AlAA 89-0631 MRSR Aerocapture Configuration Design and Packaging Constraints S. Lawson, NASA-Johnson, Houston, TX

27th Aerospace Sciences Meeting January 9-12, 1989/Reno, Nevada

For permission to copy or republish, contact the American Institute of Aeronautics and Astronautics (J 370 L'Enfant Promenade, S.W., Washington, D.C. 20024

MARS ROVER SAHPLE RETURN AEROCAPTURE CONFIGURATION DESIGN AND PACKAGING CONSTRAINTS AIAA-89-0631

Shelby J. Lawson* NASA Lyndon B. Johnson Space Center

Houston, Texas 77058

Abstract

This paper discusses the aerodynamics requirements, volume and mass constraints that lead to a biconic aeroshell vehicle design that protects the Mars Rover Sample Return (MRSR) mission elements from launch to Mars landing. The aerodynamic requirements for Mars aerocapture and entry and packaging constraints for the MRSR elements result in a symmetric biconic aeroshell that develops a L/D of 1.0 at 2 7 . 0 - angle of attack. A significant problem in the study is obtaining a cg that provides adequate aerodynamic stability and performance within the mission imposed constraints. Packaging methods that relieve the cg problems include forward placement of aeroshell propellant tanks and incorporating aeroshell structure as lander structure. The MRSR missions developed during the pre-phase A study are discussed with dimensional and mass data included. Further study is needed for some missions to minimize MRSR element volume so that launch mass constraints can be met.

81-A/D

B2 Areal

D1 Areal

D1 Local

DZ-P/A

D2-P/F

ERV LJD MAV MCO MRSR RCS RDM SRC SCA TPS

Nomenclature

Configuration 81 Aerocapture at Mars/Direct entry at Earth Configuration 82 Aerocapture at Mars with Areal Rover configuration D1 Aerocapture at Mars with Areal Rover Configuration D1 Aerocapture at Mars with Local Rover Configuration D2 Propulsive capture at Mars/Aerocapture at Earth configuration D2 Propulsive capture at Mars/Propulsive capture at Earth Earth Return Vehicle Li€t/nrng Rat io ~~ ~ ~ ~ ~

Mars Ascent vehicle Mapping/Communications Orbiter Mars Rover Sample Return Reaction Control System Rendezvous Docking Module Sample Return Capsule Sample Canister Assembly Thermal Protection System

Introduction

collect and return Mars soil samples to Earth. The MRSR program allows for the development, design and testing of advanced technology that will be required for a manned Mars missions.

Significant mass savings can be achieved by using the Mars atmosphere to reduce the vehicle velocity prior to orbit insertion. This method of orbit capture is defined as aerocapture as opposed to propulsive capture.

Reference Missions Description

L,

The pre-phase A study considered a variety of vehicle element launch combinations, capabilities and features. Three different Rover designs were considered with varying size, capability and range. The local Rover is small (1.5 m x 1.0 m x 1.0 m) with four 0.5 m diameter wheels, has a range of 100 m from the lander and weighs 150 kg. The areal Rover is a two segment vehicle ( 4 . 2 m x 4 . 0 m x 1.35 m) with six 1.0 m diameter wheels, has a range of 2 0 - 4 0 km and weighs 8 4 2 kg. The heavy Rover has the same characteristics as the areal Rover, but more equipment and capability

kg and the Rover range to 100 km. increases the heavy Rover mass to 1,500 W

The Sample Return Capsule (SRC) returns the sample for Earth recovery and has three different return options. One SRC option enters Earth atmosphere similar to the Apollo command module and, by using its two parachutes, is air-snatched by aircraft. Another SRC option either aerocaptures or propulsively captures into Earth orbit and is retrieved by the Space Shuttle or Space Station OMV. The on-orbit SRCs are sized to loiter in Earth orbit for three months by using rechargeable batteries and solar panels.

~ ~ ~~

~ ~~~~.~~~

The major MRSR elements such as the Mars Ascent Vehicle (MAV), Rover, Earth Return Vehicle/Sample Return Capsule/ Rendezvous Docking Module (ERV/SRC/RDM) segment and Mapping/Communications Orbiter (MCO) can be divided among two Titan IV launch vehicles to emphasize different design philosophies. For example, element grouping can assure partial mission success even if one launch fails, or international

The United States has explored Mars participation without technology with unmanned probes in search of life. transfer. Of these possible element Although no life was found, a manned grouping and capabilities, three exploration of the red planet is an missions were considered optimum and ultimate goal and challenge. The MRSR will be discussed in this paper program is a precursor mission that will (fig. 1).

* AIAA member Copyright 0 1989 by the American l n~ t i i u t~ of Aeronautics and AslrOnaUtiCS, 1°C. No copyright is asserted in the

United Starer under Title 17. U.S. Code. The U.S. Govern- ment has a royalty-free license to ercreisc all rights under rhc copyright claimed herein for Goveinmenml purposes.

All other rights are reserved by the copyright owner. 1

The three MRSR missions show variations in element grouping and capability. Areal B mission uses two Titan IV first stages with the Centaur G' upper stage to launch the

Areal. The first Titan IV launches Configuration 82 Areal in 1998 with the Rover and Mapping/Communications Orbiter (MCO) aerocapturing into Mars orbit (Fig. 2 ) . Once in orbit, the MCO maps a landing site. Then the Rover, using the aeroshell, enters and lands on Mars (Fig. 3 ) . The second Titan IV launches Configuration Bl-A/D in 1998 with the Mars Ascent vehicle (MAVl and the direct

- configurations known as Bl-A/D and B2

entry ERV/SRC/RDM aerocapturing into Mars orbit (Fig. 4 ) . The ERV/SRC/RDM seqment is iettisoned from the aeroshell to-remain in Mars orbit while the MAV uses the aeroshell to land on Mars (Fig. 5 ) . The mission sequence continues with the samples being gathered by the Rover and transferred to the MAV. The MAV is launched from the Mars surface to rendezvous and dock with the ERV/SRC/RDM segment. The MAV transfers the samples to the ERV/SRC/RDM. The ERV/SRC then leaves the RDM and MAV in Mars orbit and returns to Earth. The direct entry SRC is shown in Fig. 6.

One main advantage of the Areal 8 mission is that the element groupings allow for international cooperation without technology transfer. Also, partial mission success is guaranteed with the loss of one of the configurations. The main disadvantage is that the Rover must rendezvous with the MAV on the Mars surface since these elements are landed separately. Element grouping for Configuration Bl-A/D makes meeting the cg requirements for both Mars aerocapture and entry with the same aeroshell a challenge due to the jettison of the heavy ERV/SRC/RDM segment in Mars orbit prior to entry.

The Areal D and Local D missions examined during the MRSR study have the same element groupings but different Rover and SRC designs. Both of these missions use two Titan Iv launch vehicles with a Centaur G' upper stage.

The Areal D mission consists of a Titan IV launch in 1996 with Configuration D2-P/P containing the MCO and the ERV/SRC/RDM that is propulsively captured at Mars. The SRC is propulsively captured at Earth. The second Titan IV launch in 1998 is Configuration D1 Areal with the MAV and areal Rover being aerocaptured at Mars (Fig. 7). Both elements land together on the Martian surface in a horizontal position (Fig. 8 ) .

The Local D mission is different from the Areal D mission in that the SRC in Configuration DZ-P/A is aerocaptured at Earth and the Rover in Configuration D1 Local is the small, less capable

local ranging Rover. The aerocapture SRC is shown in Fig. 9. Configuration D1 Local is shown in Fig. 10. Due to the size of the local Rover, the elements land together using a vertical lander similar to Configurations Bl-A/D and 82 Areal. Once on the Mars Surface, the mission sequence is similar as described for the Areal B mission.

There are several advantages and disadvantages to the two D missions. An advantage to both missions is a greater probability of transferring the Rover samples to the MAV since the Rover and MAV land together. Both elements that remain in Mars orbit are launched together so that no complicated on-orbit deployment is required. Obtaining the required cg for Mars aerocapture and entry is simpler since no elements are jettisoned after aerocapture. Propulsive capture at Mars allows the D2 configurations to be launched two years before the D1 configurations which depends on aerocapture to achieve a Mars orbit. This provides the MCO two years to locate and map a suitable landing site.

A major disadvantage of the Areal D mission is the additional complexity involved in landing the MAV and Rover together onto the Martian surface. A complex horizontal lander is designed for D1 Areal to accommodate the payload size and mass. Landing the MAV in a horizontal position requires a rotation mechanism and support structure to place the MAV in a vertical attitude for Mars launch.

The disadvantage of the Local D mission is the lack of canabilitv and range inherent in the design and-size of the local Rover.

Aeroshell Design

the MRSR elements to Mars was determined by several factors. The aeroshell had to fit inside the Titan IV payload fairing's dynamic envelope (Fig. 11) to satisfy launch load requirements. The payload fairing's dynamic envelope has a nose cap length of 3.81 m and extends to 19.05 m using up to four 3.05 m segments. The envelope's diameter is 4.57 m. Based on these envelope restrictions the aeroshell could not be longer than 19.05 m OK greater than 4 . 5 1 m in diameter.

The aeroshell vehicle that carries

The MRSR vehicle elements do not need the volume that a 19.05 m length aeroshell provides, so the length considered for the aeroshell is considerably shorter. The limiting factor in sizing the aeroshell is the Titan IV and Centaur G ' launch mass capability. The MRSR study is restricted to a total launch mass of 1,000 kg.

2

Atmospheric data obtained from Viking and other previous Mars missions allow for the design of aerocapture vehicles that use the Martian atmosphere instead of propulsion for deceleration and orbit insertion. The shape of the Titan IV payload fairing prompted the use of a bicpnic aeroshell for the MRSR mission. ESS presents the results of aerodynamic analysis of various biconics. The analysis indicated that the biconics providing satisfactory L/D capability and adequate lateral stability would be considered for the aeroshell shape.

Trade studies varying nose radius, total vehicle length and aft cone angles determined the vehicle shape and volume for aerocapture and element packaging. The aft cone base diameter was set at 4.57 m and the aeroshell length was set at 9.09 m. Aerodynamic results show that the L/D decreases with increasing nose radius (Fig. 12) and decreasing aeroshell length (Fig. 13). Increasing aft cone angle increased L/D but decreased available volume for packaging since the aeroshell length remained constant (Fig. 14).

ments are to provide sufficient L/D for Mars aerocapture and entry based on the expected navigation, guidance and atmosoheric disoersions. ESS indicates

The primary aerodynamic require-

that H biconic kith L/D of approximately 1.0 is preferred. If further study indicates that a lower L/D is adequate for aerocapture and entry then an aeroshell with a larger nose radius may be considered to alleviate forward packaging problems. Also, a larger trim angle of attack could be flown which reduces the ballistic coefficient, aerodynamic heating and allows a more aft cg location.

the MRSR study is shown in Fig. 15. The mid L/D aeroshell (known as Configuration M biconic) is a symmetric biconic with a length of 9.906 m, a nose radius of 0.305 m, a forward cone angle of 23.9O and an aft cone angle of 4'. The forward cone base diameter is 3.12 m while the aft cone base diameter is 4.51 m. As the study matured, it became possible to cut off aft sections of the aeroshell with a corresponding mass reduction of 176 kg/m.

performance variation with cg. Based on a desired L/D of 1.0 with an angle of attack of 27.0' at Mars aerocapture and entry, the aeroshell cg is located at 53.4% of !he aeroshell as measured from the nose. Table 1 shows the mass and cg for a L/D of 1.0 at 27.O0 angle of attack for various aeroshell lengths. Reductions in the aeroshell length decreases mass and shifts the cg aft, but also decreases lateral stability

The aeroshell vehicle baselined for

Fig. 16 shows the biconic aeroshell

ma rg i ns . Configuration Packaging

W' The placement of the cg to obtain

the desired L/D and trim angle is a major concern for packaging the MRSR elements. The primary problem is limited volume in the aeroshell nose and heavy MRSR elements placed in the aft end of the aeroshell. Elements such as the MCO and ERV/SRC/RDM segment limit packaging flexibility. These elements are placed in the aft end of the aeroshell so they can be deployed after aerocapture into Mars orbit. Since the MAV and Rover land on the Mars surface, they need to be placed as far forward in the aeroshell nose as possible. The MAV's aerodynamic shape uses the aeroshell forward nose section volume effectively.

The aeroshell propellant tanks are used to obtain the desired cg for each Configuration. They are located as far forward in the aeroshell as practicable. In B2 Areal they are placed in the aeroshell nose section because the Rover does not completely fill that section. The propellant tanks are almost full during aerocapture and help to balance the cg with the MCO in the aft end. The B2 mission requires a large A'? from the aeroshell following aerocapture and most of the'propellant is used prior to entry after the MCO has deployed. Additional complexity is added to the aeroshell propulsion system since the tanks are separated from the engines.

The MCO is packaged compactly inside the aft end of the aeroshell in B2 Areal. It's eventual design may require the aeroshell to be extended. With the legs packaged as in Fig. 2 the aft aeroshell must be retained for thermal protection during entry. This requires the MCO to perform imaging through a window or door in the top portion of the aeroshell since it remains attached to the aeroshell until right before entry of the Rover. The MCO is required to perform the optical navigation during Mars approach and surface imaging in the 1 Sol orbit prior to transferring to the 1/3 sol orbit. Surface imaging while still attached to the aeroshell may require special considerations in the MCO and aeroshell structural interface.

v

The lander structure incorporates a portion of the aeroshell structure to decrease mass and increase volume efficiency. A structural analysis is needed to verify that the option is viable and can save mass when compared to a more conventional lander that jettisons the aeroshell before landing. Disadvantages to this design is a

entire aeroshell structure. And, if the aeroshell lands in a horizontal

propellant mass increase to land the - 3

attitude, as in D1 Areal, the aeroshell must be able to withstand both horizontal and vertical structural loads.

I

Configuration Packaging Design Results

In the three MRSR missions, there are four configurations that are aerocaptured at Mars. These configurations, B1-A/D, B2 Areal, D1 Areal and D1 Local are affected by the aeroshell packaging and cg constraints. The design results for these four configurations include subsystem definition and design, cg identification for Mars aerocapture and entry, and mass estimates.

Subsystem Description

All of the aerocapture configurations have a number of similar subsystems. Each configuration uses the Configuration M biconic as the aeroshell vehicle. Aeroshell structusal and TPS mass are based on 12.2 kg/m of aeroshell surface area. This mass relationship was developed by using historical mass data from previous aerospace programs. Based on packaging volume requirements, the configurations used the biconic M aeroshell that was either 8.08 m or 8 . 6 9 m in length.

each lander is similar. The system is a pressure-fed, bi-propellant system with multiple engines that delivers a Mars T/W of 3 for deceleration and maneuvering. A bi-propellant system was selected over a mono-propellant system for two reasons. A bi-propellant system has better performance and the MRSR lander is anticipating less planetary protection restrictions than were required for the Viking lander. Multiple engines provide flexibility for attitude control by pulsing the engines and reducing surface erosion when compared to a large single engine. Based on the lander payload, D1 Areal uses 1 2 lander engines, D1 Local uses 9 engines, 81-A/D uses 8 engines, and 82 Areal uses 6 primary propulsion engines.

The lander structure is a conventional legged structure that uses crushable honeycomb struts to absorb energy. The lander primary structural mass is based on historical mass relationships using Apollo and Viking mass data. All the landers use a portion of the aeroshell structure as lander structure. B2 Areal is designed with three legs due to the complex Rover deployment (Fig. 3 ) . One of the lander legs is designed to stroke an additional distance after landing which brings it closer to the surface for Rover _- deployment. A two segment ramp deploys which allows the Rover to roll off the lander. The second ramp segment is folded beneath the first segment. D1

The descent propulsion system for

Local and B1-A/D (Fig. 5 ) have four legged landers for additional stability while launching the MAV. Although D1 Areal launches the MAV, its lander has three legs due to the lander size and problems with packaging the lander legs (Fig. 0 ) .

The lander parachute is a reefed Disk-Gap-Band chute that is packaged in the aeroshell nose for D1 Local, Bl-A/D and BZ Areal. The parachute is packaged above the Rover in D1 Areal due to its horizontal landing. The parachute system mass was determined from mass scaling relationships developed by Pioneer Systems. Portions of the aeroshell are jettisoned during parachute deployment.

The thermal control system rejects excess RTG heat during Earth-Mars cruise phase, Mars aerocapture and entry to keep the aeroshell subsystems cool. A passive thermal control system consisting of a radiator, heat pipes and louvers cool the RTGs. Since the louvers are closed during aerocapture and entry an active thermal control system is required for the RTGs during these mission phases. The preliminary concept uses a water boiler system. The steam is vented to the aeroshell's exterior. The thermal control system mass varies with the number and location of RTGs in the aeroshell.

The aeroshell propulsion system uses 16 Marquardt R-4D engines for AV maneuvers and attitude control. The aft facing engines are canted inward IOo to provide a longer moment arm for pitch and yaw control. The lateral firing jets provide roll, pitch and yaw control and conform to the aeroshell's contours.

Center of Gravity Assessment

The cg for Mars aerocapture and entry was based on the mass estimates for each configuration. D1 Areal and B2 Areal use the 0 . 6 9 m aeroshell so the cg required for Mars aerocapture and entry is 54.2% of the aeroshell length as measured from the nose. This cg produces a L/D of 1.0 at a 2 7 . 0 ° angle of attack.

D1 Areal has a cg of 5 3 . 9 % for aerocapture and 5 3 . 5 % for entry. The cg shift is small because no elements are being left in Mars orbit hefore entry and the aeroshell propellant tanks are close to the cg. Moderate adjustments in subsystem locations can be used to obtain the exact cg requirements.

82 Areal has a cg of 5 5 . 2 % for aerocapture and 5 4 . 4 % for entry. The forward placement of the aeroshell propellant tanks and folding the Rove1 RTG section over the Rover trailer assembly help to shift the cg forward. Since 8 2 Areal requires large aeroshel dv maneuvers, the propellant mass is

4

significant (1,200 kg) and affects the cg during aerocapture and entry. To obtain the required cg the mass of the forward aeroshell propellant tanks and the MCO is balanced dependent on attach- ment of the MCO during aerocapture or deployed in Mars orbit prior to entry.

Bl-A/D and D1 Local have 8.08 m aeroshell length. The cg required for this aeroshell length is 54.8% of the aeroshell as measured from the nose with the same L/D and trim angle as the other aeroshell.

The preliminary design for 81-A/D has a cg of 56.6% for aerocapture and 48.2% for entry. Bl-A/D has the worst problem with cg for aerocapture and entry. The problem is inherent in the element grouping. The MAV with lander subsystems and ERV/SRC/ RDM segment have similar masses. The slightly heavier ERV/SRC/RDM segment is deployed in orbit which shifts the entry cg forward.

There are some possible solutions to Bl-A/D cg problems. For aerocapture the 56.6% cg would trim the aeroshell at a higher angle of attack (40°) with a lower L/D (0.7). This may be an acceptable solution if the lower L/D is adequate for aerocapture. For entry, the aft section of the aeroshell could be jettisoned which shifts the cg aft. But, disadvantages include adding another aeroshell RCS system to replace the jettisoned RCS (more mass and complexity) and a new lander leg design s o they are not exposed during entry.

The D1 Local aerocapture cg is 54.95% while the entry cg is 50.2% of the aeroshell length as measured from the nose. The cg shift is due to using the aeroshell propellant. Jettisoning the aft aeroshell section would shift the cg aft for entry as in B1-A/D. The same disadvantages would apply to D1 Local that are true for Bl-A/D. A more complex solution allows the MAV/Rover/ Lander system to shift in the aeroshell to obtain the desired cg.

Mass Statements

Table 2 shows the mass for each configuration during the mission phases. D1 Local has the lightest launch mass at 5,173 kg and the greatest launch mass margin of 35.3%. 82 Areal is similar with 5,207 kg launch mass (34.4% margin). D1 Areal, with a heavy payload of 2,978 kg, has the heaviest launch mass of 6,481 kg and a mass margin of 8.4% which is below the 20% margin allowed by the MRSR program. 81-A/D is similar with a launch mass of 6,457 kg ( 8 . 4 % margin).

Mass tables were created for each configuration. The first mass statement table summarizes the subsystem mass and defines the dry, inert and gross mass

for each configuration. Detailed mass tables were also generated for the subsystems, including: Structure and TPS, dry Propulsion, Power, Control, Data and Environment. The Other table W includes the parachute, bioshield, launch adapter and unique mechanisms. Tables for Payload, usable and inert propellants complete the statements. The detail mass statement for B1 A/D is shown in Table 3 and represents all the configurations. 82 Areal (Table 4). D1 Areal (Table 5) and D1 Local (Table 6) summary mass statement are included. As the study progresses the configuration mass will be defined to a greater level of detail.

Conclusion

Four MRSR configurations were designed and packaged in the configuration M biconic aeroshell that has a L/D of 1.0 with a 27.O0 angle of attack during Mars aerocapture and entry. The major MRSR element designs, constraints and functions predetermined their placement.in the aeroshell which created a challenge in obtaining the cg required for aerocapture and entry. The biconic M aeroshell could be modified to improve cg requirements and by increasing the nose radius although a reduction in L/D does result. The mass can be reduced by more efficient element design and packaging to reduce the aeroshell length. To accomplish this,

minimizing the element volume. emphasis needs to be placed on v

References

ESS, Robert H. and Munday, Stephen R.: "Aerodynamic Requirements for a Mars Rover Return Aerocapture Vehicle,'' AIAA Aerospace Sciences Meeting, Jan. 1989. AIAA Paper No 89-0630.

5

LANDER

FIG1

D l - A

m /M\

B 1 - A / D

piZG-j AEROSHFLL

LOCAL D AREAL B

JRE 1: MRSR MISSION SUMMARY

ROVER

D 1 - A

LANDER

AREAL 0

ORBITER

AEROSHELLTA

METER SCALE AEROSHELL

FIGURE 2: CONFIGURATION 82 AREAL SIDE VIEW

FIGURE 3: CONFIGURATION 82 AREAL DEPLOYED LANDER

6

LANDER

ERVISRCIRDM

FIGURE 4: CONFIGURATION B1 N D SIDE VIEW

FIGURE 5: CONFIGURATION 61 DEPLOYED LANDER

PARACHUTES

METER SCALE

0 1

FIGURE 6: DIRECT ENTRY SRC SIDE VIEW

1

FIGURE 7: CONFIGURATION D1 AREAL SIDE VIEW

FIGURE 8: CONF. D1 AREAL DEPLOYED LANDER

FIGURE 9: AEROCAPTURE SRC SIDE VIEW

METER SCALE

Q i 2 3 4 FIGURE 10: CONFIGURATION Di LOCAL SIDE VIEW

AEROSHELL

12.2 (40.0)

(10.0) 18.25 (60.0)

UNITS: METERS (FEET)

FIGURE 11: TITAN IV PAYLOAD FAIRING

LID

ANGLE OF ATTACK (DEG)

FIGURE 12: NOSE RADIUS VARIATIONS ON U D

9

1.0

0.8

0.8

VD

0.4

0.:

0.'

- RNs0.3 m. THETA2=O deg

- L1=12.20 m

----- LI=Q.I~ rn

- - - - e L1=6.10 m

FIGURE 13: VEHICLE LENGTH VARIATIONS ON UD

Length (Lg) = 32.5 11.

UD VS THETA2 Volume vs. Theta 2

1.6 -

0.6 . '

0 2 4 6 8 IO 1 2

P 4000- +,

E a,

z

> -

3000 -

2 o o o i . . , . . , . . I . . , . . I . . 0 2 4 6 8 10 1 2

-- FIGURE 14: HALF CONE ANGLE (THETA 2) VARIATIONS ON L/D AND VOLUME

10

FIGURE 15: CONFIGURRATION M AEROSHELL

Ealllfllc Coelllclsnl Trlm Angle 00 60

50

40

ua 30

axr 7ca em

W I C M [*v.w n

am 10

m I W IO

0.50 0,sa 0.64 0.66 0 . u 0.10 m

LlH lo Drag Rallo

0.50 0.52 0.54 0.68 0.58 0.60 w

Moment Coelllclent per Trlm Angle

t .2 0.W

1.1

1.0

0.S

0.0

0.7

0.8

0.5 4.m

4.01

M Cmn

4.02

9.50 0.52 0.54 0.56 0.58 0.60 0 50 0.52 0.54 0.511 0.58 0 60 xron w

FIGURE 16: CONF M AERODYNAMIC PERFORMANCE

11

CONFIGURATIOI

6.08 4.43

7.47 4.17

6.66 3.92

6.25 3.67

5.M

1 BlCONlC AERO:

% BODYLENGTH

FROMNOSE

53.4%

53.72%

54.22%

54.6%

55.84%

57.12%

56.89%

61.36%

.LL

MASS.KG

1.296

1,190

1,066

984

863

766

690

595

TABLE 1: CG AND AEROSHELL STRUCTURE MASS FOR CONFIGURATION M BlCONlC

AEROCAPTURE MASS SUMMARY

SEGMENT MASS, KG D1 E1 0 1 02 LOCAL N D AREAL H E A W

LANDED PAYLOAD 2151 1662 2976 1620

LANDED MASS 3002 2519 3629 2160

MASS AT CHUTE 4216 3787 5\27 3761 DEPLOY

ENTRY MASS 4442 4049 5566 4313

INJECTED MASS 4955 6215 6208 6219

LAUNCH MASS 5173 6457 6461 6467

TABLE 2: SUMMARY OF CONFIGURATION MASS

02 81 AREAL N P 1162

1978

2854

3356

SOW

5201

1662

2519

3789

4051

6349

6595

CONFIGURATION 81 N D MASS STATEMENT

TABLE 3 CONFIGURATION 81 ND MASS STATEMENT

12

STAR TPACKER (2)

TABLE 3: CONF €31 NO MASS STATEMENT (CONT)

C’

13

OTHER.

PARACHUTE

PAYLOAD I 3.1961

312 312

70 S d t W rrom 70 simeer smsm

MAV

comingenzy Sample Pa+ J Goodlng.JSC I I I I I *I

1.456 N Lance. JSC

TABLE 3: CONF B i AID MASS STATEMENT (CONT)

ERV

14

1.218 N.Lance.JSC 1.218

P A L I O W D E R "En c.=&lEn

C O W E M I S SYBS"lEULCOLPONm7 -,xo - STRUCTURWTPS 1.274 1,066 186 . L'

PROPULSION 469

POWER 48

260 229

CONTROL 44 35 8

DATA 116 116

ENVIRONMENT 63 63 ________I__

-. OTHER 244 244

- Sw"IT€-DRI(""I -.so

IIElA€WrAU03*TaW W W . 8

*sea wmil m"sR ORB8TEO - PROPULSION 514 62 432

I ENVIRONMENT I 51 I I 5 i I I I I

POWER 111 1 1 1

PAYLOAD 1.860 I 1.710 15" I

~uBnllTLyx(xRwwI uu.I.xo COWNIS

STRUCTURWTPS 1.260

PROPULSION 405

TABLE 6: CONFIGURATION D i LOCAL MASS STATEMENT

L Y U M n u l O w I I I i

M O f R W"n) ORBII6R

964 276

185 220 -

1 5

POWER 128 128

CONTROL 44

DATA ,140 140 - OTHER E-. 298 298

DRY MASSTOTAL: 2.326 1,553 773 D 1