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NASA CONTRACTOR REPORT

/NASA CR-66662

i F,._L.EPO.Ti; STUOYOFO,.ECTVE.SUSO.B,TALENT.YFOR . MARS MISSIONS

[_'_ VoJume JV - Appendix B- Entry and TerminaJ Phase

_t_i1, _ Performance Analysis

Prepared by

MARTIN MARt ETTA CORPORATION

DENVER, COLORADO

_;.:: for

:,_ _ Langley Research Center

NATIONAL AERONATUICS AND SPACE ADMINISTRATION * WASHINGTON, D.C. AUGUST 1968

https://ntrs.nasa.gov/search.jsp?R=19680022362 2018-07-17T08:29:53+00:00Z

NASA CR-66662

FINAL REPORT

STUDY OF DIRECT VERSUS ORBITAL ENTRY FOR MARS MISSIONS

VOLUME IV: APPENDIX B - ENTRY AND TERMINAL PHASE

PERFORMANCE ANALYSIS

By Charles E. French and Duane A. Howard

Distribution of this report is provided in the

interest of information exchange. Responsibil-

ity for the contents resides in the author or

organization that prepared it.

Prepared under Contract No. NASI-7976 by

MARTIN MARIETTA CORPORATION

Denver, Colorado

for

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

FOREWORD

This Final Report for the "Study of Direct Versus Orbital

Entry for Mars Missions" (NASA Contract NASI-7976) is provided

in accordance with Part III A.4 of the contract schedule as

amended. The report is in six volumes as follows:

NASA CR-66659 - Volume I - Summary;

NASA CR-66660 - Volume II - Parametric Studies, Final Analyses,

and Conceptual Designs;

NASA CR-66661 - Volume III - Appendix A - Launch Vehicle

Performance and Flight Mechanics;

NASA CR-66662 - Volume IV - Appendix B - Entry and Terminal

Phase Performance Analysis;

NASA CR-66663 - Volume V - Appendix C - Entry Configuration

Analysis;

NASA CR-66664 - Volume Vl - Appendix D - Subsystem Studies

and Parametric Data.

ii

APPENDIX B

CONTENT S

pa_eAPPENDIX B -- ENTRY AND TERMINAL PHASE PERFORMANCE

ANALYSIS ........................ 1

I. ENTRY TRAJECTORY ANALYSIS ............. 1

2. TERMINAL PHASE SYSTEM ANALYSIS .......... 77thru

312

B1

B2

B3

B4

thru

BI8

BI9

thru

B33

Normalized Axial Force Coefficient versus Mach

Number ...................... 3

Mars Model Atmospheres .............. 5

Ambient Temperature and Speed of Sound versus

Altitude .................... 6

Time from Entry to 20 000-ft Altitude ....... _thru

2324

Downrange Angle Entry to 20 000-ft Altitude thru

1thru Peak Drag Acceleration ...............

B43

B44 }thru Velocity at 20 000-ft Altitude ...........

B58

B59 }thru Altitude at Mach 2.0, 3.0, 5.0 ..........

B70

B71 Mars Skipout Boundary ...............

B72 Propellant and Thrust/Weight as a Function of

Parachute Terminal Velocity ............

38

39

thru

48

49

thru

63

64

thru

75

79

83

B731 184thru Aerodecelerstor Size Limit ............ thru

B78 89

B79 Aeroshell Ballistic Coefficient .......... 91

B80 Minimum Time on Parachute versus Altitude Change 92

B81 } I 94thru Aerodecelerator Size Limit ............ thru

B92 _ 105

B93 Entry Ballistic Coefficient ............ 115

thru Terminal Phase System Performance ......... thru

B194 _ 219

B195 Ablator Weight Discontinuity Example ....... 220

iii

APPENDIXB

Pa_eB196 _ { 221

thru _ Landed Equipment Weight _thruB200 _ 224

B201 Aerodecelerometer Performance, Entry from Orbit 226

B202 Aerodecelerometer Performance, Entry from Orbit 227

B203thru_ WLE and WE versus Aeroshell Diameter for Orbit _thru{228

B216 _ Mode Aerodecelerators ............... _ 241

B217thru__ WLE and W E versus Aeroshell Diameter for Direct _thru_242B232 Mode ....................... _ 257

B233 Effect of Terrain Height ............. 259

B234 Effect of Terrain Height ............. 260

B235 Effect of Entry Velocity ............. 261

B236 Retrodecelerator, Entry from Orbit ........ 262

B237 Retrodecelerator, Entry from Orbit ........ 263

B238 _ _ 264

thru_ Retrodecelerator, Direct Entry .......... _thruB241 _ 267

B242 _ _ 268

thru I Retrothrust/Initial Mass ............. _thruB245 _ 271B246 Comparison of Flaps and Inflatable Afterbodies 272

B247 _ _ 274

thrul WLE versus W E with and without Margin ....... {thruB256 _ 283

B257 _ _ 284

thru _ Maximum I0% Margin .............. _thruB260 WLE' _ 287

B261

B262

B263

B264

B265

B266

B267

B268

B269

B270

B27!

Comparison of Monopropellant and Bipropellant

Verniers ..................... 288

Maximum WLE Envelope ............... 289

Aerodecelerator WLE Envelope ........... 290

Retro WLE Envelope ................ 291

Retro WLE Envelope ................ 292

Retro/Parachute/Vernier System .......... 293

Two-Stage Retro Performance ............ 294

Effect of Deployment Mach Number ......... 296

Effect of Deployment Mach Number ......... 297

Terminal Phase Summary, Aeroshell Diameter =

8.5 ft ...................... 298

Terminal Phase Summary, Aeroshell Diameter =

15 ft ....................... 299

iv

APPENDIX B

B272

B273

B274

B275

B276

B277

B278

B279

P_geTerminal Phase Summary, Landed Equipment Weight

= 600 ib ..................... 301

Terminal Phase Summary, Landed Equipment Weight

= 600 ib ..................... 302

Terminal Phase Summary, Entry Weight = 1500 Ib 304

Terminal Phase Summary, Entry Weight = 1500 ib 305

Performance Summary, Aeroshell Diameter = 8.5 ft 307

Performance Summary, Aeroshell Diameter = 15 ft 309

Performance Summary, WLE = 600 Ib ......... 310

Performance Summary, WLE = 1500 Ib ........ 311

V

APPENDIX B

ENTRY AND TERMINAL

PERFORMANCE ANALYSIS

APPENDIX B

The entry trajectory includes that portion of the mission

from atmospheric entry (800 000 ft altitude) to terminal phase

initiation, occurring in the vicinity of 20 000 ft altitude.

The terminal phase is that portion of the trajectory from entry

trajectory termination to touchdown on the surface. The objec-

tives of the entry and terminal phase performance analysis are:

i) Determine entry trajectory characteristics to define

the entry environment and terminal phase initial

conditions for the two mission modes;

2) Compare mission modes for several terminal phase de-

celeration systems;

3) Compare terminal phase deceleration system performance

for each mission mode.

The entry trajectory analysis assumes a 70 ° half-angle cone

aeroshell CD = 1.64 and the VM-I thru VM-8 range of Martian

atmosphere models. Entry variables include velocity, ballistic

coefficient, and flightpath angle. The environmental character-

istics during entry (dynamic pressure, acceleration, etc.) pro-

vide data for aeroshell and subsystem analysis and design. The

trajectory characteristics (velocity, dynamic pressure, Mach

number, flightpath angle, altitude) provide initial conditions

for the terminal phase analysis.

The terminal phase system provides deceleration from velocities

at the end of the entry phase to zero velocity at the Martian sur-

face. The types of decelerator systems studied generally include

two basic types -- aerodynamic with retro-vernier and all-retro.

Performance is measured by variables such as aeroshell diameter

required, flightpath angle that can be tolerated, landed weight,

and entry weight. A number of constraints are imposed; these

will be discussed in the appropriate section. Required entry

weights are compared with launch vehicle capability (Appendix A).

Limiting entry flightpath angles are correlated with the entry

corridors resulting from the error analysis (also from Appendix

A).

i. ENTRY TRAJECTORY ANALYSIS

• =_=_L=UL_U LL=_ =LLL_7 LL=jectory u=L= are _Lese[[Leu iH L[l±s

section of Appendix B. Entry data are presented in the form of

carpet plots at a constant value of entry velocity with the

APPENDIX B

dependent variable a function of entry flightpath angle and entry

ballistic coefficient. A range of entry velocities covering the

orbital and direct entry modes is presented. These summary plots

are derived from a parametric matrix of entry trajectory calcula-

tions performed on the CDC 6400 computer. The simulation model

is a point mass program and, therefore, does not include the ef-

fects of vehicle dynamics. For the small angles of attack at

entry achieved with an active attitude control system, the point

mass simulation provides entirely satisfactory data for the entry

environment parameters presented herein.

Entry vehicle drag coefficient used in the trajectory calcula-

tions is a function of Mach number. The normalized drag coeffi-

cient as a function of Mach number is shown in figure BI. Indi-

cated on the curve is the hypersonic drag coefficient level for

a 60 and 70 ° half-angle cone.

V E = entry velocity (inertial) at entry altitude (800 000

ft);

7E = entry flightpath angle (inertial) at entry altitude

(negative downward);

B E = entry vehicle hypersonic ballistic coefficient,

m/CDA, slug/ft';

C A = axial force coefficient of entry vehicle;

A = reference area, plan area of entry vehicle.

Velocity and flightpath angles at initial entry conditions

are defined in an inertial system with the coordinate system

origin at the center of Mars. Entry trajectories consider the

rotation of the planet and are computed in a coordinate system

that _otates with the planet. Velocities and angles during the

entry trajectories arc therefore relative to the atmosphere, which

is considered to rotate with the planet. The trajectories from

which the data presented in this report are taken, enter in the

equatorial plane with an east (posigrade) heading.

The following planetary characteristics are used in the entry

trajectory calculations:

l) Planet radius, 3393 km (ll 131 754 ft);

2) Gravitational constant, 1.51254 x i015 ft3/sec°;

3) Rotation rat:c, 7.088]33 x lO -5 tad/see;

4) Atmosphere models, VM-I, -2, -3, -4, -7, and -8.

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The atmosphere models are the VM atmosphere as specified in

RFP-VO-6-4509. Important characteristics are tabulated below.

Model VM-I VM-2 VM-3 VM-4 VM-7 VM-8

Surface pressure, mb 7.0 7.0 i0.0 i0.0 5.0 5.0

Surface density, slug/ft 3 x I05 1.85 3.59 2.65 4.98 1.32 2.56

Surface temperature, °R 495 360 495 360 495 360

Tropopause altitude, ft _ I0 -_ 63.3 61.0 63.3 56.1 63.3 61.0

Stratosphere temperature, °R 360 180 360 180 360 180

Composition, % (by mass)

CO e 28.2 i00.0 28.2 70.0 28.2 i00.0

Ne 71.8 0 71.8 0 71.8 0

A 0 0 0 30.0 0 0

Molecular weight 31.2 44.0 31.2 42.7 31.2 44.0

Specific heat ratio 1.38 1.37 1.38 1.43 1.38 1.37

Inverse scale height, lO-5/ft 2.15 6.07 2.15 5.89 2.15 6.07

Figures B2 and B3 present density versus altitude and tempera-

ture and speed of sound versus altitude for the six mode] atmos-

pheres. Reference to figure B2 shows that VM-8, VM-7, and VM-3

bracket the range of atmospheric densities and upper atmosphere

density scale heights for the six models described above. VM-8

bounds the low scale height for the upper atmospilere and has the

lowest density down to approximately 45 000 ft where VM-7 repre-

sents the minimum density down to mean surface level. VM-3 pro-

vides the highest stratosphere density as well as the highest

density scale height. This reasoning, together with a limited

number of entry trajectory runs in all six atmospheres, has in-

dicated that over the range of ballistic coefficients and entry

flightpath angles studied in the report, atmosphere VM-8, -7, and

-3 represent limiting entry environment and provide terminal phase

critical design conditions. Parametric data presented in this

report are therefore limited to VM-3, -7, and -8. Speed of sound

in VM-8, figure B3, is lowest of all models. This characteristic,

in combination with the low stratosphere density and low density

scale height, makes VM-8 critical for Math number-dependent func-

tions for terminal phase system design. For this reason, par_unetric

data presenting altitude for Mach 2, 3, and 5 are shown only lot

the critical VM-8 atmosphere.

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Parametric data presented are for the conditions tabulated

below.

V_E, fPs BE_ slug/fte ___Z_, deg Atmospheric model

14 000 0.I, 0.2, 0.3, 0.4, 0.6 14, 16, 20, 24 VM-3, -7, -8

16 000 0.i, 0.2, 0.3, 0.4, 0.6 16, 20, 24 VM-3, -7, -8

18 000 0.i, 0.2, 0.3, 0.4, 0.6 20, 30, 40 VM-3, -7, -8

21 000 0.i, 0,2, 0.3, 0.4, 0.6 20, 30, 40 VM-3, -7, -8

24 000 0.i, 0.2, 0.3, 0.4, 0.6 20, 30, 40 VM-3, -7, -8

Time from entry altitude (800 000 ft) to 20 000 ft altitude

is shown as a function of ballistic coefficient, entry flightpath

angle, and entry velocity for VM-3, -7, and -8 in figures B4 thru

BI8. The terminal condition for the entry trajectory is taken as

20 000 ft as it is a representative altitude for initiation of

the terminal phase decelerator system.

Figures BI9 thru B33 present downrange angle as a function of

ballistic coefficient, entry flightpath angle, and entry velocity

for VM-3, -7, and -8. The downrange angle is the central angle

covered from entry to 20 000 ft altitude, the termination of the

basic entry trajectory. The downrange angle shown is calculated

for an equitorial posigrade trajectory (east heading) and includes

the effect of planet rotation.

Figures B34 thru B43 present peak drag deceleration as a func-

tion of ballistic coefficient (where applicable), entry flight-

path angle, and entry velocity for VM-3, -7, and -8. In accord-

ance with the approximate expression,

V 2 sin YE

where h S is the scale heightgmax. = - 2 e g_ h S

(shown in atmoshpere table above),

peak g is not a function of ballistic coefficient for constant

values of scale height. For VM-3 and VM-7 atmospheres, peak g

occurs along the constant scale height region, (i.e., in the

stratosphere). In the case of VM-8, peak deceleration occurs at

altitudes below the tropopause altitude (fig. B2) for the higher

ballistic coefficients and steep entry angles. This leads to peak

drag deceleration with nonlinear characteristics as a function of

ballistic coefficient, as shown in figures B35, B37, B39, B41, and

B43. The increase in scale height below the tropopause accounts

for the reduction in peak G at the higher ballistic coefficients

and steeper entry flightpath angles.

APPENDIXB

Velocity (relative) at the terminal decelerator systemini-tiation altitude of 20 000 ft is presentedas a function of bal-listic coefficient_ entry flightpath angle and entry velocityfor VM-3,-7, and -8 in figures B44 thru B58.

Figures B59 thru B70present the altitude at which Mach2.0,3.0, and 5.0 occur as a function of ballistic coefficient, entryflightpath angle, and entry velocity in VM-8. As discussedearlier, Machnumberat a given altitude for the sameentry con-dition will always be highest in VM-8. Machnumbersensitivefunctions for terminal phasedeceleration systemssuchas para-chute deploymentwill occur at the lowest altitude (critical) inVM-8. Representativedata showingthis are tabulated below.

Entry condition

VE = 14000 fps, BE = 0.30, 7E= -20°

VE = 21 000 fps, BE = 0.30, 7E = -20°

Altitude for Mach 2.0 a ftVM-3 VM-7 VM-8

73 250 41 580 19 000

78 600 48 000 24 000

As mentioned above, the entry trajectory data are used to

define terminal phase initial conditions. In addition, the entry

environment data (peak g, time, etc.) are factored into subsystem

design studies.

APPENDIXB

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APPENDIXB

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APPENDIX B

This page intentionally left blank.

76

APPENDIXB

2. TERMINALPHASESYSTEMANALYSIS

Objectives

Theterminal phasesystemdecelerates the capsule from veloc-ities at the end of the entry phase(approximately 20 O00-ftaltitude) to zero velocity at the Martian surface. The objec-tives of the terminal phaseanalysis are to evaluate the effectof mission mode(i.e., direct mode,entry from the approachtrajectory; and orbit mode,entry from orbit) on terminal phaseperformanceand to compareseveral candidate systemsfor eachmission mode. The terminal phasesystemsconsidered include sub-sonic-type parachuteplus vernier, tuckbackballute plus vernier,all-retro propulsion landing system, and two-burn systems, makinguse of a solid rocket motor for braking before parachute or all-retro systemdeployment.

Terminal phaseinitial conditions are determinedby entrytrajectory characteristics and are related directly to entry con-ditions (velocity, flightpath angle, and ballistic coefficient).Entry corridors (flightpath angle vs. velocity at entry) resultfrom the targeting and error analysis (AppendixA). Correlationwith these values defines the required values over the parametricrange used in the terminal phaseanalysis. Terminal phaseper-formanceis measuredin terms of entry weight, aeroshell diameter,anduseful landedweight. Entry weight, converted to capsule sys-temweight, allows correlation with launch vehicle capability(AppendixA). Thus, the endpoint of the terminal phaseanalysisdefines useful landedweight as a function of aeroshell diameter,entry conditions anddispersions, and required launch vehicle.Sensitivity to design parameterssuchas landed terrain heightand decelerator size is also factored into the analysis alongwith the constraints underwhich the systemis assumedto perform.

SystemDescription, GroundRules and Constraints

The terminal phasesystemsconsidered in this study are:

i) Mach2.0 deployedsubsonic-type parachutewith mono-propellant or bipropellant vernier landing rocketmotors;

2) _,_L_-_3.n_a_A_._......._ n d_plny_ tuckback ballutes with mono-propellant vernier motors;

77

APPENDIXB

3) All-retro propulsion bipropellant decelerator andvernier system (three-engine arrangement);

4) Two-burnsystememployinga high-thrust solid rocketmotor in front of either the MD = 2.0 parachuteor all-retro system.

The two-burn systemconcept is investigated for the direct modeonly; the other systemsare investigated for both direct andorbit modes.

Groundrules and constraints related to the entry trajectoryanalysis are as follows:

i) Entry velocities of 4.5 km/sec (14 800 fps, orbitmode)and 18 000 to 24 000 fps (direct mode);

2) Entry flightpath angles of -16 to -20° (orbit mode)

and -26 to -38 ° (direct mode).

Entry ballistic coefficients are determined as a function of

terminal phase system limitations and are discussed separately

below. The entry trajectory analysis also defines the attitude

of deployment of the aerodecelerator terminal phase systems. The

VM-8 atmosphere is critical in defining the Mach number-sensitive

deployment conditions because of its lowest upper altitude density

(above 44 000 ft results in highest velocities) and its low speed

of sound.

Before discussing terminal phase, a word of explanation is

required relative to the philosophy adopted for defining the entry

flightpath angle versus entry velocity profile to be used for the

deorbit/ejection maneuver strategy. The analysis performed dur-

ing our Voyager Phase B effort indicated the sensitivity of landed

payload to entry flightpath angle. To avoid the problem of analyz-

ing terminal phase system performance as a function of entry veloc-

ity and entry flightpath angle as well as system parameters, the

maneuver strategy was devised, which desensitized the terminal

phase system initial conditions to V E and IE" An example is

shown in figure B71, which shows that there are VE, _E contours

that result in essentially identical flight conditions below 30 000

ft. All of the targeting analysis discussed in Appendix A are

based on V-y contours that have these characteristics.

78

APPENDIXB

oq

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79

APPENDIX B

80

The point is essential here to relate quoted entry flightpath

angles in preceding sections to those that appear below. For ex-

ample, the entry corridors defined above for the two orbits con-

sidered defined the maximum 7E = 16.2 ° and 18.1 ° for the

lO00x15 000- and i000x33 070-km orbits, respectively. These

FE occur at entry velocities of approximately 14 400 and 15 i00

fps, respectively. The terminal phase system analysis was based

on an entry velocity of 4.5 km/sec (14 764 fps). The 7E to be

used in the terminal phase system comparison is obtained by taking

the above values and following the trend shown in figure B71 from

the actual VE to 4.5 km/sec. Thus, the entry corridor limits

for the two orbits change from -16.2 and -18.1 ° to -16.6 and -17.8 °

as far as the discussion below is concerned.

The analysis is performed for a range of aeroshell diameters

as follows:

I) Orbit mode 6.5, 8.5, 12, and 15 ft;

2) Direct mode 6.5, 8.5, 12, 15, 20, 25, and 30 ft.

The launch vehicle shroud size limitation (16 ft diam) has been

interpreted in this parametric analysis to limit the aeroshell

diameter to approximately 15 ft. Aeroshell diameters greater

than 15 ft are obtained by deploying flaps or extendible after-

bodies.

Additional ground rules common to both orbit and direct mode

analyses are:

i) VM-7 and VM-8 atmospheres;

2) Terrain heights at 0 and 6000 ft above the mean planet

surface;

3) Wind velocity of 220 fps at aerodecelerator separation

or retrosystem ignition.

The VM-7 and VM-8 atmospheres are critical for the terminal

phase analysis. The VM-8 atmosphere determines the critical de-

ployment altitude (Mach number sensitive) for the aerodecelerators.

Either VM-7 or VM-8 atmosphere is critical for aerodecelerator

altitude loss (they size the aerodecelerators). In addition, the

VM-7 and VM-8 atmospheres result in the maximum velocity entering

the terminal phase region. Terrain heights of both zero and some

positive value (6000 ft) are studied because an altitude initiation

trigger must be designed on some altitude other than zero. It is

therefore necessary to know terminal phase performance sensitivity

to terrain height.

APPENDIX B

Additional groundrules and constraints related to the terminalphasedecelerator types are described in the following paragraphs.

Aerodecelerator system. - Ground rules and constraints in-

clude:

l) Aerodecelerator completes its job at 4000 ft above

terrain;

2) Flightpath angle at aerodecelerator separation must

be steeper than -60°;

3) Time on aerodecelerator must be at least 16 sec or

longer;

4) The M D = 2.0 system must be sized to accomplish

2) and 3) above or separate aeroshell, whichever is

larger.

All-retro system. - Ground rules and constraints include:

i) High thrust braking phase to zero horizontal velocity

with thrust vector aligned along velocity vector at

initiation;

2) Vertical drop at one Mars G for 3 sec followed by 3

Mars G braking to zero velocity;

3) Initial thrust for high thrust braking phase provides

an acceleration equal to the drag acceleration (i.e.,

minimum allowable thrust-to-weight ratio).

An introductory system description was given at the beginning

of this section. A more comprehensive description is presented

below.

Vernier System

The aerodynamic decelerators (parachute, ballute) use a retro

propulsion vernier for final deceleration to touchdown at zero

velocity on the surface. Vernier system performance is based on

the following ground rules and assumptions:

Ignition at 4000 ft above the surface;

._i__. at ig _ _ I P_ t_mp_ oarachute terminal

velocity;

Flightpath angle at ignition is -60°;

81

APPENDIXB

4) A 220-fps horizontal velocity due to wind at ignition;5) Aerodecelerator separation at vernier ignition.

A monopropellant three-engine systemis assumedfor the major-ity of the terminal phase analyses. The propellant weight is de-

termined from the normalized propellant shown in figure B72 and

the equation:

F Wnorm v

OW =p I

sp

The thrust-to-weight ratio (T/W) is also shown in figure B72 as

a function of parachute terminal velocity,

2g d BDE C2 =

VT p

Vernier system engine and tankage weights are determined from data

that appear in Appendix D. A bipropellant system is also used to

compare with monopropellant performance for the parachute/orbit

case. The equations and curves discussed above and the data in

Appendix D are also applicable to that configuration. Specific

impulse is 225 sec for the monopropellant system and 285 sec for

the bipropellant.

Parachute/Vernier System

The parachute is deployed either at M = 2.0 in VM-8 or at the

same altitude in VM-7. The parachute is assumed to have ac-

complished its purpose when a relative flightpath angle of -60 °

is reached in the most critical atmosphere. Figures B73 thru

B78 show BDE C or parachute size required to reach these final

conditions for various ?E and BE . For example, in figure

B73 an entry ballistic coefficient, BE = 0.4 slug/ft _ at

_E = -18° and hT = 0 requires a BDE C size of 0.045 slug/ft e.

Any larger BDE C will fail to reach ? = -60 ° at 4000 ft above

terrain.

82

APPENDIXB

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/

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8

5v

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4 m

18i00 200 300 400 500

Parachute terminal velocity, VT, fps

Figure B72.- Propellant and Thrust/Weight as a Function

of Parachute Terminal Velocity

I600

83

APPENDIXB

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84

APPENDIX B

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85

APPENDIX B

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86

APPENDIX B

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87

APPENDIX B

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88

APPENDIX B

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89

APPENDIXB

In landedequipmentweight calculations, the parachute sizeused is alwayssufficient to separate the systemfrom the aero-shell at M = 0.8. Therefore, in those caseswhere the parachutesize required for separation is larger than shownin figures B73thru B78, the size required for separation is used. Therelativeacceleration required to separate from the aeroshell is assumedto be 30 ft/sec e. Therequired parachutesize is determined from:

- 'DEC ___/S_ = 30 ft/sec e

qD [ roDEC mA/S ]

30 = qD BA SBDEC

BA/S is shown in figure B79 as a function of aeroshell weight.

The dotted line in figure B75 shows BDE C required for separa-

tion. Larger BDE C will not separate the aeroshell.

A minimum time on the parachute criterion is also assumed.

There must be 16 sec on the parachute to allow for four opera-

tions. Deployment to aeroshell separation requires 5 to 7 sec.

Three seconds is required to separate the system i00 ft from the

aeroshell, assuring that no more than one radar beam is blocked.

It is planned to have an offset load in the aeroshell to allow

the aeroshell to clear the path of the radar beams. Three seconds

is allowed for TDLR lockup. Another 2 to 3 sec is allowed for

vernier ignition and verification. Figure B80 shows altitude

loss on the parachute versus BDE C. The boundaries shown are for

16, 30, and 50 sec on the parachute. These boundaries were formed

from all the trajectories calculated in this study, both orbit and

direct. An example of the data variation of these points is shown

for the 30-sec boundary. Any point above a boundary will have

times that are equal to or greater than the minimum time indicated.

Three typical contours taken from the B E versus BDE C (figs.

B74 and B76) are superimposed. These contours intersect the

t = 16 sec boundary. At that point, the BE versus BDE C bound-

ary must be modified to follow the t = 16 sec boundary as shown

in figures B74 and B76. These show that for a particular 7E a

larger size parachute must be used for a given BE to obtain at

least 16 sec on the parachute.

90

APPENDIX B

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Figure B79.- Aeroshell Ballistic Coefficient

91

28 x 103

APPE_IX B

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VE - 6.0 km/sec; 7E _ -24°

VE - 6.0 km/sec; mE _ -30* |i

I I I I I I0 .04 .08 .12 o16 .20

BDE C, sl/f t2

Figure B80.- Minimum Time on Parachute versus Altitude Change

92

APPENDIXB

Although a completedetermination of minimumtime limits wasnot calculated for all conditions, a checkof the study resultsindicates that all casesof interest meet the time requirement.

Ballute/Vernier System

Ballute deploymentMachnumbersinvestigated are M= 3 andM= 5 in VM-8and at the samealtitudes in VM-7. A tucked-backballute with a 10%burble fence is used in the study. Nopro-vision is madeto size the ballute to separate from the aeroshellas in the parachutesystem. Plots of BDEC versus BE, similarto those for the parachutesystem, are shownin figures B81thruB92. Final conditions are the sameas for the parachute (at least7 = -60° at 4000 ft aboveterrain).

It maybe seenin comparingMachnumbersof 2, 3, and 5 thatwith increasing deploymentMachnumbers,larger BE and smallerdecelerator sizes are required at the sameentry angle. Ingeneral, the orbit moderequires smaller decelerators than thedirect mode. For the direct modeat high 7E, the limiting entryvelocity is 24 000 fps. For lower 7E, the limiting entry ve-locity is 18 000 fps.

93

APPENDIX B

__+

I.20 L......

.16

._ .12

br.1

.08

.O4

BE, sl/ft 2


94

APPENDIX B

Cq

r-N

b

.08

Note: i. MD = 3.

2. V E = 4.5 km/sec.

3. _ I 6000 ft.


95

APPENDIX B

.2O

.16

4J

U_

b

.08

.O4

0 .1

-?,,

"4

A ,

\

\\

1

l

.2 .3

,-[Note: I. MD = 5.

, 2. VE I 4.5 km/sec.

3. _lO. I--

\

__u

.4 .5 .6 .7

BE , sl/ft 2


96

APPENDIXB

.20

.16

0J

.12

.08

.O4

0

\

Note: I. % - 5.

2. VE I 4.5 km/sec.

3. _ I 6000 ft.• . . #

D-

\

o __

%

.2 .3 .4

BE, .sl/ft e

I...........

.6 .7


97

APPENDIXB

Note: I. MD n 3.

2. VE n 18000 fps.


98

APPENDIX B

.20

.16

MD " 3.

VE I 18 000 fps.

= 6000 ft.

I.....

.12q_

bC_

.08

.04

\.i .2 .3 .4 .5

BE , sl/ft e

.6


99

APPENDIX B

BE , sl/ft 2

Figure B87.- Aerodecelerstor Size Limit

i00

APPENDIX B

4-J

r-q

b

.2O

.16

.12

.08

.04 /

J/

Note: i. M D - 3.2. V E = 24 000 fps. __

.i .2 .3 .4 .5 .6

B E , sl/ft 2


i01

APPENDIX B

.2O

.16

.12

bm .08

.04

Note: I. MD-5.

2. VE - 18 000 fps

3. h_-0.

A \

.I .2 .3 .4 .5 .6

BE , sl/ft e


.7

102

APPENDIXB

b

.2O

.16

.12

.O8

.O4

I

L

7

_L. L

.I

Note: I. MD = 5.

2. VE I 18 000 fps.

3. hT =, 6000 ft.

,\\ ,, \.2 .3 .4 .5 .6

BE , sl/ft e

.7


103

APPENDIX B

,20

.16

C_

_o .12

.O8

.O4

i

\/

_j

\

J

,/,,u'l

/

\

\]

-- -4

\_f.i .2 .3

l[ L

2. V E 24 000 fps. J

3 _ o I

I

.2-

gr'_.,, i

\

•4 .5 .6

_____]

BE , sl/ft e


104

APPENDIXB

Y_

b

.20

.16

.12

.08

.O4

\

\

,\\

f

.I .2 .3

Note:

2 V E 24 000 fps.

3 hT 6000 ft

_J

\\ \I _\.4 .5 .6 .7

BE , sl/ft 2


105

APPENDIXB

Single-Stage Retro

The single-stage retro systemassumesa three-engine liquidbipropellant systemwith a specific impulseof 300 sec. Theaeroshell is assumedto separate at retro initiation, though noattempt wasmadeto size the thrust to ensure this. The thrustto initial massis set equal to the drag acceleration at retroinitiation. This represents the minimumallowable thrust-to-weight ratio. An initiation altitude basedon this thrust levelis found by iteration, which results in zero velocity at theterrain surface.

The flight sequenceused in the calculations consist of threephases-- a braking phaseto removethe horizontal velocity com-ponent, a vertical descent at 1.0 g_ for 3 sec, and a final ver-tical braking at 3.0 g_. Thoughthis sequencewould not be usedin an operational design, it gives performancevalues very closeto a real design.

Two-StageRetro

The two-stageretro systemconsists of a first stage solidrocket motor (SRM)and a secondstage liquid rocket motor (LRM).The aeroshell is assumedto separate at SRMignition. At LRMignition, the SRMassemblyseparated. The secondstage reacheszero velocity at the terrain surface. The solid motor thrustto initial massis assumedequal to the drag acceleration atignition. However,the burntime and solid propellant used mustbe found in combinationwith the liquid rocket motor size inorder to find the maximumlanded equipment.

Retro/Parachute/Vernier

This systemconsists of a solid stage retro, a parachute, anda monopropellantvernier. The parachutedeploymentaltitude ischosensufficiently high to ensure at least p = -60° at 4000 ftaboveterrain (vernier ignition altitude). Theretro maneuveris similar to the solid phaseof the two-stage retro. Thrust toinitial massis equal to drag acceleration at ignition. For thisstudy a BDEC = 0.30 slug/ft _ deployedat 20 700 ft aboveter-raill is assumed. Retro specific impulse is 280 sec.

106

APPENDIXB

Extendible Flaps

Extendible flaps are investigated as a meansof increasingthe aeroshell diameter. Theyare deployedbefore entry. Theirpurpose is the sameas the inflatable afterbody. Flap sizes withdrag areas equal to 20-, 25-, and 30-ft aeroshell diameters areinvestigated. Theyare studied with the parachute, ballute, andsingle-stage retro systemsfor the direct modeonly.

Inflatable Afterbody/Retro

Theinflatable afterbody/retro systemconsists of a tucked-backballute without burble fence deployedbefore entry and asingle-stage liquid retro identical to that described above. Theaeroshell and ballute are assumedto separate from the lander atretro ignition. The retro thrust to initial massis equal to thedrag acceleration at ignition. Theballute is sized for dragareas equivalent to 20-, 25-, and 30-ft diameter aeroshells. Theballute weight is basedon the maximumdynamicpressure encounteredduring the trajectory.

Techniquesand Equations

Sin_le-sta_e retro. - The ground rules and assumptions are as

follows:

i) Flat planet;

2) Constant gravitational acceleration;

3) Constant thrust;

4) Constant thrust direction;

5) Zero aerodynamic forces.

The diagram below shows directions of forces

Z

THorizontal

Surface

X //i/i11/_/ 107

APPENDIXB

The single-stage retro systemflight sequenceconsists of threephases:

I) Phase1 - A braking phase, _ = "7, F/m o ffisystem

drag acceleration at retro initiation, final condi-

tion Xf = o;

2) Phase 2 - Vertical descent at 1.0 gw for 3 sec;

3) Phase 3 - Final vertical braking at 3.0 g_.

Phase i equations are as follows

= V cos 70 + VO O W

Zo = Vo sin 70

C. =g IJ sp

if = 0 = X + cos e in iO

in = (i F t ) Vom C. = - 6_.o j j

F t

m C°o j

X =0O

-VojCj= 1 - e = mass of propellant used for Phase 1

then with

Zf = Z - gt - C sin 0 in (i F t )o j m C = grit

o j

or

(-V°/Cj)= _--_ i - e = Vertical velocity at end of Phase i

f (F/too)

Finally:

hf = h + C sin ?o i - - o/ j _ ½ _ =o -V f

i \gd/ j 1- e o/Cj 1

= altitude at end of Phase i

108

APPENDIXB

Phase2 equations:

hfe = hfl + Zflte = hfl + Zfl (3.0) = altitude at end of Phase2

To get the propellant usedin both Phases2 and 3, considergeneral equations for vertical descent at constant acceleration.

mZ= F - gm

"" FZ = m-- gd = gd (n - i)

Integrating

if = g_ (n - I) t + Zo

and, from

-F -F dmg I =_. =dt

sp ]

dm - F/M

m C.J

dt --n gd

C.J

dt

Integrating

(m)nin m _C.J

m

-P---I m__ i em m

O O

C °

J

= I - e

-3 g_

C°

]= propellant used during Phase 2

109

APPENDIXB

Phase3 equations:

Zf3 = go_ (n - i) t3 + Zo3= gcf (n - I) t3 + Zfl

t3 =-Zf_

gc_(n - i)-Zf_ = burning time for Phase32g_

Then, with

Zf3 = g$ (n - i) tz + Zo

Integrating:

Zf = ½gd (n - i) t_ + Zot3 + h°

or

hf_ = hf2

= + _ g_ (n - i) t_ + Zflt3hf 3 ho 3

_ -fl = altitude at end of Phase 3 with n = 34g

The altitude at the end of the retro maneuver is:

where

• rl Ihf3 = h° + Z fl g_ 1 - e -v 4g_

VO

C.J

The h0

below.is found by iteration as is noted on the figures described

II0

APPENDIXB

= 1 - exp -n gc_t3_7 = 1 - e_p

3 J _2Cj gd_

= propellant used during Phase 3

Note: the m in the previous equations are masses at the begin-o

ning of the particular phase. Correcting to the mass at retro

igntiion, the total propellant used is:

m I )IPTOT 1.0 - e -v exp Cj \2g_ -m

ol

Two-stage retro. - The solid rocket motor phase is similar to

the first phase of the single-state retro. Thrust equals drag

acceleration at ignition and _ = - y . The three phases of the_O

LRM are the same as the single-stage retro.

The final altitude for the SRM is:

Zfs = ZO + _ot ½ gd t2 (Cot sin 7o ) [i -

For the LI_M (first phase):

ZfL l = Zfs + ifst - ½ gd te - (Cj.t sin 7i

i - in °Ll

(m)in i - Ps

m o

d+ in

o)

mp ]in 1 -m _-L

+ ----_ -----°el

iii

APPENDIXB

The final altitude is:

Zf2hf = Zf + 3Zf +

L1 4g_

The liquid propellant weight will nowbe found.

XfL = 0 = if + C. cos Fo in - = 0

S JL OL/

_+ C. cos 70 in - Ps + C.3 s 3L

cos 7o in

%/

C0

-v m- s)o = _ = in - Ps

C. cos 70 C.JL JL

+ inPL

m m

o PsS

where

PL _ o

Solving for MpL, the liquid propellant weight,

e-Vo/C j

C0

Ps

m o

The liquid propellant weight depends on Mps, the solid

propellant weight. Therefore, the optimum combination of solid

and liquid systems is found by varying the solid system and find-

ing the liquid system size that gives the maximum landed weight.

The initiation altitude must be found by an iteration procedure

as mentioned in the previous section.

112

APPENDIXB

Retro/parachute/vernier.- TheMachnumberat parachutedeploy-ment is assumedto be 2.0 in VM-8. The altitude at the end of theretro maneuveris found from:

Vf = 2Vc= 2X speedof sound

mdV= - F - mgd sin 70

dO= (- n - sin 7o) gd

2V =V=Vc o (n + sin 70) g_t

t b =(V° - 2Vc)

g<{(n + sin 70)

or

hf =ho +

with

hf = ho +

Vo (V ° - 2Vc) sin 7o

gd (n + sin 7o )

p

½(V ° - 2Vc) sin 7o

gc_ (n + sin )o)

(V ° - 2Vc) sin 7o

gd (n + sin 7o )½ V + Vc] = altitude at endo

of retro

F__= g I =C-m sp j

m qdt

Integrating

(_o) n g¢_ tIn mf = -- C .

J

where

F- n g_ = constant

dm

113

APPENDIXB

Finally

m_= m expl O I -n (V 0 - 2Vc) I

Cj (n + sin 70 )

Basic Parametric Data

The approach used in this analysis is to compare the useful

payload on the ground for the various systems on the basis of

entry weight and aeroshell diameter. The parameter used to

compare the systems is landed equipment weight, WLE. This

parameter is defined as entry weight, W E , minus the following:

i) Aeroshell weight Function of diameter, ballistic

coefficient, 7E;

2) Total aerodecelerator system - Function of size and

deployment dynamic pressure, Mach number, etc.;

3) Vernier or retro system - Function of thrust level

and propellant loading;

4) Entry thermal control and ACS - Function of diameter

and weight;

5) Landed structure and legs - Function of landed weight;

6) Pyro subsystem, cabling, etc. Constant plus function

of diameter.

Thus, WLE is the effective usable weight on the ground com-

prised of entry G&C, all communications and data handling subsys-

tems, power subsystems, surface thermal control, and surface

science subsystems. The parametric weight equations used for the

delivery system weights (i.e., W E WLE ) are given in section Iof Appendix D.

Data are presented for fixed aeroshell diameters of 6.5, 8.5,

12, and 15 ft. Direct mode analysis also includes aeroshells with

effective diameters of 20, 25, and 30 ft obtained by flaps or

inflatable afterbodies on a 15-ft basic aeroshell. The analysis

uses entry weight as the independent variable. The relationship

between entry weight and B E is shown in figure B93 for refer-_IICe .

114

APPENDIXB

00

,r4U,r4

(13OCD

U.r4.IJU].r4,-4

c_

_J

r-U

I

{D

.,_

115

APPENDIXB

Basic data are presented in figures B94 thru B154for all ofthe systems. Thedata are presentedas landed weight ratios(WLE/WE)so that a consistant set of scales could be used forall figures.

Symbolsused to define the aeroshell diameters on the machineplotted figures (figs. B94thru B154) are defined below.

_mbol Aeroshell diameter_ ft

_- 6.5

X 8.5

12

15

2o

A 25

-F 30

An example is shown in figure BII3.

The landed equipment weight may be found by using the overlay

supplied inside the back cover of this report. Of most interest

on these plots is the point of maximum landed weight ratio and

the point of maximum landed equipment weight. It should be noted

that maximum landed equipment weight always occurs at a higher

entry weight than maximum landed weight ratio. The ballistic

coefficients used in these figures were obtained from the BE

versus BDE C curves shown earlier. Therefore, all points on

the weight curves represent systems with the smallest aerodecelera-

tors that will reach final conditions at 7 = -60° at 4000 ft

above terrain. For the parachute cases, the parachute size used

is always sufficient to separate from the aeroshell. For ballute

cases, the aeroshell remains with the ballute during the aero-

decelerator phase. However, ballute size required for aeroshell

separation has been calculated for the hT = 6000 ft cases. The

boundary where the ballute will not separate is illustrated by

fences beginning with figure BI07. If no boundary is shown in

these figures, the ballute will separate from the aeroshell for

all points.

116

APPENDIXB

For direct entry, a heat shield on the backface of the aero-shell wasemployed. This separatesat aerodecelerator deployment.In the figures for the parachute/vernier, direct modeonly, thelanded equipmentweights mustbe corrected by subtracting thebackfaceweights. Thebackfaceweights are as follows:

i) D = 6.5 ft AWBF= 15 ib;

2) D = 8.5 ft _WBF= 26 ib;

3) D = 12 ft _WBF= 52 ib;

4) D> 15 ft _W = 81 lb.= BF

For aerodecelerator cases, the verniered weight is the entryweight less the following:

i) Weightof aeroshell (seeAppendixD);2) Weightof aerodecelerator, expendable(see AppendixD);

3) Weightof ACSpropellant

WACSE= 0.00007WE DA/S+ 0.00085WE;

4) Weight of science in aeroshell (13 ib).

The landedweight is the verniered weight less the vernierpropellant. The landed equipmentweight is the landedweight lessthe following:

i) Weightof lander structure

WSTR = 20 + K (landed wt)°" 7whereK = i.i (orbit);1.57 (direct)

2) Weight of aerodecelerator, fixed,a) Parachute(see AppendixD),b) Ballute (seeAppendixD);

3) Weight of ACS,fixed

WACSF= 23.5 + 0.00022WE DA/S+ 0.0027WE

117

APPENDIX B

4) Weight of telecommunications cabling

WTE = 0.7 DA/S (DA/S = 15 ft max.)

5) Weight of entry thermal control

WTH = 13 + K (DA/s)2 ; (DA/S = 15 ft max.) K = 0.i

(orbit), 0.02 (direct)

6) Weight of pyrotechnic control

Wpy = 25 + 2 DE/S (DA/S = 15 ft max.)

Figures B155 thru B186 show retro system weights for both

orbit and direct modes. Ballistic coefficients of 0.i, 0.2,

0.3, 0.4, and 0.6 slug/ft 2 are used. The retro initiation alti-

tude for each BE is printed out below the B E . The F/M ° of

the retro is equal to drag acceleration at initiation altitude

and is constant for all diameters at a given B E . The weight

breakdown is the same as outlined above for the aerodecelerators

except for deletion of the aerodeceleration system and substitu-

tion of bipro_ellant motors for monopropellant motors. The ACS

weights are not included in the retro figures, and they should be

subtracted from the landed equipment weight for a correct value.

Equations for ACS weight are shown in the above parachute break-

down. Landed weight ratios for the inflatable afterbody/retro

system are shown in figures B187 thru B194.

The ablator thickness is based on the design condition of

minimum flightpath angle (Appendix C). The weight equations re-

flect an abrubpt change in ablator thickness at B E = 0.3 due to

flow transition. This is illustrated in figure B195. Curves were

arbitrarily faired to create a smooth curve as shown in the figure.

Much of the data presented herein is in terms of WLE, landed

equipment weight. This is defined with the aid of figures B196

thru B200, which illustrate the components subtracted from entry

weight, WE, to arrive at WLE. Weight breakdowns are shown for

aerodecelerators with 8.5-, 15-, and 25-ft diameter aeroshells,

and a retro system with a 8.5-ft aeroshell. It may be observed that

the largest differences between configurations is in aeroshell

weight. For the aerodecelerators, the large increase in aero-

decelerator weight determines the peak landed equipment weight.

118

APPENDIX B

0

r5 _ "-_ 0

Z _ 0 _ I

_ .... II

>-CO

t I

Ix.

6

/J

t t I I I + I t

LD U3 _ [13

6 S 8 6oT_e= _4_a_ papue_

t

nJ

0

I

8

c3

P.

_.yI I

68

S8Pd

88

S

0

0

119

APPENDIX B

120

II II 11

I--

t I I I

r_ t.D

6 o

4,--

U-1 '_r m

o 6 6oTa_a _qgT_ papue X

I

_U

6

I

.f e

6

0

G

APPENDIX B

II II

i I I I

r'_ In

6 o

I t t I t

o_:le= _q$_.e_ pepuex

I

CU

0

i

6

6

c3

CD

0

o

i

g

121

APPENDIXB

122

APPENDIXB0

o

}--

0

t I t t

b. kD

6 6

I

k_

c3

t I I t I I

_ m Iu

6 6 o

I

6

S

_S0

r_

u

.&

t

0

0

0

o

i

&

123

APPENDIX B

.,.i

t t I I

r,,,

c5 6

I

I I t I I t I I I

I,,I'1 _ m I'Ll

6 6 6

I

c_

c_

0

0

124

APPENDIX B

o

_o •

c3 _ L, , ,,Z

Z

3"--

.JO_W

6

I

t.O

c5

I I t t I I t' t I

IJ'] _" m FLI

o c_ 6 c_ 6o_ea _q_$a_ papue_

c_

b,,

tO.=

9

6CD

C)FtJ

x-.q

cb

CD

c5

r_

c_

i°

oo

125

eAPPENDIX B

_11 II l

I I t "t

8 8

I '4 I I ,I

8 8

I I t t

flJ ,,-I

o 8

,r.t

8

0

0

126

APPENDIXB

=> g

.1_ i:1. o _ iu 0

I---

C3

I t I

c5

t { I - t I t I I I

kD t.n '_r crl t"U

o 6 6 6 6OT_= _q_T_ papuex

I

6

c3

_q

O

6

6 o

am

O_ ,

C_

68CU

600

6

0

6

127

APPENDIX B

I.--

6in u_ ,t o_

o _ 6 60T_ _H_T_ papu_q

L

c_ c_

0

0

6

o

e_

i

i

128

APPENDIXB

¢

_ qJo -- o

o oC.J m _ It n ItZ _

I I ! t

r_ U3

6 o

).(,

t

t_

6

I I t I I t I t

6 6 & 6

O

S<D

8

CD

8

O

O

O

2

t.

o_

g

129

APPENDIXB

N II n

I--

g

t I t I

r,. CI_

c5 c5

x

\

I t I t I I _ t I

o;_I _qST_a p_pu_q

I

..-4

6C9

F'..

ca

i

6

130

APPENDIXB

.,-i m

o

H H O

Z

0.._1rl

N

I I I I

6 o

/

<

I t ] t I t I

I._ _ Crl n.I

o 6 6 6o_e= _q_Ta_ papuex

.rq

6

6

-H

r-_

T

i.

C)

$

&

C)

&

131

APPENDIX B

1.1

m_47

! !

t_

6

I

LD

0

I I 'I + - I t I I I I

LFI "_ m ['U .,-4

6 6 6 o c3oT_ _q_Ta_ p_pueq

.r4

6

0

6

u_

c_

E_

i

132

APPENDIX B

m _o.;Vm,-_ | | |

t t I I I I I I I I I I t

6 6 6 ooTIS1 ]q_T_a p_pueq

_4

o

i

0_

0

6

133

APPENDIX B

e

_0 -,T on 11 0

t I

c5

/

L.D t.n _ 0"1 ['U

6 _ 6 6 6oI_e_ _q_I_ p_pue_

I

6

u

0

0

c_

134

APPENDIXB

_o.., "_,

. _ g II N

Z

N

P_

6

I

kD

6

I'---l'-------f_-------'l---- i I J -_-------'I------+----+ "----_

LI_ _ C'_ _ ._I 0

6 6 6 o 6 6oT_ _q_Ta_ papu_

0c:30

I°00O'3

CD0

IB

s

ca

r_

I°

8

BJ

,58

6

135

APPENDIXB

,_ H I H

K-

>-

r_ tD

6 8

t I I I I I I I I

I/1 _r m l'lJ _-I

8 6 _ o G

0

0

6

c

o

c_

s0J

136

APPENDIX B

> o o

,_ I M 1

F-ZIM

2,.-[B.-J['1_.WD

+-_--I---- t

t_

_5

I

O

I t I I I I I I I

_ N" frl t-U ._q

c5 c5 d_ c5 c5o!_e= _q$Iem pDpue%

6000

60003

t .

t°008[}

is

o

8

I

000DJ

t,:5_g..,-I

0

6

137

APPENDIX B

0

_> oo

I I I I

b. in

6 o

I

t.n

_5

+

I I I I

'_r f.n

6 6oT_= _q$$_ p_pu_x

u_

I I i I I

rIJ

6 6

,-q

6

0

6

138

APPENDIX B

e

o

_o _,

2_ _

.I

! I I I

c5 o

I

t.n

c5

<<4 I I I

c; c5oT_el _q_T_ p_pueq

I I I I

FU _-4

c_

[3J

-rq

++CD

0

c_

o

B_

I

139

APPENDIXB

W II !

I t

b.

6

I I _ t I t I I I I

o 6 6 6 6

I i

6

s

i°

6

O

6

_40

APPENDIXB

I I I t

r-.. LD

8 o

t

kn

8

t!4-

'\

\

I I I I

8 8OT:_e:t :_qST_ p_pue'i

i.

88

I I I

8 8 8

o

rn

F4

i.

141

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+- I I i

6 o

I

6

w

o.,4

>

m

oz

nlZl

I .... -+------+----+-----+- I

•_" O] OJ

6 6 oOT_te.z _q_Ta_ p;_puw"I

!

,r-t

cS

6OlDCDOlD

C)CI]

i

c_CD

6

6OlDC;D

6Ji

c3

215

APPENDIXB

6

o_o

6_

I

ZH

6 6

o

_ 6 6

1___.I e

.,-g

0

! ! I

_I 0

216

APPENDIXB

6

_q

LOoAJ

_O°o°

F--31

2!H<n_

o

o

I

g

80

Q4-- t

p,,,.

6

I

t._

(Z)

I I t t I

c5 _5 _5oI_I _q_I_ p_pu_x

t

FU

0

-_ CD

6 6

217

APPENDIX B

66 _

2§.

,_ R m r-

6

!

z

-o

_q

[]

_o

l

g

_-. d3 UD _r

6 & & 6 6o_e= _qSTo_ p_pueq

218

(D

FU _-4 0

6 6 6

L_ -

6_qLEfL

APPENDIX B

6_

o4a

_o

p-.ZI2

Z

<i

l_ L_ _-_k-- EF. _T_

Q-F - _-k + H------+------F +-----+-_-+---

S o d S 6oT_I lq_T_ pepue_I

o

4a

,.-.i

.,.i

.,-iul

Qz

6 6

, 0+0

+©

&, 0

+0

_ a0+o _

_ g+ _ _

+

1-

-.q

+

+

O

219

APPENDIX B

.4

.3

O

_z_o

"_ .2

_D

_D

.i

i •

Note:I. MD=3.

2. 7 E -26 °

3. V E = 18 000 fps.

4. h T = 0.

5. DA/S 12 ft.

/ _.)

1

0 i000 2000 3000 4000

Entry weight, Ib

Figure B195.- Ablator Weight Discontinuity Example

220

,--40

..C

0

<

",,, \\

\

4J

_-_ 000

• 00

E II II

o0

!

II

APPENDIX B

E

>_ cc O0t_ o9

CD _-4

0 .__-- C ----_0---- C-

O 0 :_

0 c_ o 0

_J _ 0

°_.x _ _---_-_--'_,_

000_o

U

E-_

L_

1 II I I

0 00 00 ,..o

\

, I

qI

00o,I,--4

'aqg!a M

\\

i

0000

\

00

000c,I

00,.0,-_ .c

J_a0

g_

1-1

M

00_0

00<1"

0

_J

08.;.q

C

E

o"

"0

n_C

!

o_

_C

221

APPENDIX B

i

_0

6 _ i_03 I I I I I

Not_____e:I. Diam = 25 ft. /

2. M D = 2. /

3. VE = 18 000 fps. /

4. 7E = -16°./

/

5. hT = 6000 ft./

Aeroshell4 \

al aerodecelerator

3 " _" Total (loaded) vernier _

f I system (monopropellant) _

.,/" I 1 1 \ 1\/"/ Thermal control and ACS \

I/ I \ \./_1.-2 / Landed structure \: 2_,_"vl

/ and legs \// |

/to subsystem, diameter- _ ____1

I sensitive cabling, ,_//_

aeroshell mounted science _//_

I l I I_.W : Landed equipment-----._

0 -- [ weight_ l i

1 2 3 4 5 6 x 103

Entry weight, Ib

Figure B198.- Landed Equipment Weight

222

APPENDIX B

E

O

4J

,-4

O

O,--4

,-4

,-4O

O

i

NIX

. Y',,\

,\ "x,

I \iI1i

ii

ii

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• • ,--4 0

II _'_ _ I

II II 11 II

Zl

I IJ I

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o)

or)

_

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-- 4.1 --

0

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i

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1,, \\ -.

\\

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4.1

0

e_°_

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l

00_._

O00 .,..40 •

o

4J

0

E

0 _

0 ,-4

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00

-.1"

0

223

OO

APPENDIX B

> cJ

m U "_

(:K)

I eJ_bo _, _ _ _ _

QJ _

o _ •0

,-C. m (_

_ " _ - _ ._ -

/ °i °

0 0 00 0 0

co _-

O

b0

.=O,0

!

O¢q

bO

224

APPENDIX B

Data Analysis

The data shown in figure B201 is a reproduction of figure

B99. These data are converted to WLE versus W E in figure

B202. Three specific contours are identified on the figures.

The first is maximum WLE for each diameter. This contour is

the maximum system performance for any given diameter. The

second contour is really the envelope of the curves. The contour

describes the maximum WLE for any given W E . The price paid

for using this contour is increased aeroshell diameter (i.e.,

minimum ballistic coefficient). The third contour lies between

the first two and is a locus of maximum WLE/WE ratio for any

given diameter. This contour is one that defines the most effi-

cient system in terms of maximum pounds on the ground per pound

entry for any given diameter. These three contours are referred

to as maximum WLE contour, maximum WLE envelope, and maximum

WLE/WEI ratio contour, respectively.

Plots of maximum WLE and W E versus aeroshell diameter for

the orbit mode aerodecelerators are shown in figures B203 to B210.

These figures are crossplots of the peaks of curves similar to

figure B202. They show the maximum landed equipment weight and

the required entry weight for a particular diameter. For instance,

from figures B205 and B206 for a diameter of i0 ft and 7E = -18 ° ,

the maximum landed equipment weight is 740 lb. The entry weight

for this condition is 1650 lb.

Plots of landed equipment weight and entry weight versus aero-

shell diameter at (WLE/WE)max.__ for orbit mode aerodecelerators

are shown in figures B211 thru B216. These figures are cKossplots

of the peaks of curves similar to figure B201. Similar data for

the direct mode are shown in figures B217 thru B232.

225

APPENDIX B

• >u

• ,-_ 0 • 0

II II II II

,°

Zl

II

0

J.J

0

0

p_

J_J

r_

tJr!

0

Q;

0_.J

(J

0

!

0c-J

226

APPENDIX B

=

• >o

co _ • =

• ,-.4 0 " 0

I ',D 0,1 _

II II 11 II 0

\-\

I

0

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01)-r4

I00

I I I J I I J i J000¢q

00

000

00

I0

227

_0

,-a

_0

.4

4J

E

_D

_q

2800

2400

2000

1600

1200

800

400

0 4

Aeroshell diameter, ft.

Figure B203.- WLE Versus Aeroshell Diameter for Orbit

Mode Aerodecelerators

228

APPENDIX B

I0 x 103

r-q

O0

--Note:i. Maximum WLE.

2. MD = 2.

3. V E = 4.5 km/sec.

4. h T = O.

I

I _

iI

7 E = -16 °

7 E = -20°

4 8 12 16

Aeroshell diameter, ft

Figure B204.- WE versus Aeroshell Diameter for Orbit


20

229

APPENDIXB

2400

2000

1600

• 1200E

800

400

I I I I I

I. Maximum WLE. [

2. M D - 2.

3. V E - 4.5 km/sec. [

4. hT = 6000 ft. [

4

//

/

8 12 16


20

Figure B205.- WLE versus Aeroshell Diameter for Orbit


230

APPENDIXB

lOxlO 3

4J

6

4

Note: I. Maxium WLE.

2. M D = 2.

3. V E = 4.5 km/sec.

4. h T = 6000 ft.

__ _7E = -16 °_

7E = -18 °

_ 7E = -20 o

4 8 12 16


20



231

.=O_

4.J

E

.rq

2800 --

2400

2000

1600

1200

80O

400

i

Note:

APPENDIX B

1i. Maximum WLE.

2. MD=3.

3. V E = 4.5 km/sec.-

4. h T = 6000 ft.

// i

/

/

I

0 4 8 12 16




20

232

APPENDIXB

1. MaximumWLE.2. MD= 3.3. VE = 4.5 km/sec4. hT = 6000ft.

4 8 12Aeroshell diameter, ft

16 20

Figure B208.- WE versus Aeroshell Diameter for OrbitModeAerodecelerators

233

28OO

2400

2000

2

°r,,II1)e 1600

4..1

E

"_ 1200

800

400

APPENDIX B

Note:

i

i. Maximum WLE.

2. MD=5.

3. V E - 4.5 km/sec.

4. h T = 6000 ft.

/

Ii V

/ o

4 8 12 16


20



234

APPENDIXB

6

coo_

4

No te : i. Maximum WLE.

2. _ m 5.

3. V E = 4.5 km/sec.

4. h T - 6000 ft.

///

4 8 12 16


Figure B210.- W E versus Aeroshell Diameter for Orbit


20

235

APPENDIXB

2OOO

1600

.Or-q

=_ 12oo

800

C

4OO

] i l [ [

Note : i. Maximum WLE/W E.

2. _ =2.

3. V E = 4.5 km/sec.

4. hT = 6000 ft

8 12 16 20




236

APPENDIXB

i0 x 103

8

i

oo

4 ....

2

0

J I r i Jt

iNot___e: 1. Max imum WLE / W E .

2. _B2.

3. VE = 4.5 km/sec.

4. h T = 6000 ft.

f --+-

i

4 8 12


16

-16 °

_18 ° ---

_20 °

20



237

,-4

O0

4J

E

2800

2400

2000

1600

1200

8OO

400

_JNote :

APPENDIX B

_i___/ I L_

1. Maximum WLE/WE. l

2. MD=3.

3. V E = 4.5 km/sec.

4. h T = 6000 ft.

i

/

i/

0 4 8 12 16




20

238

APPENDIXB

,.-4

.=

OO

i0 x 103i

6

4

1 T--_-- 1 INote : i. Maximum WLE / W E . _........

2. MD=3.

3. VE = 4.5 km/sec.

4. h T = 6000 ft.

!iiiIlli...........i I ....

_ I - 1- t -

I

i-

4 8 12 16 20




239

Z

E_L

_r

28O0

2400

2000

1600

1200

APPENDIX B

800

400

____L__i__J__h_

Note,i _xi_umWLE/WEI2. MD=5.

3. VE = 4.5 km/sec.

4. hT = 6000 ft.

/

// O_

_4

0 4 8 12 16




20

240

APPENDIX B

_e,-4

o0.r4

I0 x I03 1

6

4

I Note: i. Maximum WLE/W E .II 2. MD = 5.

' 3. V E = 4.5 km/sec.

I 4. h T = 6000 ft.

i

--- T ....... '

II

I I.......I0 4 8 12 16

I

Ii

II

i

I

I

20




241

APPENDIXBeq¢,%

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242

APPENDIX B

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243

f00

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244

APPENDIX B

oO

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245

APPENDIX B

4..1

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246

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APPENDIX B

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247

APPENDIX B

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248

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APPENDIX B

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249

APPENDIX B

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APPENDIX B

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APPENDIX B

/

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252

APPENDIXB

d_J

00

I0 i_ 103

i I ! I

! Note: I. M D = 5.

2. VE = 18 000 fps.

8 ----- 3. h T = 6000 ft.

6

2

o

///

/

i

8 12 16 20 24 28 32


Figure B228.- WE versus Aeroshell Diameter for Direct Mode

253

/

¢qI

II

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oO OO

•_ II II U

il

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APPENDIX B

ao

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II

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=_0

0 0 0 0 00 0 0 0 60 _ _ _ ",,1"

ql 'lq$._ _uomd_.nbo popue_l

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254

APPENDIX B

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255

APPENDIXB

! I

r ._./ , ./

:I t /---/I! I

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ql '3q_! _ auamdInba papuex

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256

0

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0

APPENDIX B

....., o_ i....

II

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0 00

I E - 00 oI _ _ _

""4 II II II

["4

Zl

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ql '_q_To_ Ka2u_

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0>

I

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°r.l

257

APPENDIX B

Figures B233 and B234 show the effect of terrain height on

maximum landed equipment weight. For the orbit mode and a 12-ft

diameter aeroshell in figure B233, a zero terrain height yields

about 250 to 400 ib more payload than h T = 6000 ft. The dif-

derence in payload is more pronounced at FE = -16° than

FE = -20°" For the direct mode in figure B234, the difference

in payload at 12-ft diameter is about i00 to 200 lb. Notice the

higher 7E of -38 ° shows the greater difference in payload. The

effect of entry angle on maximum landed equipment weight is shown

by many of the preceding figures. It shows that increasing entry

angle is a powerful factor in decreasing landed equipment. This

conclusion could be considered true of all conditions both orbital

and direct. However, figure B234 shows that increasing terrain

height is also an important factor in reducing landed equipment

weight and must be considered in design studies. Another param-

eter to be considered is entry velocity. For the small range of

entry velocities for the orbit mode (14 000 to 16 000 fps), that

effect is small. The effect of entry velocity for the direct

mode is shown in figure B235. The figure shows that for aeroshell

diameters greater than 12 ft an entry velocity of 18 000 fps yields

landed equipment weights about 200 ib greater than for 24 000 fps.

Plots of landed equipment weight and entry versus aeroshell

diameter for the single-stage retro are shown in figures B236 thru

B241. Curves of constant thrust to initial mass and BE have been

plotted. The landed equipment weights should be considered a

maximum for the particular F/Mo shown. Plots of the F/Mo used

in weight calculations for each aeroshell diameters are shown in

figures B242 thru B245.

Figure B246 compares the retro with flaps and the inflatable

afterbody/retro at VE = 18 O00 fps and "E = -30°" It may be

seen that the flaps have a clear advantage over the inflatable

afterbodies. This is also true at other entry conditions.

258

,-4

.=OO

0J

_Jr__JEr_

O __J

_O_J

r_

.a

2400

2000

1600

1200

800

400

APPENDIX B

-- Note:

Legend.

i. Maximum WLE.

2. V E = 4.5 km/sec.

3. MD = 2.0. /

I I

/h T = 0 ft /

h T = 6000 ft /

//

/ /// /

/ /'/ // /

7E = -16 °

\/_, 7 E = -20°f

/

4 8 12


16 20

Figure B233.- Effect of Terrain Height

259

APPENDIX B

oo0coi

II

\%

II

to0_

00

• oO

E•_ II It

• ° •

000

0 00 0_o c_,_ ,-4

q1 _:ltt_a :lu_mdTnb_ p_pue_I

//

II

,,, \ _,--4

-_ r_ 0 _.o

;. I_ J -.,

° L

0

m0

(D<

=

E_

0

I

¢q

260

APPENDIX B¢q

oo¢q

¢q

000

00

• 0

11 II×

• ° °

I I

00

4-J°r.4

0

0 _

4-1

,,....4 _ _

,--.4 _

0 •

¢q

(D

O0

oo

IO

261

_ZOO

E

_D

_D

2800

2400

2000

1600

1200

800

400

APPENDIX B

Note: I. Maximum WLE.

2. V E = 4.5 km/sec.

3. h T -- 6000 ft.

I I IL_nd:

7 E = -16 °

7E = -20 o

//,s

f -! /

/

t

F/M o, ft/sec 2

I/ 150

! I// Loo

// 1 55/

//

/ I!

// I/ /

i!;'1

4 8 12 16


Figure B236.- Retrodecelerator, Entry from Orbit

20

262

APPENDIXB

i0 x 103

6,=O0.,q

= 4

I

t

lIii

i

iI

Note: I. VE = 4.5 km/sec.

2. h T = 6000 ft.

I I I I

7E = -16 ° ]7E -20 °

Il

F/M o, ft/sec 2

g8 _I

- 150 !

__;X __!

I

I/ ' i

: !'_j/ /fi I"I .7

.4_/4/t',_"- I I_i r

!

4 8 12

I

ill1 i _ooI i i

55li

16 20


Figure B237.- Retrodecelerator, Entry from Orbit

263

o

0r-I

-- u'h--

oO

\\

Ii %

\ %

APPENDIX B _q

oo

\

_ \\ \!

\ \ \

.1_1

00 0

000 0

",.o

II o0 II

i

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b.-qv

APPENDIX B

I

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APPENDIX B

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266

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APPENDIX B

6o,-q

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--- 0 CD

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267

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u_

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APPENDIX B

\

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0 0 0 0',.0 N cO

e_as/_ '°_I/_ 'ssem leI_3uT/_sn_ql

cOO

r_

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,=

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268

APPENDIX B

5OO

4O0O

O

%

300

H

•,_ 200

.r4

i00

i0 x 10 3

Figure B243.- Retrothrust/Initial Mass

269

APPENDIXB

600

lOC

4 6

Entry weight, Ib

I0 x 103


270

o

_o

#

E

.r4

.,4

.,4

4J

Note:

APPENDIX B

I. VE = 18 000 fps.

2. 7E = -40 o.

3. h T = 6000 ft.

4. VM-8.

8oo I i I

/ / /700

6.5 8.5 i

12 15

600

Aeroshell

500 / diametei, ft /

/ ?t300

200 /_

,oo///.

0 2 4 6 8 I0 x I0 a

Entry weight, Ib


271

\\ o

c4

L0

0

0

o9

0

0

\

\

0

0

APPENDIX B

4J

q-4

0 0 o __

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,-_ _0 I

II O0 II II

-

_q

0

I

ql ,aN_ioa au_mdinba p_puex

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00

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r

xO

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:>.,.-I

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o

o

i

¢,i

0o

272

APPENDIX B

A conservative design requires margin to account for weight

uncertainties. To define this effect, the delivery system weight

WE_WL E is increased by 10%. Curves of WLE versus W E with

and without this 10% margin are shown in figures B247 thru B256 for

the orbital parachute system. Consider, as an example, fE = -16°

and hT = 0 in figures B247 and B248. A landed equipment weight

of i000 ib requires at least a 9-ft aeroshell, while a 9.5-ft

aeroshell is required with a 10% margin. Entry weight needed

goes up from 2050 to 2250 lb. Conversely, a 9-ft aeroshell with-

out margin yields I000 ib landed equipment weight, but only 900 Ib

with a 10% margin. Plots of the corrected maximum landed equip-

ment weights versus aeroshell diameter and entry flightpath angle

are shown in figures B257 thru B260. These plots can be used

for design purposes to provide margin. Another means of achieving

a margin is by using the left-hand slope of the diameter curves

discussed earlier in this section. This WLE versus W E slope

is somewhat off the maximum WLE and achieves a similar margin

to that just discussed. Figures B261 thru B265 show these slopes

for orbit and direct modes for aerodecelerators and retros. Figure

B265 shows a decreasing slope and, therefore, decreasing landed

equipment weight with increasing entry angle for the retro. This

trend also holds true for the aerodecelerators. Increasing de-

ployment Mach numbers also have decreased slopes. A comparison

of monopropellant and bipropellant verniers is made in figures

B261 and B262. They show the bipropellant to be superior in

amount of landed equipment weight.

Figure B266 shows the retro/parachute/vernier system perform-

ance for 7E = -20 ° and -30 ° , VE = 24 000 fps, h T = 6000 ft

for an 8.5-ft diameter aeroshell. Figure B267 shows the two-

stage retro for the same conditions less 7- = -20°. The boundary

shown is where the system fails to decelerate to zero velocity at

the surface. Also shown is the single-stage retro and the para-

chute system. For a comparable thrust-to-weight ratio, the single-

stage retro yields the highest landed equipment weight. However,

the two-stage retro has a larger growth potential.

273

APPENDIXB

2000

1600

OD

"_ 1200

800_0

40O

0 2 3

Entry weight, ib

5 x 10 3

Figure B247.- WLE versus W E without Margin

274

APPENDIX B

2000

1600

#m

g

•_ 1200

r_

800

400

Note :i. MD.= 2.

2. 7E = -16 o.

3. V E = 4.5 km/sec.

4. hT= 0.

/20.5

Aeroshell dzameter,

8.5

ft

2 3

Entry weight, Ib

5 x 10 3

Figure B248.- WLE versus W E with Margin

275

APPENDIXB

oO

_r

2000

1600

1200

8OO

400

f J__

Note: I.

- 2.

3.

4.

/

/

il9.11o9

8.5

___I E i__

M D = 2.

VE = 4.5 km/sec.

7E = -16 °.

h T = 6000 ft.

112

10.5


I 2 3 4

Entry weight, Ib

5 x I0 e


276

APPENDIXB

1600

INote"

I I I rI. MD= 2.

2. 7E = -16 o.

3. VE = 4.5 km/sec.

4. hT = 6000 ft.

¢Jr_

°_

r_

1200

8OO

400

/

/"_eros diiameter,

8.5

ft

2 3

Entry weight , Ib

4 5 × 103


277

APPENDIXB

2000

1600

.c

.-4

g

_0•_ 1200

C

E

cr800

400

Not____e:

1 i L__

i. MD= 2.

2. 7E = -18 °

3. hT=0.

/

L9.5

8.5

/\12

10.5


i 2 3 4

Entry weight, lb

5 x 10 3


278

APPENDIXB

1600

1200

N?-,-4

4J

¢_ 800

-,-I

_J

OJ

400

/

I

Note:

r J i i

i. MD= 2.

2. 7E = -18 °.

3. VE = 4.5 km/sec.

4. h T = 0.

_ 10.5

@ 9.5

8°5


1 2 3 4

Entry weight, Ib

5 x I0 a


279

APPENDIXB

20O0

1600

"_ 1200

E-_

Er

800_D

_a

4OO

/

INote:

i0

12


I

2 3

Entry weight, Ib

5 x 103


280

APPENDIXB

200(

160_

•_ 1200

4-J

.,4

/X

8.5

9.5 i

9

[Note:

I I II. MD=2.

2. 7E = -20 o.

3. VE -- 4.5 km/sec.

4. h T = 0.

/\12

10.5 Aeroshell diameter, ft

2 3

Entry weight, Ib

5 x 103

Figure B254.- WLE versus WE with Margin

281

APPENDIXB

.Q

J.=

4J=

o"

1600

1200

800

400

/I

Not__.__e:

iI" MD=2' ]

2. V E 4.5 km/sec.i--

3. 7 E = -20 ° .

4. h T = 6000 ft.

12

_ I0,5

10

9.5

8.5


2 3 4 5 x I0 _

Entry weight, Ib


282

APPENDIXB

,-4

_ZOO.r4

4J

-r4

1200

800

400

/Aeroshel

% 9

8.5

12

[ i I 1

Note:

I diameter, ft

I. M D = 2.

2. 7E = -20 °

3. VE = 4.5 km/sec.

4. hT = 6000 ft.

0 i 2 3 4

Entry weight, Ib

5 x i0 _


283

APPENDIXB

2000

1600

E

1200

800

0

Note: I. MD = 2.0.

2. V E = 4.5 k: /sec.

3. h T = 0.

/400

0 4 8


0 ---7

I

i

1 @/8

// ---

I12 1_

Figure B257.- Maximum WLE , 10% Margin

284

AIq_I_NI) IX F,

,.Q

_a

E

QJ

"0

0_,IJ

l,a

0

2000

1600

1200

800

400

I I

, Note: I °

2.

3.

I I I

MD= 2.0.

V E = 4.5 km/sec.

h T = 6000 ft.

f

"/_/

i ! J

///// ............



285

8OO

4OO

,.Q

QI

e"

E

.,.4

Cr

e_

800

0r_.)

400

10

APPENDIX BI 1 r I I

Note: i. MD = 2.0.

2. VE = 18 000 fps.

3. hT = O.

\

20 30

Entry flightpath angle, 7E, deg

40

'8 12 16



286

APPENDIX B

8OO

..c

r_

400

O0

E_ 0°r4

"t3

800

0

t._)

400

Note: I. M D = 2.0.

2. VE = 18 000 fps.

3. h T = 6000 ft.

\

i0 20 30 40

Entry flightpath angle, _E' deg

o/ss /0 J

g 12 16


Figure B260,- M_im, m WLE , 10% Margin

287

APPENDIX B

ooco

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rj

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pcl

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q1 ':lq_taa :luamd_nba papuerI

o

288

APPENDIX B

• • °

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0

0'43

I I II0 0 0

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ql C]q_!ea _uamdynba papue X

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289

APPENDIX B OO

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290

APPENDIXB

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291

APPENDIXB

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u_,,0¢,,I

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292

APPENDIXB

600

J.J

E

.,4

¢J

5OO

400

300

2OO

: I. VE = 24 000 fps.l__-

-- 2. h T = 6000 ft.

3. DA/S = 8.5 ft.

--F/M , Earth g-o

i00

400 800 1200 1600

Entry weight, Ib

Figure B266.- Retro/Parachute/Vernier System

2000

293

APPENDIXB

.=

E

.r4

600 --

5OO

400

300

200

i00

I I I I I

Note:l

I. VE = 24 000 fps_I

2. 7E = -30 °.

3. hT = 6000 ft.

4. DA/S = 8.5 ft.

8 g-----,

3.3 g_.

Single-stageRet ro j"

v¢I

ParachUteM= 2 /

)/

'///_', /2o

400 800 1200

Entry weight, Ib

1600

Figure B267.- Two-Stage Retro Performance

2_4

APPENDIXB

A comparisonof the terminal phasesystems(subsonic parachute/vernier, ballute/vernier, all-retro) maybe madeon the basis ofseveral parametersfor both orbit and direct modes. In general,the parametersof interest are landed equipmentweight, entryweight, aeroshell diameter, and entry flightpath angle. Thecom-parison is madeat a constant terrain height (6000ft); the ef-fect of varying terrain height was shownearlier. Figures B268and B269comparemaximumlanded equipmentweight versus aeroshelldiameter for orbit (VE = 4.5 km/sec) anddirect (VE = 18 000 fps)modes,respectively. Representativemaximumandminimumentryflightpath angles are included: 7E = -20° (orbit) and -38°(direct), YE= -16° (orbit) and -26° (direct). Comparisonismadebetweenparachute/vernier (MD = 2), ballute/vernier(MD = 3,5), and the all-retro decelerator systems. Theretrosystemthrust to initial massis approximately i00 ft/sec e. Thesefigures are compositesof earlier data and are used an an exampleof the types of comparisonleading to the summarycharts.

The terminal phasesystemscapabilities are summarizedas afunction of assumptions. Thebasic WLE capability is shownin figures B270and B271for fixed diameters of 8.5 and 15.0 ft.Thedata are basedon the maximumWLE contour and, therefore,represent maximumWLE for the given diameters. The data shownin figure B270for the 8.5-ft diameter comparethe orbit anddirect modesfor the various systems. Both entry weight and WLEare shownas well as the sensitivity of both of these parametersto entry flightpath angle. For the orbit mode,the Mach5 balluteprovides the highest performanceper diemeter, but also has thegreatest entry weight. Theretro systemis slightly better thanthe MD = 2.0 parachutesystem. Thedirect modedata showthesametrends, except the retro systemis relatively better. Noneof the direct modeperformanceis acceptable at a diameter of 8.5ft. Similar data are shownin figure B271for a 15-ft diameter.Theorbit modedata showhigh performancecapability with thesamecharacteristics across systemsas wasexhibited for the8.5-ft diameter data. Thedirect modedata, for a ?E maximumof -24 to -26°, showacceptable performance,with the retro sys-tem showingthe best performancecapability.

295

2400

2000

f

16004.J

OJ

.v,,l

I1)

L200

8OO

400

-! MD2

Im, 3

I

-- . Retro

!

¢

I

4 8 12 16


20

Figure B268.- Effect of Deployment Mach Number

296

.

Ii

l

I

i

I0

00

¢xl

\\\

APPENDIX B

//

//I/I!

ql _q_ _uamdInb_ p_puex

cq

//

7////

o

I

U 0

,---I

m0

\ °

",.1"

0

4..1

Z

0

e_

0

I

297

2000

2400

2000

I000

.C:_C

1200

80O

400

APPENDIX B

I. Maximum WLE.

2. h T = 6000 ft.

Orbit

V E = 4.5 km/sec

W E at

7E = -16 o

= -20 °

WLEat

Direct

V E = 18 000 fps

= -26 °

= _38 °

Figure B270.- Terminal Phase Summary, Aeroshell Diameter = 8.5 ft

298

7000

6000

5000

4000

t_

3000

20OO

i000

APPENDIX B----[ .... Z ....._------T- -_ .......

Note: i. Maximum W ]LE" i

2. h T = 6000 ft.l .......I

V E = 4.5 km/sec

--W E at--- ]

7E = -16 °

i

_- 7E = -20 °

WLE at

= -16 °

7E = -20 °

]

!

Direct

V E = 18 00_ fps

i i

O

W E at4J

O

--------+_ _ o- 7E = -26°

IIII

. -, _rrlll I" _1 IIIlile_ Illlll

__ 7E = .38 o

WLE at

_-/_-7 E = -26 =

\y/ : -3_o

| | , ,

Figure B271.- Terminal Phase Summary, Aeroshell Diameter = 15 ft

299

APPENDIXB

It is clear from these data that any of the systemscan deliveruseful payloads (WLE) in the 600-1bclass with diameters in the8.5 to 15.0-ft range. Generally, the ballute performanceissuperior for the orbit mode,while all-retro systemsare mostfavorable for the direct mode. At fixed diameter, the orbit modeprovides the greatest WLE and continues to provide maximumpoundson the groundper poundentry weight.

Thesecondcomparisonis madeon the basis of fixed landedequipmentweight of WLE= 600 lb. This is summarizedin figureB272in terms of minimumrequired entry weight. Onceagain, thesensitivity to entry flightpath angle is indicated by the shadedportion of the bars. For the orbit mode,the MD = 2.0 parachutecase requires the lighest entry weight with the all-retro systema close second. For the direct mode,all of the required entryweights are greater than for the orbit mode,with the retro sys-tem requiring the smallest entry weight. Theretro systemdataas presented, are somewhatdeceiving relative to its apparentinsensitivity to 7E. In reality, the data presentedearliershowthat the variation in 7E require a wide variation in re-quired thrust-to-weight ratio.

Therequired aeroshell diameters for the abovecomparisonareshownin figure B273. For the orbit mode,the MD = 2.0 parachutesystemrequires the greatest diameter of the aerodecelerator sys-tems, with the all-retro systemrequiring the greatest diameter.Onceagain, the retro data canbe somewhatmisleading becausethe data are basedon using the minimumuseful thrust-to-weightratio. There is a tradeoff betweenthrust level and diameter.For the direct mode,the required diameters are all greater thanthose for the orbit modewith the MD = 2.0 parachute systemrequiring the largest diameter. Again, assuming -?E maximumof 18° for the orbit modeand 26° for the direct mode,the cor-respondingdiameter ranges are 8.1 to 9.5 ft for the orbit modeaerodecelerators and 12.3 to 13.6 ft for the direct mode. Thesedata correspondto the caseof landing 600 Ib of WLE at thelowest entry weight, WLE, without regard to aeroshell diameter.

300

APPENDIX B

4 x I0z

3

J.J

oo

0

Orbit

V E = 4.5 km/sec

i

eq ¢_ zrl J O

II I| II I b_

I

I I 7 I I I

[Note: i. Maximum WLE/W E envelope.

2. h T = 6000 ft.

ID ire c t

tV E = 18 000 fps

I -_E -20° \\ _'_

-- _-- - t .......

)E = -16°

t_

c-j

II

-- ---n r_-

__ I i

kFi0

II - II bl.IJ

L

° __

FE = -38°

i

FE = -26°

Figure B272.- Terminal Phase Summary, Landed Equipment

Weight = 600 Ib

301

APPENDIX B

24

20

16

q4

E

"_ 12

,-4

.C

O

<

8

0

Note: i. Maximum WLE/W E envelope.

2. h T = 6000 ft.

Orbit

VE = 4.5 km/sec

I! X 41 O

_20 °

.16 °

.... ...........

J

i l

Direct

V E = 18 000 fps

O

II

II II II

0

Figure B273.- Terminal Phase Summary, Landed Equipment Weight = 600 Ib

:]02

APPENDIX B

The final comparison made here is on the basis of a fixed

entry system weight of 1500 lb. The maximum allowable landed

weight for the systems and mission modes is shown in figure B274.

The efficiency of the M D = 2.0 parachute is once again apparent

in providing the greatest pound on the ground per pound entry for

the orbit mode, while the all-retro system is best for the direct

mode. The corresponding required diameters are shown in figure

B275. Once again, the orbit mode diameters are smaller than the

direct mode cases with the M D = 2.0 requiring the largestdiameters.

The variations in these results are only indicative of the

kinds of tradeoffs that can be made. A generalization that can

be drawn from the analysis is that the M D = 2.0 parachute is

best when compared on the basis of pounds on the ground per pound

entry. The supersonic ballutes provide the maximum pounds on the

ground per foot of aeroshell diameter. The all-retro systems are

always competitive, but care must be taken in their evaluation

to consider their high sensitivity to required thrust-to-weight

ratios.

Before concluding this summary discussion of the data, some

mention is made on the degree of conservatism implied in the

analysis from a mission profile viewpoint. There are two as-

pects to this question discussed here. The first, that all of

the results presented above are based on the following:

i) The VM-8 atmosphere plus design terrain height define

the altitude mark to trigger aerodecelerator deploy-

ment;

2) The VM-7 atmosphere in combination with 220-fps hori-

zontal wind design the aerodecelerator size and

vernier propellant requirements. The VM-8 atmosphere

defines the deployment, design dynamic pressure;

3) The 3o steepest entry flightpath angle is used in

conjunction with 30 orbit or approach trajectory

emphemris uncertainty.

303

APPENDIXB

i0

80(

I I i l [ L_I

Note: I. Maximum WLE/WE envelope.

2. h T = 6000 ft. I

Direct

V E = 18 000 fps

_0

•,_ 60(

E

r_

20(

/_i 7 E "20° _-TE = -26°

ff "-TE -38 °

I

. , , o i " " "J

Figure B274.- Terminal Phase Summary, Entry Weight = 1500 ib

304

_.J

C

E

<

2O

16

12

4

APPENDIX B

Note:

I__ ,. Orbit

f J

fJ

I

t i

II

I. Maximum WLE/W E envelope.

2. h T = 6000 ft.

Direct

.......... V E = 18 000 f_s

V E = 4.5 km/secii ................. \\

fJ

7 E = -20 o

r I

...... ---- ----4-- --?_ -16 °

%%/

t._ 0 - -

I

II

i

II II

m

II _-I

Z=_®l

!

7 E = -38 °I1

I

/ _7E = -26 o

II

Figure B275.- Terminal Phase Summary, Entry Weight = 1500 Ib

305

APPENDIX B

All of these things must occur simultaneously. These constraints

really fall into two categories from a design viewpoint. The

first is that the worst entry flightpath angle and atmosphere are

assumed in direct combination. The system might still work if

either one or the other is worse than expected. The second

category is the definition of altitude mark for aerodecelerator

deployment. This must be based on an assumed design terrain height

and worst atmosphere. If the terrain height is higher than assumed,

the probability of successful landing is low. An altitude mark

must be set before flight, and there is less inherent adaptability

in the system to cope with surprises. Even so, the combination of

events that must occur simultaneously that have been assumed in

this analysis should reflect as some degree of conservation in

the results.

The second aspect of conservatism that has been partially in-

vestigated here is the uncertainty in the weight equations dis-

cussed earlier (10% weight margins).

The maximum performance characteristics of the various terminal

phase systems and launch vehicles (in terms of capsule system

weight) are compared for fixed aeroshell diameters of 8.5 and 15.0

ft, fixed landed equipment weight of WLE = 600 ib and fixed entry

weight of W E = 1500 lb. As in the previous summary type (bar)

charts, the constant diameter data are based on maximum WLE for

a given diameter. They, therefore, result in the best diameter

at a cost of increased entry weight. The fixed WLE comparisons

are based on the maximum WLE per pound entry weight. They re-

suit in the most efficient system from an entry weight veiwpoint

at a cost of increased (nonoptimum) aeroshell diameter.

The 8.5-ft diameter condition is summarized in figure B276.

These data show the range of total capsule system weight as a

function of ?'E (highest values on bars correspond to shallow-

est )E)" The corresponding WLE are also shown on the bars

for reference. The scale up the middle of the chart shows the

performance capability of the various launch vehicles. It is

clear from figure B276 that all of the orbit mode cases require

tile Titan lllC/Centaur launch vehicle but with considerable launch

vehicle margin. The high deployment Mach number aerodecelerator

has the most WLE performance capability. The direct mode perform-

ance requirements fall into the Titan IIIF/Stretched Transtage

capability, but have negligible landed weight capability in all

cases.

306

,-4

O0°_

E

_0

7OOO

6000

5000

4OOO

3000

APPENDIX B

TlllF/Centaur/

I

IIIII

I

TlllC/Centaur /

///!

Orbit mode

16°< (-7 E) < 20 °

I-

!

I

Ib

Note: h T = 6000 ft. I

2000

i000

,_ TIIIF/ Direct modei

_ _ Stretched oTranstage 260 < (-TE) < 38

C_ ii Z 0 /

_ x-_ "TIIIC _ ,

WL E _ _-_ xx -I--- ,

_ 7_J_X_ WEE

Figure B276.- Performance Summary, Aeroshell Diameter = 8.5 ft

307

APPENDIXB

Similar data for the 15-ft diameter aeroshell are showninfigure B277. Theorbit mode WLE is large, but unfortunatelythe total capsule systemweight exceedseven the Titan IIIF/Centaur capability.

Thedirect modeperformancerequirements clearly fall withinthe Titan lllC/Centaur range, with the corresponding WLE compar-able to that obtained with the 8.5-ft diameter orbit mode. Again,the Titan lllC/Centaur is the launch vehicle on this basis andagain, with large launch vehicle margin. For the direct mode,the all-retro terminal phasesystemis the best performer.

Thenext comparisonshownin figure B278is madeon the basisof fixed landed equipmentweight of 600 lb. In this case, thehighest capsule systemweight correspondsto the steepest entryflightpath angle. Thedata showlowest total systemweight forthe orbit modewith the MD = 2.0 parachute is the best of theterminal phasesystems. Thediameters for these casesare givenOn figure B273. The performance requirements for all of the sys-

tems and both modes fall into the Titan lllC/Centaur capability

with large launch vehicle margins.

The final comparison, shown in figure B279, is on the basis

of fixed entry weight of 1500 lb. On this basis, the total sys-

tem weights are generally comparable with the M D = 2.0 parachute

and landed equipment weight for the orbit mode. The most competi-

tive configuration for the direct mode is the all-retro system.

The diameters are shown in figure B275.

It is clear from these comparisons that the tradeoff between

aeroshell diameter, landed equipment weight, and total capsule

system weight can be made in many ways. However, the following

generalizations can be made:

i) Orbit mode - The MD = 2.0 parachute is most efficient

from a weight viewpoint, while the ballutes are favored

from an aeroshell diameter viewpoint;

2) Direct mode - The same generalization relative to aero-

decelerators is true here, but the all-retro system

is competitive. The all-retro system is sensitive to

thrust-to-weight ratios (throttling requirements go

up). Relative to ballutes versus M D = 2.0 parachutes,

the latter are generally preferred on the basis of more

straightforward packaging and release considerations;

308

r_bO

E

co

16 (

700O

6000

5000

4000

3000

2000

WLE

I000

[OrbiL mode----

i i

'< (-?E) < 20 °

APPENDIX B

8590 lb

TIIIC

__ _ oil il It

ITl_IF/Centaur

/

Note:

26 o

/Centaur

TIIIF/Stretched

Transta_e

//

i,,am. ¢

TIIIC

hT = 6000 ft.'Ii

Direct mode---

< (-?E) < 38 °

CK'T

oco Lr_

II r iiii

o[-- 1 ] _

Figure B277.- Performance Summary, Aeroshell Diameter = 15 ft

309

t_

E

_o

7OOO

6000

5OOO

4000

3000

2000

i000

APPENDIX B

3590 Ib

TIIIF/Centaur/

////

TIIIC _Cent aur/

/

.aP-

/I

Orbit modet

16° < (-_E) < 20 °

/I

TIIIF/

__ StretchedTranstage

/4--/

II II nqJ

WLE l I T I IIC 7//,

Not_____e:h T = 6000 ft]

Direct mode

26 °< (-7 E) < 38 °

cq ¢o u_ O

- II _ II II--

WLE

0

Figure B278.- Performance Summary, WLE = 600 Ib

310

O0

E

>_co

C.O

7OOO

6O00

5000

40OO

3000

Orbit mode

APPENDIX I}

JI_- -I.....8590 ib

//

TIIIF/Cent aur

I/ -/ J/ !

!TIIIC/Centaur

!!l

hT = 6000 ft.[

16°<(-7E ) < 20 °

Direct mode

26 °< (-7 E) < 38 °

20O0

i000

TIIIF/ ,i-'_

=z_ _'_ Stretched _: Transtage

u'_ 0 / II MII II _ II II .i_I

Fa}

WLE _:_ _,.,.,.,TlllC

I I I I l I , , ,

Figure B279.- Performance Summary, W E = 1500 ib

WLE

311

APPENDIXB

3) Direct vs. orbit mode Both modes require the Titan

lllC/Centaur launch vehicle. On this basis, the

greater inherent flexibility and adaptability of the

orbit mode makes it the more desirable.

312

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