRY

7.5

7.0

(POi/P02) = (20.2/2.5)KW

POi : FIRST STAGEP02: SECOND STAGE

TITAN HID/CENT

SOLID^RETRO

CAPTURE

SEPSPIRALCAPTURE

4.5

10 15

SURFACE SAMPLE,KG

FIGURE 18. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 950-DAY MSSR MISSION WITH SEP STAGING

42

MARSDEPARTURE 310 ORBIT WAIT

TO EARTH(SEP) EARTH

DEPARTURE

TO MARS(BALLISTIC)

EARTH• AT MARS,ARRIVAL /

\

\

\\

MARSARRIVAL

EARTH DEPARTUREMARS ARRIVALMARS DEPARTUREEARTH ARRIVAL

DEC. I, 1981SEPT. 17, 1982JULY 24, 1983JULY 18, 1984

FIGURE 19. TRAJECTORY PROFILE FOR 960d MSSR MISSION

43

TO MARS(BALLISTIC) DEPARTURE \

\

\

\

XIX )

. MARS/ ARRIVAL /

/ / / T O/ / / EARTH/ ' ' (SEP)

EARTH DEPARTURE DEC I, 1981MARS ARRIVAL SEPT. 17,1982MARS DEPARTURE SEPT 27, 1982EARTH ARRIVAL APR. 9, 1984

MARSARRIVAL

MARS DEPARTURE

I0d ORBIT WAIT

FIGURE 20. TRAJECTORY PROFILE FOR 860d MSSR MISSION

44

\\\

\

EARTHDEPARTURE

TO/ MARS

/ (BALLISTIC)

EARTH ATMARS ARRIVAL

// EARTH / // (SEP)//

EARTH DEPARTURE

MARS ARRIVAL

MARS DEPARTURE

EARTH ARRIVAL

DEC. I, 1981SEPT. 17, 1982SEPT. 27, 1982OCT. 12, 1983

MARS1

ARRIVAL

MARSDEPARTURE

I0d ORBIT WAIT

FIGURE 21. TRAJECTORY PROFILE FOR 680a MSSR MISSION

o*ooo

ui

XUJ>ozooUJ

CrtCrt

LUcr

o.LJoior<UJ

8

DIRECT REENTRY3I.9KW

ITANinDl7VCENT

BALLISTIC EARTH-MARS TRANSFERSDIRECT MARS ENTRY ( FIRST VEHICLE)CHEMICAL RETRO CAPTUREAND ESCAPE

EARTH RECOVERY MODEORBIT CAPTURE

680"S/S RETRO

26.8KW

860°S/S RETRO

I4.6KW

960d

SOLIDRETRO

-o—2.6KW( RETURN SEP STAGE)

_L JL10 15

SURFACE SAMPLE, KG

20 25

FIGURE 22. DUAL LAUNCH SOLAR ELECTRIC PERFORMANCEFOR MSSR MISSIONS

46

orbiter/return stage; the propellents considered are as indicated.It will be noted that the Titan IIID/Centaur has adequate per-

formance capability in the dual launch mode for either the 860-day or 960-day missions. The longer mission, in particular, is

attractive in that the solid retro system can be employed and theSEP return stage requires less than 4 kw power.

Figure.23 shows several concepts for the 680-day missionprofile which uses a single launch and chemical retro capture

and escape at Mars. An INT-20/Centaur launch vehicle would berequired for the out-of-orbit entry mode. The Centaur upperstage would not be needed for the Mars direct entry mode. The

SEP return stage power is greated than 26 kw for a sample sizegreater than 5 kg.

The performance characteristics for two Mars lander

probes launched on the same vehicle are shown in Figure 24, "againusing the 680-day mission profile. Note that only one orbiter/

SEP return stage is employed. This would rendezvous with eachMars ascent vehicle from the two different landing sites. If theMars direct entry mode is employed an INT-20/Centaur launchvehicle would be capable of returning 50 kg of samples - 25 kgfrom each landing site. However, the SEP return stage powerrequirement is greater than 40 kw.

Figure 25 shows a mission profile which uses an oppo-sition-type ballistic transfer to Mars and a Venus swingby return

to Earth. The Mars-Venus transfer is SEP and the Venus-Earthtransfer is ballistic. The launch date is December 21, 1981 with

a total mission time of 600 days and a 20-day stop-over time at

Mars. It will be noted that the return trajectory in this case

.is similar to the all-SEP return on the 680-days mission profile(see Figure 21).


47

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• BALLISTIC EARTH-MARS TRANSFER• CHEMICAL RETRO CAPTURE

AND ESCAPE

SOLID .. . : .SPACE STORABLE

INT-20/CENT

EARTH ORBIT CAPTURE

OUT-OF-ORBIT ENTRY

MARS DIRECTENTRY

Po = 25.6KW(RETURN SEP STAGE)

5 10 15

SURFACE SAMPLE,KG

FIGURE 23. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 680-DAY MSSR MISSION

48

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BALLISTIC EARTH-MARS TRANSFERCHEMICAL RETRO CAPTURE AND ESCAPE

SOLIDSPACE STORABLE

EARTH ORBIT CAPTURE

OUT-OF-ORBITENTRY

22fnNT-20/CENT

18

14

10

INT-20.

4I.4KW

J

MARS DIRECT.ENTRY^""

= 28.8KW(RETURN SEP STAGE)

1 _L0 5 10 15 20

SURFACE SAMPLE, EACH PROBE, KG

25

FIGURE 24. SINGLE LAUNCH/TWO PROBE SOLAR ELECTRICPERFORMANCE FOR 680-DAY MSSR MISSION

49

//

//

/ TO MARS / /(BALLISTIC) / /TO EARTH' * /(BALLISTIC);

EARTHDEPARTURE

MARSARRIVAL MARS

DEPARTURE

EARTH DEPARTURE DEC 21, 1981MARS ARRIVAL JULY 29, 1982MARS DEPARTURE AUG. 18, 1982VENUS SWINGBY FEB. 24, 1983EARTH ARRIVAL AUG. 13, 1983

FIGURE 25. TRAJECTORY PROFILE FOR 600d MSSR MISSION WITHINBOUND VENUS SWINGBY

50

Performance data for the Venus swingby concept is pre-

sented in Figure 26. A chemical retro stage is used for both

Mars capture and escape, and the Earth recovery mode shown is via

orbit capture. The INT-20 would be adequate for a space storable

retro, but the INT-20/Centaur (off scale) would be needed if a

solid retro were utilized, The power of the SEP return stage

lies in the range 15-19 kw for a 5-25 kg surface sample.

The next section will summarize several baseline mission

examples representing a spectrum of the various concepts and

performance data just described, >

I IT R E S E A R C H INST ITUTE

51

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• BALLISTIC EARTH-MARS TRANSFER• CHEMICAL RETRO CAPTURE 8 ESCAPE• SOLAR ELECTRIC MARS-VENUS ONLY• EARTH ORBIT CAPTURE

XXSPACE STORABLE

RETRO

Pq -14. 6KW(RETURN SEP STAGE)

J_10 15

SURFACE SAMPLE, KG

20 25

FIGURE 26. SINGLE LAUNCH SOLAR ELECTRIC PERFORMANCEFOR 600-DAY MSSR MISSION WITH INBOUNDVENUS SWINGBY

52

5. BASELINE MISSION EXAMPLES

The previous section presented a set of representative

MSSR mission concepts in the form of parametric payload data at

Earth departure as a function of sample size. Tables 3 through

8 describe design-point baseline missions taken from the set of

mission concepts. The examples selected encompass the spectrum

of MSSR missions in terms of flight duration, propulsion modes,

launch vehicle, etc. For each baseline mission, a sample size

was assumed which would allow a nominal margin between total

system mass and launch vehicle capability. The masses of vari-

ous subsystems were then calculated on the basis of desired

sample size. All of the baseline missions considered assume the

direct entry option for the Mars lander vehicle and the Earth

orbit capture mode for sample capsule recovery.

Table 4 presents an 1155-day mission which will return a

10 kg sample using a Titan IIID/Centaur single launch. SEP is

used for all major propulsion phases and the total SEP thrusting

time is 784 days. The power requirement of the SEP stage is

18.5 kw at 1 -a.-u. The launch vehicle margin is approximately

100 kgs.

Table 5 presents a 1055-day mission to return a 10 kg

sample. A solid propellant retro system is used for Mars capture,

In comparison with the previous example, note that even though

the SEP thrust time has decreased because of no spiral capture,

the SEP system requirements (mass and power) have increased. The

launch vehicle for this mission is the Titan IIID(7)/Centaur, and

the margin is 425 kgs.

Table 6 lists data for a 950-day mission returning a

10 kg sample. All major propulsion phases are again SEP. The

total thrusting time has been reduced to 586 days because of the


53

TABLE 4 .

BASELINE MISSION 1 - SUMMARY

Sample Size

Launch Vehicle

Mission Duration

10 kg (Earth Orbit Capture)

Titan IIID/Centaur (Single Launch)

1155 Days

Earth-Mars Transfer (SEP)

Mars Capture Spiral (SEP)

Mars Stay Time

Mars Escape Spiral (SEP)

Mars-Earth Transfer (SEP)

FLIGHT TIME

550 days

98

34

113

3601155

System Weight Breakdown

Mars Lander/Ascent Probe (Direct Entry)

; Aerobraking/Propus lion 558

Lander 847

Rover 68

Ascent Vehicle 1330



SEP Stage

Propulsion System (18,5 kw) 555

Propellant + Tankage 985

Spacecraft Equipment Module

Earth Capture Stage (Solid)

Earth Departure Vehicle

Titan IIID/Centaur Capability

I IT RESEARCH INSTITUTE

54

SEP TIME

496 days

98

113

77784

2803 kg

188148:...

1540

4531465278 kg5380 kg

Sample Size

Launch Vehicle

Mission Duration

TABLE 5



Titan IIID(7)/Centaur (Single Launch)

1055 Days

FLIGHT TIME SEP TIME

Earth-Mars Transfer (SEP)

Mars Stay Time

Mars Escape Spiral (SEP)

Mars Earth Transfer (SEP)



Aerobraking/Propulsion 558

Lander 847

Rover 68

Ascent. Vehicle 1330



Mars Capture Stage

SEP Stage

Propulsion System (22.5 kw) 675

Propellant + Tankage . 1033




Titan IIID/Centaur Capability

550 days

49

963601055

496 days

96

76668

2803 kg

188

'148

9791708

4531466425 kg6850 kg

ST T E

55

TABLE 6


Sample Size 10 kg (Earth Orbit Capture)

Launch Vehicle Titan IIID(7)/Centaur (Single Launch)

Mission Duration 950 Days


Earth-Mars Transfer (SEP) 350 days 292 days

Mars Capture Spiral (SEP) 118 118

Mars Stay Time 22

Mars Escape Spiral (SEP) 100 100

Mars-Earth Transfer (SEP) 360 76950 586


Mars Lander/Ascent Probe (Direct Entry) 2803 kg


Lander 847

Rover 68

Ascent Vehicle 1330

Sterilization Canister 188

Probe Mounting Structure 148

SEP Stage 1378



Spacecraft Equipment Module 453

Earth Capture Stage (Solid) 146

Earth Departure Vehicle 5116

Titan IIID(7)/Centaur Capability 5995


56

faster Earth-Mars transfer. The SEP power requirement is 20.5 kw

at 1 a.u. The launch vehicle is again the seven-segment Titan/

Centaur with a margin of nearly 880 kgs.

The fourth example, Table 7, presents a 960-day mission

which utilizes the dual launch concept. The returned sample

size is 20 kgs. Solar electric is used only during the inbound

transfer, and the system power requirement is relatively low at

3.9 kw. The vehicle for each launch is a Titan IIID/Centaur,

and the weight margins are approximately 360 kgs and 1175 kgs.

Table 8 presents a 680-day mission which will return

10 kgs. SEP is used only for the inbound transfer. The SEP

thrusting time is 245 days, and the system power requirement is

nearly 28 kw at 1 a.u. The mission utilizes a single launch

vehicle, the Intermediate-20, with a weight margin of nearly

1000 kgs.

The final example, Table 9, is a 600-day mission which

will return a 25 kg sample. This mission uses a Venus swingby

during the inbound transfer, with SEP used only for the Mars-

Venus leg. The SEP thrust time is only 157 days and the power

requirement is nearly 20 kw at 1 a.u. The required launch

vehicle is the Intermediate-20/Centaur with a margin of over

4000 kgs.

Table 10 summarizes the more pertinent aspects of the

six baseline missions selected as examples./ ' -

For purposes of comparison, Figures 27 and 28 present

three all-ballistic mission concepts. The two concepts in

Figure 27 use the same conjunction type Earth-Mars and Mars-

Earth transfers. Earth departure date is Nov. 23, 1981, with a

total mission duration of 1040 days; the Mars stay time is 420

days. As can be seen, with the phase options indicated, the


57

Sample SizeLaunch VehicleMission Duration

TABLE 7


20 kg (Earth Orbit Capture)Titan IIID/Centaur (Dual Launch)960 Days

Earth-Mars Transfer (Ballistic)Mars Stay TimeMars^Earth Transfer (SEP)

FLIGHT TIME

290 days

310 .

360

SEP TIME

960287 days287


A, First Launch



Lander 912

Rover 68

Ascent Vehicle 1590




Bo Second Launch

Mars Capture Stage (Solid)

Mars Escape Stage (Solid)

SEP Stage

Propulsion System (3,9 kw) 117

Propellant + Tankage 50Spacecraft Equipment Module

Earth Capture Stage

Titan IIID/Centaur CapabilityNT RESEARCH INSTITUTE

3200 kg

215

169 '4534037 kg

1896 kg

500

167

453

2093225 kg4400 kg

58

Sample SizeLaunch VehicleMission Duration

TABLE 8


10 kg (Earth Orbit Capture)Intermediate-20 (Single Launch)680 Days

Earth-Mars Transfer (Ballistic)Mars Stay TimeMars-Earth Transfer (SEP)

FLIGHT TIME

290 days

10380

680


Mars Lander/Ascent Probe (Direct Entry)Aerobraking/Propulsion 558Lander 847

Rover 68

Ascent Vehicle 1130 .

Sterilization CanisterProbe Mounting Structure

Mars Capture Stage (Solid)Mars Escape Stage (Solid)

SEP TIME

245 days

245

2803 kg

SEP StagePropulsion System (27.8 kw)

Propellant + Tankage

Spacecraft Equipment ModuleEarth Capture Stage (Solid)


Intermediate-20 Capability

830631

188148

48071294

1461

453146

113.00 kgs

12250 kgs


59

TABLE 9


Sample Size

Launch Vehicle

Mission Duration


Intermediate-20/Centaur (Single Launch)

600 Days


Earth-Mars Transfer (Ballistic)

Mars Stay Time

Mars-Venus Transfer (SEP)

Venus-Earth Transfer (Ballistic)

220 days

20

190

170

600



Aerobraking/Propuslion 675

Lander 932

Rover 68

Ascent Vehicle 1725



Mars Capture Stage (Solid)

Mars Escape Stage (Solid)

SEP Stage






Intermediate-20/Centaur Capability

I I I R E S E A R C H I N S T I T U T E

157 days

157

3400 kgs

228

180

9066

1421

835

453217

15800 kgs20000 kgs

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TITAN ffl.D(7)/CENT

TITAN HID/CENT

_L10 15

SURFACE SAMPLE, KG

20

SINGLELAUNCH

^DUALLAUNCH

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FIGURE 27. ALL-BALLISTIC PERFORMANCE FOR1040 DAY MSSR MISSION

62

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20

18

16

14

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INT- 20/CENT

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DIRECT MARS ENTRY

EARTH ORBIT CAPTURE

SOLID RETRO STAGES

I

10 15

SURFACE SAMPLE, KG

20 25

FIGURE 28. ALL-BALLISTIC PERFORMANCE FOR 625 DAYMSSR MISSION WITH INBOUND VENUS SWINGBY

63

Titan IIID(7)/Centaur single launch can marginally return a 5 kg

sample, whereas the Titan IIID/Centaur dual launch concept is

capable of returning the full range of sample size.

Figure 28 presents payload data for a mission which uses

a fast, opposition type Earth-Mars transfer and a Venus swingby

return to Earth. The launch date is Nov. 17, 1981, with a total

mission time of 625 days and .a 30-day Mars stay time. The

Intermediate-20/Centaur is required to return samples greater

than about 7 kgs.. Comparing this with the 600-day mission which

uses a SEP/Venus swingby Earth return and solid retro option

(see Figure 26), the all-ballistic mission provides slightly

better performance. This is due largely to the fact that the

SEP stage is being used only for the Mars-Venus transfer and

must be carried as inert mass from the launch phase through the

Mars escape maneuver. A fast, opposition type mission with a

direct Mars-Earth transfer was also examined, but the energy

requirements were much too high for a practical mission

application.

CONCLUSIONS

Solar electric -propulsion can be used effectively to

accomplish the Mars Surface Sample Return 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 vehicle and sample size.

The major results of this study are listed below:

(1) A sample of 10 25 kg can be returned to

an Earth orbit compatible with manned

spacecraft recovery operations.

I l l R E S E A R C H I N S T I T U T E

64

(2) State-of-the-art chemical propulsion

systems may be utilized; solid

propellants for retro maneuvers and

earth-storable liquid propellants for

ascent from the Martian surface.

(3) Titan IIID/Centaur vehicles (5 or 7segment) can be employed in the

single-launch mode provided that SEP

is used for both outbound and return

interplanetary transfers and, at least,

the Mars escape maneuvers.

(4) The above mission concept requires a

total trip time of 2.5 to 3 years, a

powerplant size of about 20 kw, and a

60-70% propulsion duty cycle.

(5) Shorter missions (1.5-2 years) can beaccomplished with the INT- 20 or

INT-20/Centaur launch vehicles.

However, SEP should be used only for

the return transfer in order to limit

the SEP power requirement.

Since a mission duration of 2.5 years does not seem unreasonable,

the best application of SEP may well be Mission Concept No. 3

which utilizes the Titan IIID( 7) /Centaur launch vehicle. There

is a healthy margin of safety between the Earth departure weightand the launch vehicle capability. The problem areas or reserva-

tions concerning this choice are the possible difficulty of

mechanizing the thrust steering program during the Mars spiral

maneuvers, and the long wait (118 days) between landing and

rendezvous with the orbital bus.

65

REFERENCES

Hahn, D. W. , et. al., "Chebychev Trajectory Optimization Program(CHEBYTOP)", Report No. D2-121308-1, Boeing Company,July 1969.

Horsewood, J. L., and Mann, F. I., "Optimum Solar ElectricInterplanetary Trajectory and Performance Data",NASA CR-1524, April 1970.

NASA/OSSA, "Launch Vehicle Estimating Factors", January 1971.

Niehoff, J. C./'Preliminary Payload Analysis of Automated MarsSample Return Missions", ASC Report No. M-13, IIT ResearchInstitute, May 1967.

Odom, P. R., "Automated Mars Surface Sample Return Mission/SystemStudy, Vol. IIA-Mission Analysis", Report No. TR-793-717,Northrop Corporation, February 1970.

Odom, P. R., et. al., "Preliminary Study of Minimum PerformanceApproaches to Automated Mars Sample Return Missions -Final Report", Report No. TR-793-842, Northrop Corporation,November 1970.

Ragsac, R. B. , "Study of Trajectories and Upper Stage PropulsionRequirements for Exploration of the Solar System, Vol. II -Technical Report", Report F-910352-13, United AircraftCorporation, September 1967.


66

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