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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
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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39
APPENDIXB
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41
APPENDIX B
40
30
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l
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LI
0.6 sl/ft 2 --
Note: Mars entry, VE = 21 000 fps,VM-3 and -7.
L20 30 40
-7E , deg
Figure B40.- Peak Drag Deceleration
45
APPENDIX B
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APPENDIXB
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APPENDIXB
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APPENDIX B
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APPENDIX B
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APPENDIXB
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APPENDIX B
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APPENDIX B
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APPENDIXB
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APPENDIX B
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71
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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FT.) t
I
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t_
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r_
-- °
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u_
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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
30
28EO
4-J
26
r-q
r_
O
(D_q
"_ 24
EI.-4
O
2:
22
20
I I I I I
Note: hI I 4000 ft above ]terrain.
/
//J
j
y/
//
8
5v
.IJ
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
APPENDIX B
.20
.16
.12
b.08
.04
0 .i .2 .3 .4 .5 .6 .7
BE , sl/ft e
Figure B75.- Aerodecelerator Size Limit
86
APPENDIX B
.08
04/t
!
.20 ......
.16
%_..12
i
Note: I. MD _ 2.
2. VE = 18 000 fps.
3. _ = 6000 ft.
0 .i .2 .3 .4 .5 .6
B E , sl/ft 2
.7
Figure B76.- Aerodecelerator Size Limit
87
APPENDIX B
0d
.2O
.16
.12
.08
.04
i
-..j
t
\_o
.I .2
\
.3
I I
-- I N°te: I" MD 2"
--I 2. VE " 24 000 fps.l
---i FhT'0., 3.I
I
°\i
BE , sl/ft 2
Figure B77.- Aerodecelerator Size Limit
88
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
i0.0_
8
7
6
5
1.01098
7
G
q_ 5
4
3
2
.I io9
8
f
G
5
.01 ii
I0
3 4 5 G 7 8 910 2 3 4 5 6 I 8 910 2
102 103
Aeroshell weight, Ib
3 4 5 6 78910
10 4
Figure B79.- Aeroshell Ballistic Coefficient
91
28 x 103
APPE_IX B
ote: _ I 6000 ft. /tmin. - 50 secI
/
/¢;: itmin _ 30 sec20
_Data
Varia io
• 16 d
12 "
tmin. - 16 sec
_..T"-Z rt // • // , / , ,1 ,/
8 -
VE = 4.5 km/sec; _E _ -20*
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
Figure B81.- Aerodecelerator Size Limit
94
APPENDIX B
Cq
r-N
b
.08
Note: i. MD = 3.
2. V E = 4.5 km/sec.
3. _ I 6000 ft.
Figure B82.- Aerodecelerator Size Limit
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
Figure B83.- Aerodecelerator Size Limit
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
Figure B84.- Aerodecelerator Size Limit
97
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
Figure B86.- Aerodecelerator Size Limit
99
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
Figure B88.- Aerodecelerator Size Limit
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
Figure B89.- Aerodecelerator Size Limit
.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
Figure B90.- Aerodecelerator Size Limit
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
Figure B91.- Aerodecelerator Size Limit
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
Figure B92.- Aerodecelerator Size Limit
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
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
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Z _ 0 _ I
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t I
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t t I I I + I t
LD U3 _ [13
6 S 8 6oT_e= _4_a_ papue_
t
nJ
0
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8
c3
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68
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88
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0
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119
APPENDIX B
120
II II 11
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t I I I
r_ t.D
6 o
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APPENDIXB0
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t I t t
b. kD
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.&
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123
APPENDIX B
o
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Z
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6
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I I t t I I t' t I
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c_
b,,
tO.=
9
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x-.q
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125
APPENDIXB
=> g
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I---
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I t I
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t { I - t I t I I I
kD t.n '_r crl t"U
o 6 6 6 6OT_= _q_T_ papuex
I
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600
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127
APPENDIXB
¢
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o oC.J m _ It n ItZ _
I I ! t
r_ U3
6 o
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t
t_
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I I t I I t I t
6 6 & 6
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129
APPENDIXB
N II n
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t I t I
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130
APPENDIXB
.,-i m
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131
APPENDIX B
1.1
m_47
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t_
6
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LFI "_ m ['U .,-4
6 6 6 o c3oT_ _q_Ta_ p_pueq
.r4
6
0
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u_
c_
E_
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132
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
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IB
s
ca
r_
I°
8
BJ
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6
135
APPENDIXB
,_ H I H
K-
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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
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c_
s0J
136
APPENDIX B
> o o
,_ I M 1
F-ZIM
2,.-[B.-J['1_.WD
+-_--I---- t
t_
_5
I
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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
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t,:5_g..,-I
0
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137
APPENDIX B
0
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I I I I
b. in
6 o
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t.n
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+
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
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2_ _
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t.n
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c; c5oT_el _q_T_ p_pueq
I I I I
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c_
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0
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B_
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139
APPENDIXB
I I I t
r-.. LD
8 o
t
kn
8
t!4-
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I I I I
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88
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8 8 8
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rn
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141
APPENDIXB
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I I I I
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I I _ I I
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8 6 6oT_ea ]q_I_a p_pue]
6
600
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6o{23
Lnb
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•_ 0
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143
APPENDIX B
¢
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r-_ LO
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dl _ m
c5 c5 6ot_e_ lq_T_ p_pue_
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t - t 'I
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i'-. a,
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0
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144
APPENDIXB
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b-Z
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6 6
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t I t _ I
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I I I '1 t
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145
APPENDIXB
0
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dWEli
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b_
c5tO
0
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8_ o_bq_ i
c.lM _
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146
APPENDIX B
e
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oo_ Z o%
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b. t.D
6 o
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tN
6_ m
c; c5o!_em _qg!am pepuel
I I I
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4-
1°0
6
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147
APPENDIXB
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6 & & 6 6OT_e= _q_lo_ papuex
t _ i t i
nJ ,-4
5 60
0
t
g
5
148
APPENDIXB
g(11o o o
Z
f2_
dN
+ -F---+ .... q- -+-
6 o
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I i I
6 6
I I t _ t
6 6 6o3:_e.I :_qg3_ pBpue'I
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149
APPENDIXB
oo> O o
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N N N
F---
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t -t I I I
6 6 6o'T_ lq_ p_pu_I
_+_____+____+__r__ F- _
_qj _-_
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cq
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+
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152
APPENDIXB
> o o
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+ ..... q____ -+
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154
APPENDIXB
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155
APPENDIXB
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6 6 6oT]e_ ]451_ p_puel
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APPENDIX B
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6
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Lf] _ m F_
6 6 6 6oI_a _q_I_ p_puPq
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=
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mm
158
APPENDIX B
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159
APPENDIX B
4-1
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160
APPENDIXB
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161
APPENDIX B
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162
APPENDIX B
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163
APPENDIXB
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6 6 6 6oT_ _qgT_ pgpue_
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6 6 6
164
APPENDIXB
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165
APPENDIX B
0
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c5 c5
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+I
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166
APPENDIXB
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167
APPENDIX B
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6 6
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168
APPENDIX B
.,-i k_l
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c5 _ 6 6o_ae_ _4$Te_ pepueq
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==
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t t I I --
D.J ",-4 0
6 d d
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170
e APPENDIX B
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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
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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 \// |
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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
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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
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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
Mode Aerodecelerators
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
Aeroshell diameter, ft
20
Figure B205.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
230
APPENDIXB
lOxlO 3
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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
Aeroshell diameter, ft
20
Figure B206.- WE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
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
Aeroshell diameter, ft
Figure B207.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
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
Aeroshell diameter, ft
20
Figure B209.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
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
Aeroshell diameter, ft
Figure B210.- W E versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
20
235
APPENDIXB
2OOO
1600
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=_ 12oo
800
C
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Note : i. Maximum WLE/W E.
2. _ =2.
3. V E = 4.5 km/sec.
4. hT = 6000 ft
8 12 16 20
Aeroshell diameter, ft
Figure B211.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
236
APPENDIXB
i0 x 103
8
i
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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
Aeroshell diameter, ft
16
-16 °
_18 ° ---
_20 °
20
Figure B212.- W E versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
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
Aeroshell diameter, ft
Figure B213.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
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
Aeroshell diameter, ft
Figure B214.- WE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
239
Z
E_L
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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_
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0 4 8 12 16
Aeroshell diameter, ft
Figure B215.- WLE versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
20
240
APPENDIX B
_e,-4
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6
4
I Note: i. Maximum WLE/W E .II 2. MD = 5.
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I 4. h T = 6000 ft.
i
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II
I I.......I0 4 8 12 16
I
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i
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20
Aeroshell diameter, ft
Figure B216.- W E versus Aeroshell Diameter for Orbit
Mode Aerodecelerators
241
APPENDIXBeq¢,%
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8 ----- 3. h T = 6000 ft.
6
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Aeroshell diameter, ft
Figure B228.- WE versus Aeroshell Diameter for Direct Mode
253
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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
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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
Aeroshell diameter, ft
16 20
Figure B233.- Effect of Terrain Height
259
APPENDIX B
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APPENDIX B¢q
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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 -! /
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I/ 150
! I// Loo
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4 8 12 16
Aeroshell diameter, ft
Figure B236.- Retrodecelerator, Entry from Orbit
20
262
APPENDIXB
i0 x 103
6,=O0.,q
= 4
I
t
lIii
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Note: I. VE = 4.5 km/sec.
2. h T = 6000 ft.
I I I I
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Il
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g8 _I
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Aeroshell diameter, ft
Figure B237.- Retrodecelerator, Entry from Orbit
263
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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
Figure B245.- Retrothrust/Initial Mass
271
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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
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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
Aeroshell diameter, ft
I 2 3 4
Entry weight, Ib
5 x I0 e
Figure B249.- WLE versus W E without Margin
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
Figure B250.- WLE versus W E with Margin
277
APPENDIXB
2000
1600
.c
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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
Aeroshell diameter, ft
i 2 3 4
Entry weight, lb
5 x 10 3
Figure B251.- WLE versus W E without Margin
278
APPENDIXB
1600
1200
N?-,-4
4J
¢_ 800
-,-I
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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
Aeroshell diameter, ft
1 2 3 4
Entry weight, Ib
5 x I0 a
Figure B252.- WLE versus W E with Margin
279
APPENDIXB
20O0
1600
"_ 1200
E-_
Er
800_D
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4OO
/
INote:
i0
12
Aeroshell diameter, ft
I
2 3
Entry weight, Ib
5 x 103
Figure B253.- WLE versus W E without Margin
280
APPENDIXB
200(
160_
•_ 1200
4-J
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/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
Aeroshell diameter, ft
2 3 4 5 x I0 _
Entry weight, Ib
Figure B255.- WLE versus W E without Margin
282
APPENDIXB
,-4
_ZOO.r4
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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 _
Figure B256.- WLE versus W E with Margin
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
Aeroshell diameter, ft
0 ---7
I
i
1 @/8
// ---
I12 1_
Figure B257.- Maximum WLE , 10% Margin
284
AIq_I_NI) IX F,
,.Q
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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
"/_/
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///// ............
Aeroshell diameter, ft
Figure B258.- Maximum WLE , 10% Margin
285
8OO
4OO
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Cr
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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
Aeroshell diameter, ft
Figure B259.- Maximum WLE , 10% Margin
286
APPENDIX B
8OO
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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
Aeroshell diameter, ft
Figure B260,- M_im, m WLE , 10% Margin
287
APPENDIX B
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0 0 00 0 004 00
q1 ':lq_taa :luamd_nba papuerI
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288
APPENDIX B
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APPENDIX B OO
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290
APPENDIXB
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291
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
Aeroshell diameter, ft
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
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7////
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I
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m0
\ °
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0
4..1
Z
0
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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