P L A N E T A R Y M I S S I O N E N T R Y V E H I C L E S QUICK REFERENCE GUIDE • VERSION 4.1

Kayla Parcero, AMA Inc.

Gary Allen, AMA Inc.

Al Witkowski, Katabasis Aerospace, LLC

Sharon McKee, NASA Ames

Leah Torres, AMA Inc.

Carol Davies, ELORET Corporation

Compliments of Advanced Concepts Lab (ACL), NASA Langley

This version adds the Phoenix, InSIGHT, LOFTID and Artemis I missions and 

updates data entries to the missions from previous versions with corrected and/or 

added data.

Content Editors: 

Parachute Editor:

Production Editors:

Original Content Editor 

2002:

Cover Photograph:

Version 4.1:

To all those who have contributed to either previous versions of the “Blue Book” or have volunteered to send in information 

to our current release, thank you. Without the support of our community, this book would not be possible.

NASA/SP-20230010341

PLANETARY MISSION ENTRY VEHICLES

Quick Reference Guide Version 4.1
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NASA STI Program 
Planetary Mission Entry Vehicles

N A S A A M E S R E S E A R C H C E N T E R

E N T R Y S Y S T E M S A N D T E C H N O L O G Y D I V I S I O N

VERSION 4.1

QUICK REFERENCE GUIDE

— i —
Back in the era before ubiquitous portable electronics with instantaneous access to all the world’s information at a fingertip, 

one had to rely on dead-reckoning to compare new concepts with historical analogs. This effect permeated all our endeavors 

and certainly so in the niche world of planetary entry vehicles. So, it was in the days of designing the Stardust return capsule

that in meetings, more often than not, someone would ask, “how does this <heat rate, cone angle, g-load...> compare to 

<Apollo, Viking, Galileo...>?” Thus was born V1 of the Planetary Mission Entry Vehicles Quick Reference Guide, thereafter, 

dubbed the “Probe Blue Book,” containing the high-level characteristics of past entry systems and easily tucked into a daily 

planner, notebook, or other carry-along work aid. 

Now that cell phones with global internet access are here, seemingly the era of the Probe Blue Book has passed. Yet 

ironically, we now suffer too much information. To obtain that one useful, comparative number becomes an exercise of 

selecting judicious search terms then reconciling conflicting results. Persisting in its convenient form and function, this 4th

version contains an expanded set of missions as well as additional information on each mission. Inevitably, there are 

ambiguities and uncertainties in the data contained herein. There are often differences between design vs. construction, 

nominal vs. upper limit, and directly measured vs. inferred. In general, preference is given to best estimates of as-built 

construction and as-flown conditions. In some cases, references are cited. In other cases, the editors made best estimates. In 

all cases, an attempt is made to state the provenance of the listed value. References are included as an entry point into the

publication record. 

The editors would be appreciative of suggested corrections or clarifications to improve future versions of the guide. While it 

must be acknowledged that an Entry Vehicle is merely a tool (usually disposable!) toward achieving a greater goal of scientific 

discovery or expanding human presence, still there is excitement in the vehicle itself; our own shooting star containing and 

preserving the wishes of our collective human enterprise of discovery. 

P L A N E T A R Y M I S S I O N E N T R Y V E H I C L E S QUICK REFERENCE GUIDE • VERSION 4.1

— ii —

Dean Kontinos

Chief Engineer

NASA Ames Research Center
Explanation of data types:

Best-estimated: Derived from information on the as-built and as-flown spacecraft configuration and mission. These values are 

based on any available flight telemetry along with Best-Estimated Trajectory (BET) and follow-on simulation data.

Design: Derived from information from spacecraft project design documents. These may deviate from what a best-estimated 

value would be due to design unknowns and conservatisms.

Editor’s Estimate: Derived from editor research and simulations, where design information is unavailable or in conflict. This is a 

credible engineering estimate.

Explanation of symbols used for each data type:

P L A N E T A R Y M I S S I O N E N T R Y V E H I C L E S QUICK REFERENCE GUIDE • VERSION 4.1

— iii —

Geometry 

Use 123 for Best-estimated

[123] for design

~123 for editor's estimate

Aerothermal 

Use 123 for Best-estimated

[123] for design

~123 for editor's estimate

TPS

Use 123 for Best-estimated

[123] for design 

~123 for editor's estimate

Use (1), (2), (3) to designate 

multiple materials

Parachutes

Use 123 for Best-estimated

[123] for design 

~123 for editor's estimate 

Use (1), (2), (3) to indicate 

multiple parachute stages, as 

per Table 3 of AIAA 2006-

6792
— iv —

Table of Contents

P L A N E T A R Y M I S S I O N E N T R Y V E H I C L E S QUICK REFERENCE GUIDE • VERSION 4.1

MISSION 

FIREII.......................................MAY 1965..........1

Apollo AS-201.............................FEB 1966.......... 2

Apollo AS-202............................ AUG 1966..........3

Apollo 4..................................... NOV 1967..........4

Apollo 6.................................... APR 1968.......... 5

Reentry F.................................. APR 1968.......... 6

PAET........................................ JUN 1971.......... 7

Viking Lander 1.......................... AUG 1975..........8

Viking Lander 2.......................... SEP 1975.......... 9

Pioneer-Venus: “North Probe”....... AUG 1978..........10

Pioneer-Venus: “Night Probe”........ AUG 1978..........11

Pioneer-Venus: “Day Probe”......... AUG 1978..........12

Pioneer-Venus: “Sounder Probe”... AUG 1978..........13

Galileo...................................... OCT 1989..........14

OREX.......................................FEB 1994.......... 15

Pathfinder..................................DEC 1996..........16

Mirka......................................... OCT 1997..........17

Huygens....................................OCT 1997..........18

ARD..........................................OCT 1998..........19

Deep Space 2............................. JAN 1999.......... 20

Stardust ................................... FEB 1999.......... 21

LAUNCH DATE PAGE

Genesis..................................... SEP 2001.......... 22

Hayabusa.................................. MAY 2003..........23

Beagle 2.....................................JUN 2003.......... 24

Mars Exploration Rovers................JUN 2003.......... 25

Phoenix.................................... AUG 2009...........26

MSL ......................................... NOV 2011..........27

IRVE ........................................ JUL 2012.......... 28

Hayabusa 2................................ DEC 2014..........29

Exploration Flight Test-1 ...............DEC 2014..........30

ExoMars 2016............................. MAR 2016..........31

OSIRIS-REx...............................SEP 2016.......... 32

InSight...................................... MAY 2018..........33

Mars2020...................................JUL 2020 ...........34

LOFTID......................................NOV 2022..........35

Artemis I.....................................NOV 2022..........36

Appendix 1: List of Space Vehicles and Missions....... 37

Appendix 2: Definitions......................................... 37

Appendix 3: Unit Conversion Factors....................... 39

Appendix 4: References........................................39

MISSION LAUNCH DATE PAGE
Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: 

-14.24° Shape spherical 

section

Velocity at 

peak heat 10.2 km/s

Forebody 

material 

designation

See note above Type N/A

Entry velocity: 

inertial & 

relative

Inertial: 

11.74 km/s

Relative: 

11.35 km/s

Nose 

radius

0.935 m

0.805 m

0.702 m

Peak 

convective 

heating

~920 W/cm2

Forebody 

thickness & 

mass

0.3 cm Deployment 

method N/A

Trim L/D 

(specify trim α)

0° (spin 

stabilized) Base area

0.354 m2

0.312 m2

0.271 m2

Peak 

radiative 

heating

~350 W/cm2 (.2-.4μ)

implying ~650 total

Traj value 487 W/cm2

Ablating

Ejected

Fore: ablated and 

melted 

Aft: ablated

Reference 

diameter / 

area

N/A

Control method

Ballistic, 

spin 

stabilized

Vehicle 

mass 86.57 kg

Integrated 

total 

heatload

16136 W/cm2

TPS 

Integration 

method

Nested/ layered 

TPS

Deployment 

Mach N/A

Ballistic co-eff. 164.1 kg/m2 Payload 

mass N/A PH stag. pressure

1.046 atm

Aftbody

material 

designation

Phenolic-asbestos 

and silicon 

elastomer

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration ~77.2 g

TPS mass 

fraction, 

inc. insul.

Peak stag. 

heating rate 1339 W/cm2

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

MISSION: FIRE II

PLANET: EARTH

LAUNCH: MAY 22, 1965 

ENTRY: MAY 22, 1965

MISSION DESCRIPTION: Pile driver flight test of a scaled Apollo CM geometry to 

measure reentry heating environment. Launched with Atlas D LV then pile driven into 

atmosphere to high speed with Antares II-A5 rocket motor.

INSTRUMENTATION: 3 forebody calorimeters, 11 forebody thermocouples,12 offset 

radiometer thermocouples and one static pressure transducer on the afterbody. 

NOTES: This aerothermal flight test was to evaluate radiative heating for Apollo. 

The reentry package consisted of three nested phenolic-asbestos heatshields 

sandwiched between three beryllium calorimeters that ablated or melted away, 

therefore the mass and OML changed with time.

— 1 —
MISSION DESCRIPTION: First unmanned suborbital flight to test the 

Saturn 1B launch vehicle, and the command and service modules.

INSTRUMENTATION: 36 pressure sensors all worked; 35 calorimeters 

worked initially.

NOTES: TPS thickness: Ablator = 4.32 cm, braised stainless-steel 

substructure (PH 15-7 MO) = 5.08 cm; Insulation: (TG-15,000) = 

2.03 cm, aluminum honeycomb (2014-T6 and 5052-H39) = 3.81 cm 

Peak heating is not at stagnation point. Manufacturer: AVCO Corp.

— 2 —

MISSION: APOLLO AS-201 

PLANET: EARTH

LAUNCH: FEB 26, 1966 

ENTRY: FEB 26, 1966 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: 

-8.58°

Relative: 

-9.03°

Shape

Capsule: 

33°

cone

Velocity at 

peak heat

5.73 

km/s

Forebody 

material 

designation

Avco 5026-39 HC Type

(1) 2 conical ribbon drogue 

parachutes, (2) 3 ringshot

pilot parachutes, (3) 3 

ringsail main parachutes

Entry 

velocity: 

inertial & 

relative

Relative: 

7.67 

km/s

Nose 

radius

4.69 m, 

3 m 

effective

Peak 

convective 

heating

186 

W/cm2 

at 

peak

Forebody 

thickness & 

mass

Ablator = 4.32 cm, 

braised stainless-steel 

substructure 

(PH 15-7 MO)=5.08 cm

Deployment 

method

(1) mortar (2) drogue 

parachute (3) pilot 

parachute

Trim L/D 

(specify

trim α)

-19.7°

< α < 

21.3°

Base 

area

12.02 

m2

Peak 

radiative 

heating

0 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: no

Reference 

diameter / 

area

(1) 4.9 m dia. (2) 2.2 m dia. 

(3) 25.5 m dia.

Control 

method

No 

control: 

rolled

Vehicle 

mass

5027.48 

kg

Integrated 

total 

heatload

7600 

J/cm2

TPS 

integration 

method

Honeycomb bonded to 

substructure; cells filled 

w/ablative compound

Deployment 

Mach ~0.7

Ballistic 

co-eff.

Payload 

mass N/A PH stag. pressure

0.85 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

Peak 

deceleration 14.3 g

TPS 

mass 

fraction, 

inc. insul.

13.7%

Peak stag. 

heating 

rate

186 

W/cm2 

at 

peak

Aftbody

thickness & 

mass

Parachute 

materials

Parachutes
MISSION: APOLLO AS-202 (APOLLO-3)

PLANET: EARTH

LAUNCH: AUG 25, 1966 

ENTRY: AUG 25, 1966 

MISSION DESCRIPTION: Second unmanned suborbital flight to test the Saturn 1B 

launch vehicle, and the command and service modules. 

INSTRUMENTATION: 36 pressure sensors all worked; 35 calorimeters worked initially. 

NOTES: TPS thickness: Ablator = 4.32 cm, braised stainless-steel substructure (PH 

15-7 MO) = 5.08 cm. Insulation: (TG-15000) = 2.03 cm, aluminum honeycomb 

(2014-T6 and 5052-H39) = 3.81 cm Manufacture: AVCO Corp.

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial -3.57°

Relative -3.74°

Entry altitude 

121.67 km

Shape Capsule 

33°cone

Velocity at 

peak heat

7.94 

km/s

Forebody 

material 

designation

Avco 5026-39 HC Type

(1) 2 conical ribbon drogue 

parachutes, (2) 3 ringshot

pilot parachutes, (3) 3 

ringsail main parachutes

Entry 

velocity: 

inertial & 

relative

Inertial: 8.75°

km/s

Relative: 8.35 

km/s

Nose 

radius

4.59 m, 

3 m 

effective

Peak 

convective 

heating

59.2 

W/cm2 

at 

peak

Forebody 

thickness & 

mass

Ablator = 4.32 cm 

braised stainlesssteel substructure 

(PH 15-7 MO) = 

5.08 cm

Deployment 

method

(1) mortar (2) drogue 

parachute (3) pilot 

parachute

Trim L/D 

(specify 

trim α)

0.372 Avg. 

Resultant L/D

Base 

area 12.2 m2

Peak 

radiative 

heating

0 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: no

Reference 

diameter / 

area

(1) 4.9 m dia. (2) 2.2 m 

dia. (3) 25.5 m dia.

Control 

method Roll modulation Vehicle mass

5424.51 

kg

Integrated 

total 

heatload

18789 

J/cm2

TPS 

integration 

method

Honeycomb bonded 

to substructure; 

cells filled with 

ablative compound

Deployment 

Mach ~0.7

Ballistic 

co-eff. N/A Payload mass

Apollo 

CM test 

flight

PH stag. 

pressure

0.11 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

Peak 

deceleration

TPS 

mass 

fraction, 

inc. 

insul.

13.7%

Peak stag. 

heating 

rate

59.2 

W/cm2 

at 

peak

Aftbody

thickness & 

mass

Parachute 

materials

Parachutes

— 3 —
MISSION: APOLLO AS-501 (APOLLO-4)

PLANET: EARTH

LAUNCH: NOV 9, 1967 

ENTRY: NOV 9, 1967 INSTRUMENTATION: 17 pressure sensors all worked, and 23 calorimeters worked 

initially. Radiometer functioned well. 

MISSION DESCRIPTION: Test of Saturn V launch vehicle and overall reentry operations. 

NOTES: TPS thickness: Ablator = 4.32 cm, braised stainless-steel substructure (PH 

15-7 MO) = 5.08 cm. Insulation: (TG-15,000) = 2.03 cm, aluminum honeycomb (2014-

T6 and 5052-H39) = 3.81 cm Manufacturer: AVCO Corp. 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial -6.93 

Relative-7.19° 

Entry altitude 

121.92 km

Shape

Capsule: 

33°

cone

Velocity at 

peak heat

10.04 

km/s

Forebody 

material 

designation

Avco 5026-39 HC Type

(1) 2 conical ribbon 

drogue parachutes, 

(2) 3 ringshot pilot 

parachutes, (3) 3 ringsail

main parachutes

Entry 

velocity: 

inertial & 

relative

Inertial: 11.14 

km/s

Relative: 10.73 

km/s

Nose 

radius

4.66 m, 

3 m 

effective

Peak 

convective 

heating

219 

W/cm2

Forebody 

thickness

& mass

TPS thickness: 

Ablator = 4.32 cm, 

braised stainless-steel 

substructure (PH 15-7 

MO) = 5.08 cm

Deployment 

method

(1) mortar (2) drogue 

parachute (3) pilot 

parachute

Trim L/D 

(specify 

trim α)

0.37 < L/D < 

0.44 

24° < α <28°

Base area 12.08 m2

Peak 

radiative 

heating

317 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: no

Reference 

diameter / 

area

(1) 4.9 m dia. (2) 2.2 m 

dia. (3) 25.5 m dia.

Control 

method

Roll 

modulation

Vehicle 

mass

5424.5 

kg

Integrated 

total 

heatload

24122 

J/cm2

TPS 

integration 

method

Honeycomb bonded 

to substructure; cells 

filled with ablative 

compound

Deployment 

Mach ~0.7

Ballistic coeff. ~340 kg/m2 Payload 

mass N/A PH stag. pressure

0.542 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

Peak 

deceleration 8.79 g

TPS 

mass 

fraction, 

inc.insul.

13.7%

Peak stag. 

heating 

rate

527 

W/cm2 

peak

Aftbody

thickness & 

mass

Parachute 

materials

Parachutes

— 4 —
MISSION: APOLLO AS-502 (APOLLO-6)

PLANET: EARTH

LAUNCH: APR 4, 1968 

ENTRY: APR 4, 1968 

MISSION DESCRIPTION: Final qualification test for launch vehicle and 

command module for manned mission. 

NOTES: TPS thickness: Ablator = 4.32 cm, braised stainless-steel 

substructure (PH 15-7 MO) = 5.08 cm Insulation: (TG-15,000) = 2.03 cm, 

aluminum honeycomb (2014-T6 and 5052-H39) = 3.81 cm.

Manufacture: AVCO Corp Saturn-V malfunctioned causing the CM to 

follow a spurious return trajectory.

Trajectory Geometry Aero/thermal TPS

Entry 

angle

Inertial: 

-5.9° Shape Capsule: 

33° cone

Velocity at 

peak heat

8.32 

km/s

Forebody 

material 

designation

Avco 5026-39 HC Type

(1) 2 conical ribbon 

drogue parachutes, 

(2) 3 ringshot pilot 

parachutes, (3) 3 ringsail

main parachutes

Entry 

velocity: 

inertial 

& 

relative

Inertial: 10.00 

km/s

Relative: 9.60 

km/s

Nose 

radius

4.69 m, 

3 m 

effective

Peak 

convective 

heating

197 

W/cm2

Forebody 

thickness & 

mass

TPS thickness: Ablator = 

4.32 cm, braised 

stainless-steel 

substructure 

(PH 15-7 MO) = 5.08 cm

Deployment 

method

(1) mortar (2) drogue 

parachute (3) pilot 

parachute

Trim L/D 

(specify 

trim α)

0.35 < L/D 

<0.4

24°< α <28°

Base 

area 12.02 m2

Peak 

radiative 

heating

43 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: no

Reference 

diameter / 

area

(1) 4.9 m dia. (2) 2.2 m 

dia. (3) 25.5 m dia.

Control 

method

Roll 

modulation

Vehicle 

mass 5424.9 kg

Integrated 

total 

heatload

32000 

J/cm2

TPS 

integration 

method

Honeycomb bonded to 

substructure; cells filled 

with ablative compound

Deployment 

Mach ~0.7

Ballistic 

co-eff. 395.8 kg/m2 Payload mass N/A PH stag. pressure

0.354 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

Peak 

deceleration

TPS 

mass 

fraction, 

inc.insul.

13.7%

Peak stag. 

heating 

rate

240 

W/cm2 

peak

Aftbody

thickness & 

mass

Parachute 

materials

Parachutes

INSTRUMENTATION:

— 5 —
Entry angle -20.78° Shape

5°half 

cone, 

3.962 m 

long

Velocity at 

peak heat 5.47 km/s Forebody material designation

Nose-tip: ATJ 

graphite 

Body: beryllium

Type N/A

Entry 

velocity: 

inertial & 

relative

Inertial: 

6.28 km/s

Relative: 

5.96 km/s

Nose 

radius 0.254 cm

Peak 

convective 

heating

32 W/cm2 Forebody thickness 

& mass Be: 1.524 cm Deployment method N/A

Trim L/D 

(specify 

trim α)

0°

ballistic Base area 0.3772 m2 Peak radiative 

heating ~0 W/cm2 AblatingEjected

Nose-tip: yes 

Ejected: no

Reference 

diameter / area N/A

Control 

method N/A Vehicle mass 272 kg Integrated total heatload

TPS integration 

method

Deployment 

Mach N/A

Ballistic coeff.

~50000 

kg/m2

Payload 

mass N/A PH stag. pressure

250 atm Aftbody material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration

TPS mass 

fraction, 

inc. insul.

Peak stag. 

heating rate 32 W/cm2 Aftbody thickness & mass

Parachute 

materials N/A

MISSION: REENTRY F

PLANET: EARTH

LAUNCH: APR 27, 1968 

ENTRY: APR 27, 1968 

MISSION DESCRIPTION: To measure turbulent heating rates and 

transition onset in a flight environment.

NOTES: Nose-tip heating rate is not relevant for this flight, which 

was designed to measure heating on a sharp cone. The nose tip was 

meant to ablate during entry 3. The beryllium heat shield melted for 

~40 seconds after entry.

Trajectory Geometry Aero/thermal TPS Parachutes

INSTRUMENTATION: Multiple thermocouples and pressure 

sensors at 21 stations on cone, 4 heat-flux gauges and 2 pressure 

gauges on base. 3 thermocouples in nose-tip assembly.

— 6 —
MISSION DESCRIPTION: To test the capability to determine the 

composition of unknown atmospheres during high-speed entry.

INSTRUMENTATION: Multiple thermocouples and pressure sensors 

at 21 stations on cone, 4 heat-flux gauges and 2 pressure gauges on 

base. 3 thermocouples in nose-tip assembly.

MISSION: PAET

PLANET: EARTH

LAUNCH: JUL 2, 1971 

ENTRY: JUL 2, 1971 

Trajectory Geometry Aero/thermal TPS

Entry angle

Relative: -41°

Inertial: 

-38.98°

93 km altitude

Shape

Blunt-nose, 

55° halfcone angle

Velocity at 

peak heat 5.7 km/s

Forebody 

material 

designation

Nose: Beryllium heatsink 

Conical frustum: ESA 3560 

ablator

Type N/A

Entry velocity: 

inertial & relative

Inertial: 6.88 

km/s Relative: 

6.66 km/s

Nose 

radius 0.457 m

Peak 

convective 

heating

~228 

W/cm2

Forebody 

thickness & 

mass

Nose: 1-2.5 cm 

Conical frustum: 0.76 cm

Deployment 

method N/A

Trim L/D

(specify trim α) 0 Base area 0.66 m2

Peak 

radiative 

heating

Negligible Ablating

Ejected

Frustum ablated 

Ejected: no

Reference 

diameter / 

area

N/A

Control method Ballistic Vehicle 

mass 62.1 kg

Integrated 

total 

heatload

Stag. pt. 

1982 

J/cm2

TPS 

integration 

method

None (solid Be) Deployment 

Mach N/A

Ballistic co-eff. 68.4 kg/m2 Payload 

mass 14 kg PH stag. pressure

0.60 atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 76 g

TPS 

mass 

fraction, 

inc. insul.

Forebody: 

13.7% 

Afterbody: 

3.5%

Peak stag. 

heating 

rate

Max: 228 

W/cm2 

(no 

ablation) 

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

— 7 —

Shadowgraph 

of PAET Model 

in Ballistic 

Range Test

NOTES:
MISSION: VIKING LANDER 1

PLANET: MARS

LAUNCH: AUG 20, 1975 

ENTRY: JUL 20, 1976 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: -17.00°

Relative: -17.76° 

Entry altitude 242.8 km 

Shape

70°

sphere

-cone

Velocity at 

peak heat

4.06 

km/s

Forebody 

material 

designation

SLA561-V Type Single DGB 

parachute

Entry 

velocity: 

inertial & 

relative

Inertial: 4.610 km/s

Relative: 4.418 km/s 

Nose 

radius

0.8763 

m

Peak 

convective 

heating

21.1 

W/cm2

Forebody 

thickness & 

mass

Variable: (max) 1.38 

cm

Deployment 

method

Main parachute 

directly deployed 

by 15 in. 

diameter mortar

Trim L/D 

(specify 

trim α)

α = -11.1° Base area

9.84 

m2

Peak 

radiative 

heating

0 Ablating

Ejected

Ablator-pyrolyzed

Ejected: yes

Reference 

diameter / 

area

16.2 m dia.

Control 

method

3-axis RCS 

rate limiting

Vehicle 

mass

982.9 

kg

Integrated 

total 

heatload

~1001 

J/cm2

TPS 

integration 

method

Honeycomb bonded 

to substructure; cells 

filled with ablative 

compound

Deployment 

Mach ~1.1

Ballistic 

co-eff.

At peak dyn. p.: 

60.41 kg/m2

Payload 

mass N/A

Peak 

heating 

stag. 

pressure

0.06 

atm

Aftbody

material 

designation

Inner cone: laminated 

glass fabric/phenolic 

resin; outer cone: 

aluminum

Deployment 

dynamic 

pressure

~380 Pa

Peak 

deceleration ~6.86 g 

TPS mass 

fraction, 

inc. insul.

2.8%

Peak stag. 

heating 

rate

Peak: 

21.1 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials Dacron

Parachutes

— 8 —

MISSION DESCRIPTION: To characterize the structure and composition of the 

atmosphere and surface of Mars. 

INSTRUMENTATION: The forebody aeroshell was not instrumented, but the wake 

enclosure (backshell) had thermocouples. There was one pressure port off stagnation 

point and one on the base cover. Temperature gauges were on the back face and on 

both backshell frustums.

NOTES: Resin Material: silicone elastomer with glass microspheres and cork; Matrix 

material: fiberglass-phenolic honeycomb. RCS was used to maintain trim angle of attack.
MISSION: VIKING LANDER 2

PLANET: MARS

LAUNCH: SEP 9, 1975 

ENTRY: SEP 3, 1976 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: -17.08° 

Relative: -17.62° 

Entry altitude 240.99 km

Shape

Blunt-nosed 

70° half 

cone

Velocity at 

peak heat

4.1 

km/s

Forebody 

material 

designation

SLA561-V Type Single GBD 

parachute

Entry 

velocity: 

inertial & 

relative

Inertial: 4.61 km/s

Relative: 4.48 km/s

Nose 

radius 0.876 m

Peak 

convective 

heating

~21.95 

W/cm2

Forebody 

thickness & 

mass

Variable: (max) 

1.38 cm

Deployment 

method

Main parachute 

directly deployed 

by 15 in. 

diameter mortar

Trim L/D 

(specify 

trim α)

0.18 at α = -11.3° Base area 9.84 m2

Peak 

radiative 

heating

0 Ablating

Ejected

Ablator-pyrolyzed

Ejected: yes

Reference 

diameter / 

area

16.2 m dia.

Control 

method

3-axis RCS 

rate limiting

Vehicle 

mass 981.6 kg

Integrated 

total 

heatload

~1043 

J/cm2

TPS 

integration 

method

Honeycomb 

bonded to 

substructure; cells 

filled with ablative 

compound

Deployment 

Mach ~1.1

Ballistic

co-eff. 61.44 kg/m2 Payload mass N/A PH stag. pressure

~0.041 

atm 

Aftbody

material 

designation

Inner cone: 

laminated glass 

fabric/phenolic 

resin; outer cone: 

aluminum

Deployment 

dynamic 

pressure

~408 pa

Peak 

deceleration

~7.2 g trajectory

~8.5 g chute snap

TPS mass 

fraction, 

inc. insul.

2.8%

Peak stag. 

heating 

rate

Max: 

21.95 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials Dacron

Parachutes

— 9 —

MISSION DESCRIPTION: To characterize the structure and composition of the 

atmosphere and surface of Mars.

INSTRUMENTATION: The forebody aeroshell was not instrumented, but the wake 

enclosure (backshell) had thermocouples. There was one pressure port off stagnation 

point and one on the base cover. Temperature gauges were on the back face and on 

both backshell frustums.

NOTES: Resin material: silicone elastomer with glass microspheres and cork. Matrix 

material: fiberglass-phenolic honeycomb. RCS was used to maintain trim angle of attack. 
MISSION: PIONEER-VENUS: SMALL ‘NORTH PROBE’

PLANET: VENUS

LAUNCH: AUG 8, 1978 

ENTRY: : DEC 9, 1978 

MISSION DESCRIPTION: A 60° N day entry to map atmosphere, 

including characterizing wind and turbulence.

INSTRUMENTATION: Thermocouples: one at 17 ° off stagnation point (0.41 cm 

below heat-shield surface); another on conical frustum ahead of shoulder (0.30 cm 

below heat-shield surface) at s/Rn = 2.2.

NOTES: Heating rates and loads are based on Tauber correlations in radiative 

equilibrium (very model sensitive). Pioneer-Venus "North Probe" experienced 466 g 

which was greatest for a successful probe.

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial & 

Relative: -68.74°

@ 200 km

Shape

Blunt-nosed, 

45° halfcone angle

Velocity at 

peak heat

10.83 

km/s

Forebody 

material 

designation

Carbon-phenolic FM 

5055 Type N/A

Entry 

velocity: 

inertial & 

relative

Inertial & 

Relative: 11.54 

km/s 

Nose 

radius 0.19 m

Peak 

convective 

heating

~3085 

W/cm2

Forebody 

thickness & 

mass

1.2 cm at 

stagnation point

Deployment 

method N/A

Trim L/D 

(specify

trim α)

0 Base area 0.46 m2 Peak radiative 

heating

~3273 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

N/A

Control 

method Ballistic Vehicle mass 91 kg Integrated total heatload

At stag. pt

~10502 

J/cm2

TPS 

integration 

method

Nose: chopped molded; 

conical section: tapewrapped

Deployment 

Mach N/A

Ballistic

co-eff. 190 kg/m2 Payload mass 3.60 kg Peak heating stag. pressure

~16.5 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 465.9 g

TPS mass 

fraction, 

inc. insul.

12.9%

Peak stag. 

heating rate 6150 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

— 10 —
MISSION: PIONEER-VENUS: SMALL ‘NIGHT PROBE’

PLANET: VENUS

LAUNCH: AUG 8, 1978 

ENTRY: : DEC 9, 1978 

INSTRUMENTATION: Thermocouples: one at 17° off stagnation point (0.41 cm 

below heat-shield surface); another on conical frustum ahead of shoulder (0.30 cm 

below heat-shield surface) at s/Rn = 2.2.

NOTES: Heating rates and loads are probably for non-ablating conditions. 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial & 

Relative: -41.50° 

@ 200 km

Shape

Blunt-nosed, 

45.2° halfcone angle

Velocity at 

peak heat ~10.79 km/s

Forebody 

material 

designation

Carbon-phenolic 

FM 5055 Type N/A

Entry 

velocity: 

Inertial & 

Relative

Inertial & 

Relative: 11.54

Nose 

radius 0.19 m

Peak 

convective 

heating

~2743 

W/cm2

Forebody 

thickness & 

mass

1.2 cm 

at stagnation point

Deployment 

method N/A

Trim L/D 

(specify

trim α)

0 Base area 0.46 m2

Peak 

radiative 

heating

~2330 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

N/A

Control 

method Ballistic Vehicle mass 91 kg

Integrated 

total 

heatload

At stag.pt

~11413 

J/cm2

TPS 

integration 

method

Nose: chopped molded; 

conical section: tapewrapped

Deployment 

Mach N/A

Ballistic 

co-eff. 190 kg/m2 Payload mass 3.60 kg

Peak 

heating 

stag. 

pressure

~4.9 atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 353.7 g

TPS mass 

fraction, 

inc. insul.

12.9% Peak stag. 

heating rate

~4906 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

— 11 —

MISSION DESCRIPTION: To map the atmosphere, including 

temperature and pressure, from a night side entry.
MISSION: PIONEER-VENUS: SMALL ‘DAY PROBE’

PLANET: VENUS

LAUNCH: AUG 8, 1978 

ENTRY: : DEC 9, 1978 

MISSION DESCRIPTION: To map the atmosphere, including radiative 

energy, from a day side entry.

INSTRUMENTATION: Thermocouples: one at 17° off stagnation point (0.41 cm 

below heat-shield surface); another on conical frustum ahead of shoulder (0.30 cm 

below heat-shield surface) at s/Rn = 2.2. 

NOTES: Heating rates and loads are probably for non-ablating conditions. 

Day Probe survived surface impact (accidentally) or 1 hour, 7.58 minutes.

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: -25.41°

Relative: -25.42°

@ 200 km

Shape

Blunt-nosed, 

45° halfcone angle

Velocity at 

peak heat 10.68 km/s

Forebody 

material 

designation

Carbon-phenolic FM 

5055 Type N/A

Entry 

velocity: 

inertial & 

relative

Inertial & Relative: 

11.54 km/s 

Nose 

radius 0.19 m

Peak 

convective 

heating

~2250 

W/cm2

Forebody 

thickness & 

mass

1.2 cm at stagnation 

point

Deployment 

method N/A

Trim L/D 

(specify 

trim α)

0 Base area 0.46 m2 Peak radiative 

heating

~1363 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

N/A

Control 

method Ballistic Vehicle mass 91 kg Integrated total heatload

At stag.pt

~13264 

J/cm2

TPS 

integration 

method

Nose: chopped 

molded; conical 

section: tape-wrapped

Deployment 

Mach N/A

Ballistic 

co-eff. 190 kg/m2 Payload mass 3.60 kg Peak heating stag. pressure ~8.1 atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 227.7 g

TPS mass 

fraction, 

inc. insul.

12.9% Peak stag. 

heating rate

3476 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

— 12 —
MISSION: PIONEER-VENUS: LARGE PROBE ‘SOUNDER’

PLANET: VENUS

LAUNCH: AUG 8, 1978 

ENTRY: DEC 9, 1978 

MISSION DESCRIPTION: This probe contained 7 experiments, including 1 to measure 

the atmospheric composition.

INSTRUMENTATION: Thermocouples: one at 17° off stagnation point (0.41 cm below 

heat-shield surface); another on conical frustum ahead of shoulder (0.30 cm below heatshield surface) at s/Rn = 2.2. 

NOTES: Heating rates and loads are probably for non-ablating conditions. 

Parachute deployed at 66.46 km to extract probe from heat shield. Parachute detached from 

probe at 1076 sec after entry Probe impacted surface (signal ended) 3261 sec. after entry.

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial & Relative: 

-32.37°

@ 200 km

Shape

Bluntnosed, 

45° halfcone angle

Velocity at 

peak heat

10.82 

km/s

Forebody 

material 

designation

Carbonphenolic FM 

5055

Type

(1) Ribless guide surface, 

8 gore pilot chute, (2) 

Conical ribbon, 18 gore 

main chute

Entry 

velocity: 

inertial & 

relative

Inertial & Relative:

11.54km/s 

@ 200 km

Nose 

radius 0.36 m

Peak 

convective 

heating

~1813 

W/cm2

Forebody 

thickness & 

mass

1.60 cm at 

stagnation 

point

Deployment 

method (1) Mortar (2) Pilot chute

Trim L/D 

(specify 

trim α)

0 Base 

area 1.59 m2

Peak 

radiative 

heating

~2468 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

(1) 0.76 m dia. (CdA=0.37 

m2) (2) 4.94 m dia. 

(CdA=10.4 m2) 

Control 

method Ballistic Vehicle mass 316.483 kg

Integrated 

total 

heatload

At 

stag.pt

~11305 

J/cm2

TPS 

integration 

method

Nose: chopped 

molded; conical 

section: tapewrapped

Deployment 

Mach ~0.82

Ballistic 

co-eff. 189.7 kg/m2 Payload mass

Science 

instr.. 29.15 

kg (9.2%)

Peak heating 

stag. 

pressure

~10.2 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

~3287 pa

Peak 

deceleration 286.4 g

TPS 

mass 

fraction 

inc.insu.

Forebody: 

8.83% Aft 

cover:1.52%

Peak stag. 

heating rate

4145 

W/cm

Aftbody

thickness & 

mass

Parachute 

materials

(1) Dacron 1.25 oz/yd2

(2) Dacron 2.25 oz/yd2

Parachutes

— 13 —
MISSION DESCRIPTION: To descend into the Jovian atmosphere, collect atmospheric data 

and relay to the orbiter. 

INSTRUMENTATION: Forebody TPS: ablation recession gauges. Afterbody TPS: 

thermocouples in the nylon phenolic. Galileo was the most difficult entry in the history of 

space exportation. Attempts to model Galileo's entry are little more than educated guesses 

with many phenomena poorly understood. 

NOTES: Reported CG estimates varied widely. "Blockage" refers to reduction in outer wall 

heating due to ablation effects.

— 14 —

MISSION: GALILEO

PLANET: JUPITER

LAUNCH: OCT 18, 1989 

ENTRY: DEC 7, 1995 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: -6.65°

Relative: -8.41°

@ 450 km

Shape

Blunt-nosed 

44.86° halfcone angle

Velocity at 

peak heat 39.4 km/s

Forebody 

material 

designation

Carbonphenolic Type (1) Pilot: 20° conic ribbon 

(2) Main: 20° conic ribbon 

Entry velocity: 

inertial & 

relative

Inertial: 59.83 km/s

Relative: 47.41 

km/s

Nose 

radius 0.222 m

Peak 

convective 

heating

Forebody 

thickness & 

mass

14.6 cm at 

stagnation

Deployment 

method

(1) Mortar launched with aft 

cover (2) Pilot parachute

Trim L/D 

(specify

trim α)

0 Base 

area 1.26 m2

Peak 

radiative 

heating

~38055 W/cm2 

[unblocked]

~16272 W/cm2

[blocked]

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

(1) 1.14 m dia. (CdA=0.51 m2) 

(2) 3.8 m dia. (CdA=5.46 m2)

Control 

method Ballistic Vehicle mass

At entry: 

338.93 kg

Integrated 

total 

heatload

~210826 J/cm2 

with ablation

TPS 

integration 

method

Nose: chopped 

molded; 

conical 

section: tapewrapped

Deployment 

Mach

g-switch malfunctioned; chute 

deployed at 16.6 km above 

one bar altitude 

Ballistic 

co-eff. ~267.6 kg/m2 Payload mass Science 8.3% PH stag. pressure 7.55 atm

Aftbody

material 

designation

Phenolic-nylon

Deployment 

dynamic 

pressure

See above

Peak 

deceleration ~228 g

TPS 

mass 

fraction, 

inc. 

insul.

Forebody: 

45.4% 

Afterbody: 5%

Peak stag.

heating 

rate

16858 W/cm2

[unblocked]

3286 W/cm2

[blocked]

Aftbody

thickness & 

mass

Parachute 

materials Dacron

Parachutes
MISSION DESCRIPTION: To collect information on the design of a reentry vehicle to 

support Japanese unmanned space shuttle HOPE.

INSTRUMENTATION: Wall catalycity measurement, electrostatic probe, and heat 

shield temperature sensors. 

NOTES: RCS was used to maintain a trim angle of attack of zero. 

— 15 —

MISSION: OREX

PLANET: EARTH RETURN

LAUNCH: FEB 4, 1994

ENTRY: FEB 4, 1994

Trajectory Geometry Aero/thermal TPS

Entry angle

Relative 

-3.17° Shape

50°

spherecone

Velocity at peak 

heat 6.4 km/s Forebody material designation

Si Coated C-C 

for nose Type N/A

Entry 

velocity: 

inertial & 

relative

Inertial: 

7.8 km/s

Relative: 

7.43 km/s

Nose 

radius 1.35 m

Peak 

convective 

heating

51 W/cm2 Forebody 

thickness & mass 4 cm Deployment method N/A

Trim L/D 

(specify 

trim α)

0 Base area 9.08 m2 Peak radiative 

heating 0 AblatingEjected

Ablated: no 

Ejected: no

Reference 

diameter / area N/A

Control 

method RCS Vehicle mass

761 kg at 

entry

Integrated total 

heatload

TPS integration 

method Tiled Deployment Mach N/A

Ballistic 

co-eff. 56 kg/m2 Payload mass N/A PH stag. pressure

0.078 atm Aftbody material 

designation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration

TPS mass 

fraction, 

inc. insul.

N/A Peak stag. 

heating rate 51 W/cm2 Aftbody thickness & mass

Parachute 

materials N/A

Parachutes
MISSION DESCRIPTION: To demonstrate a simple, low-cost system for 

placing a science payload on the surface of Mars. 

INSTRUMENTATION: TPS instrumented with thermocouples only 

Rover called "Sojourner”.

MISSION: MARS PATHFINDER

PLANET: MARS

LAUNCH: DEC 4, 1996

ENTRY: JUNE 4, 1997

NOTES: Spin stabilized. 

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial -14.06°

@ 130.86 km Shape

70°

sphere-cone

Velocity at 

peak heat

Relative: 

6.62 km/s

Forebody 

material 

designation

SLA-561 V Type DGB

Entry 

velocity: 

inertial & 

relative

Inertial: 7.26 km/s

Relative: 7.48 

km/s

Nose 

radius 0.66 m

Peak 

convective 

heating

100.69 

W/cm2

Forebody 

thickness & 

mass

1.9 cm Deployment 

method

Mortar 

fired

Trim L/D 

(specify

trim α)

Ballistic 

(0° average) Base area 5.51 m2

Peak 

radiative 

heating

~5.27 W/cm2 

(with ablation)

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

12.7 m 

Control 

method Ballistic Vehicle mass

At entry: 

585.3 kg

Integrated 

total 

heatload

~3834 J/cm2 

with ablation

TPS 

integration 

method

Honeycomb 

bonded to 

substructure; cells 

filled with ablative 

compound

Deployment 

Mach 1.71

Ballistic 

co-eff.

At peak dyn. p.: 

~61.5 kg/m2

Payload 

mass N/A PH stag. pressure

~0.19 

atm

Aftbody

material 

designation

SLA-561S

Deployment 

dynamic 

pressure

588 Pa

Peak 

deceleration 16 g 

TPS mass 

fraction, 

inc. insul.

Forebody 6.2% 

Backshell 2% 

Peak stag. 

heating 

rate

105.96 

W/cm2

Aftbody

thickness & 

mass

1.27 cm Parachute 

materials

Nylon, 

polyester, 

Kevlar

Parachutes

— 16 —
MISSION DESCRIPTION: To qualify a reentry heatshield concept, to 

assess and validate relevant gas-surface-interactions models, and to assess 

tumbling processes with scientific and engineering experiments conducted 

by German researchers. 

INSTRUMENTATION: 3 acceleration sensors, 3 angular rate sensors, 24 

thermocouples, RAFLEX (pressure, temperature & heat flux sensors) and 

PYREX (pyrometric temperature measurements).

NOTES: CFRP: Carbon Fiber Reinforced Plastics. SPA: Surface Protected Ablator. 

This was the first successful Western European reentry mission.

— 17 —

MISSION: MIRKA

PLANET: EARTH

LAUNCH: NOV 9, 1997

ENTRY: NOV 23, 1997

Trajectory Geometry Aero/thermal TPS

Entry angle Relative: -2.51° Shape

Spherical 

1 m 

diameter

Velocity at 

peak heat

6.51 

km/s

Forebody 

material 

designation

CFRP: carbon fiber 

reinforced plastics SPA: 

surface protected ablator 

and fiber ceramic cover

Type unknown

Entry 

velocity: 

inertial & 

relative

Inertial: separation 

Velocity: 7.3 km/s

Relative: 7.6 km/s at 120 km

Nose 

radius 0.5 m

Peak 

convectiv

e heating

120 

W/cm2

Forebody 

thickness & 

mass

Front: 3 cm 

Back: 2 cm

Deployment 

method

Trim L/D 

(specify

trim α)

0 Base 

area 0.785 m2

Peak 

radiative 

heating

0 Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

Control 

method Ballistic Vehicle mass 154 kg

Integrated 

total 

heatload

12000 

J/cm2

TPS 

integration 

method

Deployment 

Mach

Ballistic

co-eff. 214 kg/m3 Payload mass N/A PH stag. pressure

0.178 

atm

Aftbody

material 

designation

Deployment 

dynamic 

pressure

Peak 

deceleration

TPS 

mass 

fraction, 

inc. 

insul.

36%

Peak stag. 

heating 

rate

Peak: 

120 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials

Parachutes
MISSION DESCRIPTION: To explore the atmosphere of Titan. 

INSTRUMENTATION: No aeroheating data. A mass spectrometer for 

atmospheric composition was deployed after the heat shield was ejected. 

NOTES: Huygens is a European Space Agency probe that was carried by the 

Cassini Saturn Orbiter. AQ60 silica fibers reinforced by phenolic resin. 

— 18 —

MISSION: HUYGENS

PLANET: TITAN 

(A moon of Saturn)

LAUNCH: OCT 15, 1997

ENTRY: JAN 14, 2005

Trajectory Geometry Aero/thermal TPS

Entry angle Inertial: -65.5°

Relative: -65.4° Shape

60°

spherecone

Velocity at 

peak heat ~5.1 km/s

Forebody 

material 

designation

AQ60/I Type

(1) Pilot: DGB

(2) Main: DGB

(3) Descent: DGB

Entry 

velocity: 

inertial & 

relative

Inertial: 6.0 km/s 

@ 1270 km

Relative: 6.0 km/s

Nose 

radius 1.25 m

Peak 

convective 

heating

~29.3 

W/cm2

Forebody 

thickness & 

mass

At stag: 17.4 mm 

On flank:18.2 mm

39 kg TPS, 76 kg 

w/ struct. 

Deployment 

method

(1) Mortar

(2) Pilot

(3) Main

Trim L/D 

(specify

trim α)

0 Base area 5.73 m2

Peak 

radiative 

heating

~0 W/cm2 Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

(1) 2.59 m dia.

(2) 8.31 m dia.

(3) 3.03 m dia.

Control 

method Ballistic Vehicle mass 320 kg

Integrated 

total 

heatload

~1397 

J/cm2

TPS 

integration 

method

Tiled Deployment 

Mach

(1) 1.46

(2) 1.36

(3) 0.14

Ballistic 

co-eff.

34.5-37.5 

kg/m2

Payload 

mass 44 kg PH stag. pressure

0.1 

atm

Aftbody

material 

designation

Prosial

Deployment 

dynamic 

pressure

(1) 309 Pa

(2) 275 Pa

(3) 10 Pa

Peak 

deceleration 12.36 g

TPS mass 

fraction, 

inc. insul.

Forebody: 

12.2% 

Aft: 1.64%

Peak stag. 

heating rate

~29.3 

W/cm2 

(140W/cm2 

design)

Aftbody

thickness & 

mass

0.03-0.31 cm, 5.2 kg 

TPS, 17 kg w/ 

struct.

Parachute 

materials

Nylon fabric

Kevlar structurals

Parachutes
— 19 —

MISSION: ARD 'Atmospheric Reentry Demonstrator' 

PLANET: EARTH

LAUNCH: NOV 21, 1998

ENTRY: NOV 21, 1998

Trajectory Geometry Aero/thermal TPS

Entry angle -2.6° Shape

Apollo-like 

capsule, 

33° cone

Velocity at 

peak heat

7.2 

km/s

Forebody 

material 

designation

Aleastrasil

Type

(1) Extraction parachute 

(2) drogue parachute (3) 

3 ringsail main parachutes

Entry velocity: 

Inertial & 

Relative

Inertial: 

7.86 km/s

Relative: 

7.45 km/s

Nose 

radius 3.36 m

Peak 

convective 

heating

120 

W/cm2

Forebody 

thickness & 

mass

40-65 mm Deployment 

Method

(1) Mortar

(2) Extraction parachute

(3) Drogue parachute

Trim L/D 

(specify trim α) -21.2° Base area 6.15 m2

Peak 

radiative 

heating

~0 

W/cm2

Ablating

Ejected

Ablated: 

yes

Ejected: 

no

Reference 

diameter / 

area

(1) 0.9 m dia.

(2) 5.8 m dia.

(3) 22.9 m dia.

Control 

method

RCS: 7 

thrusters

Vehicle 

mass

2715 kg 

(inc.1017 kg 

front heat 

shield)

Integrated 

total 

heatload

~24000 

J/cm2

TPS 

integration 

method

Tiled Deployment 

Mach ~0.8

Ballistic co-eff.

~347 

kg/m2

Payload 

mass N/A PH stag. pressure

0.22 

atm

Aftbody

material 

designation

Norcoat

622-50FI 

(See note)

Deployment 

dynamic 

pressure

Peak 

deceleration ~11 g

TPS 

mass 

fraction, 

inc. insul.

23% 

(626 kg 

TPS)

Peak stag. 

heating 

rate

120 

W/cm2

Aftbody

thickness & 

mass

19 mm Parachute 

materials

Parachutes

MISSION DESCRIPTION: To undertake a complete space flight cycle for ESA, 

with emphasis on reentry technologies. 

INSTRUMENTATION: The capsule afterbody was instrumented with 7 surfaces 

pressure sensors, 2 thermal plugs with 2 thermocouples each on the back cover, and 

4 surface-mounted copper calorimeters on the cylindrical section. The front cone 

contained 18 pressure sensors, 14 thermal plugs with 3 or 5 TC each.

NOTES: Aleastril: silica fibers with phenolic resin; Norcoat: cork powder and phenolic 

resin. 4 experimental Ceramic Matrix Composite (CMC) tiles and samples of Flexible 

External insulation (FEI). 
MISSION: DEEP SPACE 2

PLANET: MARS

LAUNCH: JAN 3, 1999

ENTRY: DEC 3, 1999

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial -13.25°

@128 km Shape

45° sphere-cone, 

spherical aft

Velocity at 

peak heat

5.94 

km/s

Forebody 

material 

designation

SIRCA-SPLIT Type N/A

Entry 

velocity: 

inertial & 

relative

Relative: 

6.9 km/s

Nose 

radius 0.0875 m

Peak 

convective 

heating

194 

W/cm2

Forebody 

thickness & 

mass

~1 cm Deployment 

method N/A

Trim L/D 

(specify

trim α)

0 Base area 0.096 m2

Peak 

radiative 

heating

N/A Ablating

Ejected

Ablated:

Ejected: no

Reference 

diameter / 

area

N/A

Control 

method Ballistic Vehicle mass 3.67 kg Integrated total heatload

8712 

J/cm2

TPS 

integration 

method

Monolithic Deployment 

Mach N/A

Ballistic

co-eff. 36.2 kg/m2 Payload mass N/A PH stag. pressure

Aftbody

material 

designation

FRCI: Fibrous 

refractory composite 

insulation

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 12.4 g

TPS mass 

fraction, 

inc. insul.

N/A Peak stag. 

heating rate

194 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials N/A

Parachutes

MISSION DESCRIPTION: To penetrate the Martian surface with two 

small probes. 

INSTRUMENTATION:

NOTES: The DS-2 aeroshells were on the failed Mars Polar Lander. 

They were to be jettisoned 5 minutes before the lander entered the 

Martian atmosphere. No signals from the probes were received. 

— 20 —
MISSION: STARDUST

PLANET: EARTH RETURN

LAUNCH: FEB 7, 1999

ENTRY: JAN 15, 2006

Trajectory Geometry Aero/thermal TPS

Entry angle -8.23° @ 134.4 km Shape

Blunt-nosed 

60° half-angle 

cone

Velocity at 

peak heat 11.1 km/s

Forebody 

material 

designation

PICA-15 Type

(1) Pilot: DGB

(2) Main: 

Triconical

Entry 

velocity: 

inertial & 

relative

Inertial: 12.8 km/s

Relative: 12.45 

km/s

Nose 

radius 0.23 m

Peak 

Convective 

heating

~817 

W/cm2

Forebody 

thickness & 

mass

5.82 cm Deployment 

method

(1) Mortar

(2) Pilot pulled

Trim L/D 

(specify

trim α)

0 Base 

area

0.52 m2, 

0.50 m2 

(ablated)

Peak 

Radiative 

heating

~71 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

(1) 0.83 m dia.

(2) 7.3 m dia.

Control 

method

Ballistic, spinstabilized at 14.5 

rpm 

Vehicle 

mass 45.8 kg

Integrated 

total 

heatload

~26210 

J/cm2

TPS 

integration 

method

PICA: singlepiece, direct 

bonded to 

substructure 

with HT-424

Deployment 

Mach (1) 1.4m +/- 0.3

Ballistic 

co-eff.

60.0 kg/m2 

60.4 kg/m2 

(ablated)

Payload 

mass N/A PH stag. pressure

0.22 

atm

Aftbody

material 

designation

SLA-561V

Deployment 

dynamic 

pressure

(1) 713 Pa 

Peak 

deceleration 32.89 g

TPS 

mass 

fraction, 

inc. insul.

22% Peak stag. 

heating rate

~880 

W/cm2

Aftbody

thickness & 

mass

Design: 1.397 

cm

Parachute 

materials

Canopy: nylon, 

Kevlar 

suspension 

lines

Parachutes

MISSION DESCRIPTION: To collect comet material from Wild 2 and return to Earth. 

INSTRUMENTATION: Stardust has the highest successful Earth return velocity. 

Most heating models breakdown for Stardust. Classic Fay-Riddell is not trustworthy 

for Stardust. 

NOTES: The rendezvous with Wild 2 occurred on Jan 2, 2004. The Stardust capsule 

made a successful return to Earth on Jan 15, 2006. Engineering methods and CFD 

disagree for heating rates. Traj. gives at stag. point 762 W/cm2 peak total heat rate 

and 122 W/cm2 peak radiative heat rate. Total heat load is 20978 J/cm2.

— 21 —
MISSION DESCRIPTION: To collect solar wind particles and return to Earth. MISSION: GENESIS

PLANET: EARTH RETURN

LAUNCH: SEP 8, 2001

ENTRY: SEP 8, 2004

Trajectory Geometry Aero/thermal TPS

Entry angle -8° Shape 59.81° Velocity at 

peak heat

9.2 

km/s

Forebody 

material 

designation

Carbon-carbon 

sheet over 

FiberForm®

Type (1) Pilot: DGB

(2) Main: Parafoil

Entry velocity: 

inertial & 

relative

Inertial: 11.0 

km/s

Relative: 10.8 

km/s

Nose 

radius 0.43 m

Peak 

convective 

heating

670 

W/cm2

Forebody 

thickness & 

mass

3.8 cm C-C sheet; 

2.2 cm carbon 

FiberForm®

Deployment 

Method

(1) Mortar

(2) Pilot parachute

Trim L/D 

(specify trim α) 0 Base area 1.78 m2

Peak 

radiative 

heating

30 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

Diameter / 

Area

(1) 2.07 m dia.

(2) 325 m2 area

Control method Spin-stabilized 

aero-ballistic

Vehicle 

mass 210 kg

Integrated 

total 

heatload

16600 

J/cm2

TPS 

integration 

method

Deployment 

Mach (2) ~1.4

Ballistic co-eff. 80 kg/m3 Payload 

mass N/A PH stag. pressure

N/A

Aftbody

material 

designation

SLA-561V

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration ~25.7 g

TPS mass 

fraction, 

inc. insul.

~18%

Peak stag. 

heating 

rate

700 

W/cm2

Aftbody

thickness & 

mass

Parachute 

materials

nylon fabric

Kevlar structural

Parachutes

INSTRUMENTATION: Thermosensitive paint strips. 

NOTES: The capsule crashed violently into the desert after failing to deploy the drag 

devices. Despite this mishap, many of the collectors remained intact and most of the 

mission goals should be accomplished. 

— 22 —
MISSION: HAYABUSA

PLANET: EARTH RETURN

LAUNCH: MAY 9, 2003

ENTRY: JUL 13, 2007

Trajectory Geometry Aero/thermal TPS

Entry angle -12.79°

@199.97 km Shape

45°

spherecone

Velocity at 

peak heat

10.61 

km/s

Forebody material 

designation

Carbonphenolic Type Cross-type

Entry 

velocity: 

inertial & 

relative

Inertial: 12.04 

km/s

Relative: 11.65 

km/s

Nose radius 0.202 m

Peak 

convective 

heating

1050 

W/cm2

Forebody 

thickness & mass 3.0 cm Deployment method Mortar fired

Trim L/D 

(specify

trim α)

0 Base area 0.128 m2

Peak 

radiative 

heating

264 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: yes

Reference 

diameter / 

area

2.8 m2

Control 

method none Vehicle mass 18 kg

Integrated 

total 

heatload

23051 

J/cm2

TPS integration 

method Segmented Deployment Mach 0.8

Ballistic 

co-eff. 133.72 kg/m3 Payload mass 1.04 kg PH stag. pressure

0.63 atm Aftbody material 

designation

Carbonphenolic

Deployment 

dynamic 

Pressure

Peak 

deceleration ~53.36 g

TPS mass 

fraction, inc. 

insul.

43% Peak stag. 

heating rate

1314 

W/cm2

Aftbody thickness 

& mass 2 cm Parachute materials

Canopy: 

polyester

Parachutes

MISSION DESCRIPTION: To collect samples from asteroid 

Itokawa (1998SF36) and return to Earth.

INSTRUMENTATION: A one-axis accelerometer for parachute 

deployment. 

NOTES: The mission's name was changed from MUSES-C. 

— 23 —
MISSION DESCRIPTION: To develop a low-cost, low-mass system for placing an 

exobiology science payload on Mars. Beagle 2 was transported by the European 

Space Agency’s 2003 Mars Express mission.

MISSION: BEAGLE 2

PLANET: MARS

LAUNCH: JUN 2, 2003

ENTRY: DEC 25, 2003

Trajectory Geometry Aero/thermal TPS

Entry angle

Inertial: 

-15.8° Shape

43.75° truncated 

conical backshell 60°

sphere-cone front 

shield

Velocity at 

peak heat

Relative: 

4.70 

km/s

Forebody 

material 

designation

Norcoat-Liege 

(EADS) Type

(1) DGB pilot

(2) Ringsail

main

Entry 

velocity: 

inertial & 

relative

Inertial: 5.63 

km/s

Relative: 5.40 

km/s

Nose 

radius 0.417 m

Peak 

convective 

heating

72.11 

W/cm2

Forebody 

thickness & 

mass

8 mm

3.9 kg

Deployment 

method

(1) Mortar

(2) Pilot 

parachute

Trim L/D 

(specify 

trim α)

Ballistic (2°

average) Base area 0.67 m2

Peak 

radiative 

heating

0.17 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

Area

(1) 3.2 m dia.

(2) 10 m dia.

Control 

method

Ballistic spinstabilized at 

14.2 rpm

Vehicle 

mass Entry: 68.84 kg

Integrated 

total 

heatload

2449 

J/cm2

TPS 

integration 

method

Tiled Deployment 

Mach (1) 1.5

Ballistic 

co-eff.

Peak heat flux:

69.9 kg/m2

Payload 

mass

11.4 kg (science) 

(33.2 kg landed)

PH stag. 

pressure

0.18 

atm

Aftbody

material 

designation

Norcoat-Liege 

(EADS)

Deployment 

dynamic 

pressure

(2) 730-750 

Pa

Peak 

deceleration

TPS mass 

fraction, 

inc. insul.

Forebody: 9.2% 

Back shell: 15.2% 

Peak stag. 

heating 

rate

72.28 

W/cm2

Aftbody

thickness & 

mass

3 to 6 mm 2kg Parachute 

materials Nylon, Kevlar

Parachutes

INSTRUMENTATION: No TPS instrumentation: axial accelerometers only. 

NOTES: Beagle 2 landed on Mars but did not make radio contact and was never 

operative. 

— 24 —
— 25 —

MISSION: MARS EXPLORATION ROVERS ‘SPIRIT’ AND ‘OPPORTUNITY’

PLANET: MARS

LAUNCH: JUN 10, 2003 & JUL 7, 2003

ENTRY: JAN 3, 2004 & JAN 24, 2004

MISSION DESCRIPTION: To place two rovers (A and B) on Mars to conduct

remote geological investigations including search for past water activity.

NOTES: MER A and MER B are two separate missions, each carrying a rover to 

Mars. Data here are for MER B, the most severe entry environment. This mission 

uses an entry aeroshell like that of Pathfinder, however the enclosed rovers are 

larger than Sojourner and are self-contained. There are 3 TIRS (Transverse 

Impulse Rocket System) covers made of SIRCA spaced around the backshell.

Trajectory Geometry Aero/thermal TPS

Entry angle -10.75° @ 

126.799 km Shape

70° spherecone

Velocity at 

peak heat

~4.96 

km/sec

Forebody 

material 

designation

SLA-561V (SLA561S for backshell) Type Single DGB

Entry 

velocity: 

inertial & 

relative

Inertial: 5.78 

km/sec

Relative: 

5.39 km/sec 

Nose 

radius 0.66 m

Peak 

convective 

heating

~37.72 

W/cm2

Forebody 

thickness & 

mass

1.57 cm Deployment 

method Mortar

Trim L/D 

(specify

trim α)

0 Base 

area 5.499 m2

Peak 

radiative 

heating

~0 Ablating

Ejected

Ablated: yes

Ejected: yes

Reference 

diameter / 

area

14.1 m dia.

Control 

method

Ballistic, 

spinstabilized

Vehicle 

mass 832.2 kg

Integrated 

total 

heatload

~3064.

45 

J/cm2

TPS 

integration 

method

Honeycomb bonded 

to substructure; 

cells filled with 

ablative compound

Deployment 

Mach 1.77

Ballistic 

co-eff.

~88.96 

kg/m2

Payload 

mass

~525 kg

(landed mass, 

rover+lander)

PH stag. 

pressure

0.08 

atm

Aftbody

material 

designation

SLA-561S for 

backshell,

backshell interface 

plate (BIP): SIRCA

Deployment 

dynamic 

Pressure

765±77 Pa

Peak 

deceleration ~4.78 g

TPS 

mass 

fraction, 

inc. 

insul.

Forebody: 

(TPS only) 

3.6% 

Back shell: 

(TPS only) 2% 

Peak stag. 

heating 

rate

~37.72 

W/cm2 Aftbody

thickness & 

mass

0.5 cm (CDR)

10.4 kg (CDR CBE)

Parachute 

materials

Disk: nylon fabric

Band: polyester fabric

Kevlar structural

Parachutes
MISSION DESCRIPTION: First NASA Scout Mission. Goals: (1) to study 

the history of water in the Martian artic and (2) search for evidence of 

habitable zone and assess the biological potential of the ice-soil boundary.

— 26 —

MISSION: PHOENIX

PLANET: MARS

LAUNCH: AUG 4, 2011

ENTRY: MAY 5, 2008

INSTRUMENTATION: 24 thermocouples, MEADS (Mars Entry 

Atmospheric Data System) for temperature and isotherm tracking. 

Trajectory Geometry Aero/thermal TPS

Entry angle [-13.°] Shape

2.65 m diameter, 

70 deg. sphere 

cone

Velocity at 

peak heat

[4.743 

km/s]

Forebody 

Material 

designation

SLA-561V Type Viking-type DGB

Entry 

velocity: 

Inertial & 

Relative

Inertial: 5.6 

km/s

Nose 

radius Rn=0.66 m

Peak 

convective 

heating

[4.8 

W/cm2]

Forebody 

Thickness 

& mass

Deployment 

Method Mortar

Trim L/D 

(specify 

trim α)

0 Base area 0.0995 m2

Peak 

radiative 

heating

Ablating

Ejected

Reference 

Diameter / 

Area

11.8 m dia.

Control 

method

Non-spinning 

passive

Vehicle 

mass 602 kg

Integrated 

total 

heatload

[2301 

J/cm2]

TPS 

Integration 

method

Deployment 

Mach 1.74

Ballistic 

co-eff. 65 kg/m2 Payload mass 59 kg PH stag. pressure

Aftbody

material 

designation

SLA-561S on 

cone, SLA-561V 

on parachute cover 

Deployment 

dynamic 

pressure

489.3 Pa

Peak 

deceleration [9.3 Earth G’s]

TPS mass 

fraction, 

inc. insul.

Peak stag. 

heating 

rate

Aftbody

thickness & 

mass

Parachute 

materials

Nylon canopy, 

Kevlar structural 

members

Parachutes

NOTES: Entry vehicle originally designed for cancelled Mars Surveyor 

2001.
MISSION DESCRIPTION: To determine if Mars was ever able to support microbial 

life.

— 27 —

MISSION: MARS SCIENCE LABORATORY (MSL)

PLANET: MARS

LAUNCH: NOV 26, 2011

ENTRY: AUG 6, 2012

INSTRUMENTATION: 24 thermocouples, MEADS (Mars Entry Atmospheric Data 

System) for temperature and isotherm tracking. 

Trajectory Geometry Aero/thermal TPS

Entry angle -16° Shape

70° spherecone

Velocity at 

peak heat

5.33 

km/s

Forebody 

Material 

designation

PICA Type DGB

Entry 

velocity: 

Inertial & 

Relative

Inertial: 6.08 km/s

Relative: 5.8 km/s

Nose 

radius 1.125 m

Peak 

convective 

heating

[197 

W/cm2]

Forebody 

Thickness & 

mass

[3.18 cm] Deployment 

Method Mortar fired

Trim L/D 

(specify 

trim α)

0.24 Base area 15.90 m

Peak 

radiative 

heating

Ablating

Ejected

Ablated: yes

Ejected: yes

Reference 

Diameter / 

Area

21.35 m

Control 

method

RCS: 8 thrusters

ballast ejection

Vehicle 

mass 3257 kg

Integrated 

total 

heatload

[5477 

J/cm2]

TPS 

Integration 

method

PICA: tiled, 

direct 

bonded to 

substructure 

Deployment 

Mach 1.7

Ballistic 

co-eff. [146 kg/m2] Payload mass 75 kg PH stag. pressure

Aftbody

material 

designation

SLA-561V 

Deployment 

dynamic 

pressure

493.6 Pa

Peak 

deceleration ~9.7 g

TPS mass 

fraction, 

inc. insul.

Peak stag. 

heating rate

Aftbody

thickness & 

mass

[1.27 cm] Parachute 

materials

Canopy: nylon, 

Suspension lines: 

Technora T221 

and Kevlar 29

Parachutes

NOTES: 
MISSION: INFLATABLE REENTRY VEHICLE EXPERIMENT-3 (IRVE-3)

PLANET: EARTH (SUBORBITAL)

LAUNCH: JUL 23, 2012

ENTRY: JUL 23, 2012

MISSION DESCRIPTION: To advance the hypersonic inflatable 

aerodynamic decelerator (HIAD) technology, reach higher than 12 

W/cm2 cold-wall heat flux on the flexible TPS, and generate 

aerodynamic lift, with the potential for guided entry applications. 

INSTRUMENTATION:

NOTES: IRVE-3 is NASA's second successful flight test of Hypersonic 

Inflatable Aerodynamic Decelerator (HIAD). 

Trajectory Geometry Aero/thermal TPS

Entry angle -74° Shape

60° half angle 

sphere-cone

Velocity at 

peak heat Mach 7.3 

Forebody 

Material 

designation

Nextel

BF-20 Type N/A

Entry velocity: 

inertial & 

relative

Relative: 2.7 km/s Nose 

radius 0.191 m

Peak 

convective 

heating

14 W/cm2

(cold wall)

Forebody 

Thickness & 

mass

0.6 cm Deployment 

method N/A

Trim L/D 

(specify trim α) 0.2 (max.) at 16° Base area 7.07 m2 Peak 

radiative 

heating

0 Ablating

Ejected Flexible

Reference 

diameter / 

area

N/A

Control method

Cold argon gas pre-entry 

alignment, roll control 

during entry. Movable 

radial mass offset entry. 

Vehicle 

mass 281 kg 

Integrated 

total 

heatload

193 J/cm2

(cold wall) 

TPS 

Integration 

method

Sewn 

fabric

Deployment 

Mach N/A

Ballistic co-eff. 58 kg/m2 Payload 

mass N/A PH stag. pressure 1 psia

Aftbody

material 

designation

N/A

Deployment 

dynamic 

pressure

N/A

Peak 

deceleration 20.2 g

TPS mass 

fraction, 

inc. insul.

0.20

Peak stag. 

heating 

rate

Aftbody

thickness & 

mass

N/A Parachute 

materials N/A

Parachutes

— 28 —
MISSION: HAYABUSA 2

PLANET: EARTH RETURN

LAUNCH: DEC 3, 2014

ENTRY:DEC 5, 2020

MISSION DESCRIPTION: To collect samples from asteroid RYUGU 

(1999JU3) and return to Earth. 

INSTRUMENTATION: A one-axis accelerometer of parachute 

deployment REMM(Reentry Flight Measurement Module) which measures 

and records the heatshield temperature, the acceleration and the attitude 

motion of SRC during the reentry. 

NOTES:

Trajectory Geometry Aero/thermal TPS

Entry angle -12.06° Shape

45° spherecone

Velocity at 

peak heat

~10.5 

km/s

Forebody 

material 

designation

Carbonphenolic TypeCrossType

Entry velocity: 

inertial & 

relative

Inertial: 11.84 km/s

Relative: 11.58 km/s

Base radius

Nose radius 0.202 m

Peak 

convective 

heating

~1110 

W/cm2

Forebody 

thickness & 

mass

Deployment 

method Mortar fired

Trim L/D 

(specify trim α) 0 Base area 0.128 m2

Peak 

radiative 

heating

~180 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

2.8 m2

Control 

method None Vehicle mass 16.03 kg

Integrated 

total 

heatload

~27000 

J/cm2

TPS 

integration 

method

Segmented Deployment 

Mach Subsonic

Ballistic co-eff. 113.5 kg/m3 Payload 

mass 1.13 kg PH stag. pressure

~0.44 

atm

Aftbody

material 

designation

Carbonphenolic

Deployment 

dynamic 

pressure

< 2kPa

Peak 

deceleration 46.3 g

TPS mass 

fraction, 

inc. insul.

Peak stag. 

heating 

rate

Aftbody

thickness & 

mass

2 cm Parachute 

materials

Canopy: 

polyester

Parachutes

— 29 —
— 30 —

MISSION: EXPLORATION FLIGHT TEST-1

PLANET: EARTH RETURN

LAUNCH: DEC 5, 2014

ENTRY: DEC 5, 2014

MISSION DESCRIPTION: To flight test Orion through two orbits of Earth 

with observation of radiation stresses and orbit debris environment.

INSTRUMENTATION: 2 broadband radiometers. 

NOTES: Aerothermal values are margined. 

Trajectory Geometry Aero/thermal TPS

Entry angle -12.6° Shape Spherical Velocity at 

peak heat

~7.53 

km/s

Forebody 

material 

designation

Avcoat Type

(1) Drogue = VPCRibbon 2x

(2) Pilot = VPCRibbon 3x

(3) Main = Ringsail 3x

Entry 

velocity: 

inertial & 

relative

Inertial: 8.5 

km/s 

Relative: 8.9 

km/s

Nose 

radius 6.035 m

Peak 

convective 

heating

Flown:170 W/cm2 

[Design: 

400W/cm2]

Forebody 

thickness & 

mass

Var. thick 2.5-

4cm ~454 kg

Deployment 

method

(1) Mortar drogue

(2) Mortar pilot

(3) Pilot main

Trim L/D 

(specify 

trim α)

0.25 Base area 19.8 m2

Peak 

radiative 

heating

Flown: 15 W/cm2 

[Design: 50 

W/cm2] 

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

(1) 7.0 m dia.

(2) 3.0 m dia.

(3) 35.4 m dia.

Control 

method

RCS 

Thrusters

Vehicle 

mass 9611 kg

Integrated 

total 

heatload

~250 

MJ/m2

TPS 

integration 

method

Honeycomb 

bonded to 

substructure; 

cells filled with 

ablative 

compound

Deployment 

Mach

(1) 0.42

(2) 0.17

Ballistic 

co-eff.

360 

kg/m2 (CD = 

1.4 at EI+50 

sec.)

Payload 

mass Minimal PH stag. pressure

~0.56 

atm

Aftbody

material 

designation

RCG coated 

AETB-8 

Deployment 

dynamic 

pressure

(1) ~5746 Pa

(2) ~1915 Pa

Peak 

deceleration ~3.2 g

TPS mass 

fraction, 

inc. insul.

Fore: 

4.7% 

Aft: 1.2%

Peak stag. 

heating 

rate

Aftbody

thickness & 

mass

Var. thick 2.5-

4.2 cm ~120 kg

Parachute 

materials Nylon and Kevlar

Parachutes
MISSION: EXOMARS 2016

PLANET: MARS

LAUNCH: MAR 14, 2016

ENTRY: OCT 19, 2016 INSTRUMENTATION: Front Shield: 4 pressure sensors; 7 thermal plugs (each with 3 

thermocouples and 1 thermistor). Back Shell: 3 thermal plugs (each with 2 thermocouples 

and 1 thermistor); COMARS+ (3 combined sensor heads + one broadband IR radiometer). 

MISSION DESCRIPTION: To demonstrate entry, descent, and landing of a payload on the 

surface of Mars.

NOTES: ExoMars 2016 comprised a trace gas orbiter and the Schiaparelli EDL 

demonstrator module. An anomaly in the measurement and navigation sequence resulted 

in Schiaparelli crash-landing on the surface of Mars, but essential telemetry was received 

during entry and descent.

Trajectory Geometry Aero/thermal TPS

Entry angle -12.48°

@120km Shape

47° truncated conical 

backshell 

70° sphere-cone front 

shield

Velocity at 

peak heat

~5 

km/s

Forebody 

material 

designation

Norcoat-Liège 

over Al 

honeycomb 

with CFRP 

skins 

Type DGB

Entry 

velocity: 

inertial & 

relative

Inertial: 

6.03 km/s 

Relative: 

5.79 km/s

Nose radius 0.6 m

Peak 

convective 

heating

~50 - 70 

W/cm2

Forebody 

thickness & 

mass

12.2 mm 

(nose) to 15.4 

mm (trailing 

edge) 37 kg 

Deployment 

method

Mortar 

fired

Trim L/D 

(specify 

trim α)

Base area

4.52 m2 (2.4 m dia. 

heatshield)

Peak 

radiative 

heating

~9 

W/cm2

Ablating

Ejected

Ablated: yes

Ejected: yes

Reference 

diameter / 

area

12 m 

dia.

Control 

method

Ballistic, 

spinstabilized at 

2.617 rpm

Vehicle 

mass 576.3 kg Integrated total heatload

TPS integration 

method Tiled Deployment Mach 2.08

Ballistic 

co-eff. 75.2 kg/m2 Payload mass 15.2 kg PH stag. pressure

~0.11 

Atm

Aftbody material 

designation

same as 

forebody

Deployment 

dynamic 

pressure

779.5

Peak 

deceleration 8.15 g

TPS mass 

fraction, inc. 

insul.

0.11 Peak stag. 

heating rate

Aftbody

thickness & 

mass

7.6 mm 

26.2 kg

Parachute 

materials

Nylon, 

Kevlar

Parachutes

— 31 —
— 32 —

Trajectory Geometry Aero/thermal TPS

Entry angle [-8.2°] Shape

Blunt-nosed 

60° half-angle 

once

Velocity at 

peak heat

11.1 

km/s

Forebody 

material 

designation

PICA-15 Type

(1) Pilot: DGB

(2) Main: 

Triconical

Entry 

velocity: 

inertial & 

relative

Relative: 

[12.2 km/s] Nose radius 0.23 m

Peak 

convective 

heating

[1070 

W/cm2]

Forebody 

thickness & 

mass

5.82 cm Deployment 

method

(1) Mortar

(2) Pilot parachute

Trim L/D 

(specify 

trim α)

0 Base area 0.52 m2 

Peak 

radiative 

heating

[130 

W/cm2]

Ablating

Ejected

Ablated: yes 

Ejected: no

Reference 

diameter / 

area

(1) 0.83 m dia.

(2) 7.3 m dia.

Control 

method

Ballistic, 

spinstabilized 

at 13 rpm

Vehicle mass ~46 kg

Integrated 

total 

heatload

[36000 

J/cm2]

TPS 

integration 

method

PICA: singlepiece, direct 

bonded to 

substructure 

with HT-424

Deployment 

Mach

(1) [1.3] 

(2) [0.12]

Ballistic 

co-eff. ~71 kg/m2 Payload mass N/A PH stag. pressure

[0.22 

atm]

Aftbody

material 

designation

SLA-561V 

Deployment 

dynamic 

pressure

(1) ~1350 Pa

(2) ~700 Pa

Peak 

deceleration ~32 g

TPS mass 

fraction, inc. 

insul.

0.22 Peak stag. 

heating rate

[1200 

W/cm2]

Aftbody

thickness & 

mass

[1.397 cm] Parachute 

materials

Canopy: nylon, 

Kevlar suspension 

lines 

Parachutes

MISSION DESCRIPTION: To return and analyze samples from asteroid 

Bennu (1999RQ36) and return to Earth.

INSTRUMENTATION:

NOTES: The OSIRIS-REx capsule outer geometry and TPS are 

build-to-print based on the Stardust capsule.

MISSION: OSIRIS-REX

PLANET: EARTH RETURN

LAUNCH: SEPT 8, 2016

ENTRY: PLANNED FOR

SEPT 24, 2023
— 33 —

Trajectory Geometry Aero/thermal TPS

Entry angle -12° Shape

2.65 m 

diameter, 70°

sphere cone

Velocity at 

peak heat

4835.8 

m/s

Forebody 

material 

designation

SLA-561V Type Viking-type 

DGB

Entry 

velocity: 

inertial & 

relative

Relative: 

5.523 

km/s

Nose 

radius 0.66 m

Peak 

convective 

heating

46 

W/cm2

Forebody 

thickness & 

mass

1.93 cm & 35.74 kg Deployment 

method Mortar

Trim L/D 

(specify 

trim α)

0 Base area 0.0995 m2 

Peak 

radiative 

heating

5 W/cm2 

stag. pt.

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

11.8 m dia.

Control 

method

Nonspinning 

passive

Vehicle 

mass

605.65 kg 

(wet)

Integrated 

total 

heatload

2514 

J/cm2

TPS 

integration 

method

Direct application of flexcore

then filler

Deployment 

Mach 1.53

Ballistic 

co-eff. 65 kg/m2 Payload mass ~50 kg PH stag. pressure

6.38 

kPa

Aftbody

material 

designation

SLA-561S lower aftbody, SLA220 parachute cone, SLA-561V 

on parachute lid

Deployment 

dynamic 

pressure

546.0 Pa

Peak 

deceleration 8.13 g

TPS mass 

fraction, 

inc. insul.

0.85

Peak stag. 

heating 

rate

~7 

W/cm2

Aftbody

thickness & 

mass

0.61 cm lower aftbody 12.46 

kg, 1.22 cm parachute cone 

3.07 kg, 0.89 cm parachute lid 

0.12 kg

Parachute 

materials

Nylon canopy, 

Kevlar structural 

members

Parachutes

MISSION DESCRIPTION: Robotic lander to study the deep interior 

of Mars.

INSTRUMENTATION: SEIS seismometer to measure Mars quakes, 

HP3 Heat Flow and Physical Properties Package to measure heat 

flowing out of deep interior, RISE radio science instrument to measure 

wobble of Mars north pole.

NOTES: Discovery program mission; InSight = Interior Exploration using Seismic Investigations, 

Geodesy and Heat Transport; EDL system reused design of Phoenix mission augmented for landing 

during dust storm season; dust on solar panels led to mission termination on December 21, 2022.

MISSION: MARS InSight

PLANET: MAR

LAUNCH: MAY 5, 2018

ENTRY: SEPT 26, 2018
— 34 —

Trajectory Geometry Aero/thermal TPS

Entry angle

Relative: 

-16.18° Shape

70° spherecone

Velocity at 

peak heat 

(PH)

4880 

m/s

Forebody 

material 

designation

PICA Type DGB

Entry 

velocity: 

inertial & 

relative

Relative: 

5.33 km/s 

@ 128.2 

km

Nose radius 1 m

Peak 

convective 

heating

68.4 

W/cm2

Forebody 

thickness & 

mass

[3.18 cm] 

[212 kg]

Deployment 

method Mortar fired

Trim L/D 

(specify 

trim α)

0.248 at 

 -16.71° Base area 15.9 m2 

Peak 

radiative 

heating

~0 

W/cm2

Ablating

Ejected

Ablated: yes 

Ejected: yes

Reference 

diameter / 

area

21.5 m

Control 

method

RCS: 8 

thrusters 

ballast 

ejection

Vehicle mass 3317.3 kg 

Integrated 

total 

heatload

1672.5 

J/cm2

TPS 

integration 

method

PICA: tiled, direct 

bonded to 

substructure

Deployment 

Mach 1.81

Ballistic 

co-eff.

143.2 

kg/m2 

Payload 

mass 1022.8 kg PH stag. pressure 24 kPa

Aftbody

material 

designation

SLA-561V 

Deployment 

dynamic 

pressure

519.6 Pa

Peak 

deceleration 10.73 g

TPS mass 

fraction, inc. 

insul.

Fore: [6.4%] 

Aft: [1.9%]

Peak stag. 

heating rate

68.4 

W/cm2

Aftbody

thickness & 

mass

[1.27 cm] 

[64 kg]

Parachute 

materials

Canopy: nylon, 

suspension lines: 

Technora T221 

and Kevlar 29

Parachutes

MISSION DESCRIPTION: Seek signs of ancient life and collect samples of rock and regolith 

(broken rock from soil) for possible return to Earth.

INSTRUMENTATION: MEADS (Mars Entry Atmospheric Data System): 1 hypersonic 

heatshield pressure transducer, 6 supersonic heatshield pressure transducers, 1 low pressure 

transducer on backshell. MISP (MEDLI2 Instrumented Sensor Plugs): 11 PICA heatshield 

plugs with 1-3 embedded TCs, 6 SLA-561V plugs on backshell with 1-2 embedded TCs, and 

3 direct heat flux sensing elements on backshell (2 total heat flux gauges and 1 radiometer). 

NOTES:

MISSION: MARS2020

PLANET: MARS

LAUNCH: JUL 30, 2020

ENTRY: FEB 18, 2021
— 35 —

Trajectory Geometry Aero/thermal TPS

Entry angle [-2.3° ] Shape

70° spherecone 

stacked 

torus

Velocity at 

peak heat 

(PH)

[7.02 km/s]

Forebody 

material 

designation

FTPS: Silicon carbide 

outer fabric, carbon 

felt/pyrogel insulation, 

TZL gas barrier

Type

(1) Pilot: 16 .6 

ft Cruciform

(2) Main: 94 ft 

Ring Sail

Entry 

velocity: 

inertial & 

relative

Relative: 

[8.0 

km/s] 

Nose 

radius 1.715 m

Peak 

convective 

heating

30 W/cm2

Forebody 

thickness & 

mass

0.5" / 114 kg Deployment 

method

(1) Mortar

(2) Pilot

Trim L/D 

(specify 

trim α)

0 Base area 28.27 m2

Peak 

radiative 

heating

2 W/cm2 Ablating

Ejected

Ablative: No

Ejected: No

Reference 

diameter / 

area

(1) 16.6 ft

(2) 94 ft

Control 

method

Ballistic, 

spinstabilized 

at 3 rpm

Vehicle 

mass 1101 kg

Integrated 

total 

heatload

2.7 kJ/cm2

TPS 

integration 

method

Sewn Fabric Deployment 

Mach

(1) 0.1

(2) 0.1

Ballistic 

co-eff. 22 kg/m2 Payload mass N/A PH stag. pressure ~0.4 psia

Aftbody

material 

designation

N/A

Deployment 

dynamic 

pressure

[415 Pa]

Peak 

deceleration ~9 g

TPS mass 

fraction, 

inc. insul.

0.17 

(aeroshell)

Peak stag. 

heating 

rate

[0.66 W/cm2 (cold 

wall)]

Aftbody

thickness & 

mass

N/A Parachute 

materials Nylon, Kevlar

Parachutes

MISSION DESCRIPTION: Low-Earth Orbit Flight Test of an Inflatable 

Decelerator (LOFTID). Technology demonstration mission to advance 

hypersonic inflatable aerodynamic decelerator (HIAD) technology by testing 

the Gen-2 inflatable structure and flexible thermal protection system (FTPS).

INSTRUMENTATION: 58 FTPS thermocouples, 24 inflatable structure thermocouples, 3 fiber 

optic sensing systems (FOSS), 4 total heat flux sensors, 1 radiometer, 5 nose pressure 

transducers, 12 loadcell pins, 7 torus pressure transducers, 6 inflation system pressure 

transducers, 6 inflation system pressure transducers, 3 inflation system RTDs, 1 inflation system 

flow meter, 2 hot film anemometers, 6 visual cameras, 12 infrared cameras, 1 up-look camera.

NOTES:

MISSION: LOFTID

PLANET: Earth

LAUNCH: Nov 10, 2022

ENTRY: Nov 10, 2022
— 36 —

Trajectory Geometry Aero/thermal TPS

Entry angle -5.66° Shape Capsule

Velocity at 

peak heat 

(PH)

~34,200 ft/2

Forebody 

material 

designation

Avcoat Type Ringsail

Entry 

velocity: 

inertial & 

relative

Inertial: 

36062.66 ft/s

Relative: 

35974.97 ft/s 

Nose 

radius 6.04 m

Peak 

convective 

heating

~175 BTU/ft^2/s

Forebody 

thickness & 

mass

Varied 

thickness 

~223 lbm

Deployment 

method Mortar

Trim L/D 

(specify 

trim α)

[0.27 at 162°] Base area 3.33 m2 

Peak 

radiative 

heating

~180 BTU/ft^2/s Ablating

Ejected Ablator

Reference 

diameter / 

area

FBC chutes (3) - 7 ft

Drogues (2) - 23 ft, 

400 ft^2

Pilot (2) - 11 ft, 95 ft^2

Mains (4) - 116 ft, 

10,500 ft^2

Control 

method 12 RCS Jets Vehicle mass

~20575 

lbm

Integrated 

total 

heatload

~70000 Btu/ft^2 

TPS 

integration 

method

Tiled Deployment 

Mach

~0.42 (FBC chutes)

~0.39 (drogues)

~0.17 (pilots/mains)

Ballistic 

co-eff. ~68.33 lbf/ft2 Payload mass

PH stag. 

pressure ~22,000 lb/ft^2

Aftbody

material 

designation

AETB-8

Deployment 

dynamic 

pressure

~3600 lb (FBC chutes)

~3500 lb (drogues)

~1000 lb (pilots/mains)

Peak 

deceleration ~13.4 g

TPS mass 

fraction, 

inc. insul.

Heatshield: 

~27%

Peak stag. 

heating 

rate

qtot ~20 BTU/ft^2/s, 

qconv ~15 

BTU/ft^2/s, qrad ~6 

BTU/ft^2/s

Aftbody

thickness & 

mass

Varied 

thickness 

1.0-1.2"

Parachute 

materials Kevlar/Nylon

Parachutes

MISSION DESCRIPTION: Artemis I is the first in a series of increasingly 

complex missions that will enable human exploration to the Moon and Mars.

INSTRUMENTATION: Thermocouples (multiple locations and in-depth), 

Pressure Transducers, and Radiometers. 

NOTES: Data is mostly from internal documentation.

MISSION: Artemis I

PLANET: Earth

LAUNCH: Nov 15, 2022

ENTRY: Dec 11, 2022
APPENDIX I List of Space Vehicles and Their Missions Unmanned Planetary Probes

For a complete list of deep space exploration go to www.nasa.gov –> Downloads –> History E-books –> 

Beyond Earth: A Chronicle of Deep Space Exploration.

History e-Books | NASA

— 37 —

Entry angle

The angle between the local horizontal plane (orthogonal to the vector from 

the planet center to the vehicle) and the velocity vector of the vehicle, V, at a 

reference altitude, h. The entry angle can be inertial or relative, depending on 

entry velocity used. Y is negative when V, is below the horizontal plane, as in 

planetary entry.

Entry velocity inertial

The vehicle velocity at reference altitude, h, assuming a non-rotating planet.

Entry velocity relative

The inertial entry velocity amended by the component of the planet’s rotation, 

assuming the atmosphere to be a solid body.

Trim L/D

The L/D where vehicle is statically stable. The vehicle will trim (restore) to trim 

angle of attack if variations occur.

Control method

(a) Ballistic: no control, subject to drag forces only, with passive stability about 

zero lift condition; (b) Controlled Ballistic: active control to maintain zero lift; 

and (c) RCS: a set of small engines called the reaction control system (RCS) 

engines.

Ballistic coefficient

The ratio of the product of drag coefficient (Cd) and projected reference area 

(A) to mass (m), giving m/CdA.

APPENDIX II Definitions

Peak deceleration

Maximum deceleration force on the vehicle during entry, stated in Earth gravity, g.

Shape

All vehicles are spherically blunted cones, or spherical, or conical.

Nose radius

The radius of the spherical nose, or the capsule radius.

Base area

The total apparent or drag area projected along the centerline.

Vehicle entry mass

The total vehicle mass of the vehicle at entry, including TPS and payload. Generally, 

the vehicle mass at entry is the same as take-off mass minus any fuel used for 

maneuvering. However, the mass can change after leaving the orbiter but before 

entry. An example is the small probe of Pioneer Venus where the spin yo-yo was 

jettisoned before entry. The mass can also change during entry if the heat shield 

material ablates. An (extreme) example was the Galileo probe that lost about 26% of 

its entry mass to ablation.

Payload mass

The proportion of payload mass (scientific instruments and may include transmitters, 

batteries etc.) to vehicle mass. For missions with landed payloads and eject-able 

heatshields/backshells, this is total mass of the vehicle carried within the entry 

system. 
TPS mass fraction

The proportion of TPS mass to vehicle mass at entry. Insulator may or may 

not be included.

Velocity at peak heat (PH)

The velocity when the vehicle reaches the maximum convective heat flux at 

the stagnation point. 

Peak convective heating

The maximum convective heat flux at the leading edge stagnation point. 

Depends on trajectory. The stagnation point is a point in the flowfield where 

the local velocity of the fluid is zero. Location depends on angle of attack and 

deviation behind the shock. The “peak convective heating” (peak convective 

heat flux) is often at the entry vehicle’s leading edge stagnation point, but not 

always.

Peak radiative heating

The maximum radiative heat flux at the leading edge stagnation point. 

Depends on trajectory. The stagnation point is a point in the flowfield where 

the local velocity of the fluid is zero. Location depends on angle of attack and 

deviation behind the shock. The “peak radiative heating” (peak radiative heat 

flux) is often at the entry vehicle’s leading edge stagnation point, but not 

always.

Integrated total heat load

The convective heat flux integrated over flight time. The highest heat load is 

usually, but not always, at the stagnation point. The integrated heat load will 

vary over the vehicle surface.

Peak heat (PH) stagnation pressure

The pressure at the time of maximum convective heat flux. This is not the 

peak pressure, which occurs later in the trajectory.

Peak stag. heating rate

The maximum combined convective and radiative heating rate during the 

trajectory. Peak convective stagnation point heating seldom occurs at the 

same time as peak radiative stagnation point heating, therefore, not the 

simple sum of those values unless one is zero. Peak radiative heating rate (if 

nonzero) always occurs before peak convective heating rate.

Forebody material designation

This can be a material trade name, defined by the manufacturer (e.g., SLA561V) or a generic designator applied to a class of materials (e.g., carbon 

phenolic). It provides little useful information about the material

other than a broad description of its constituents.

Forebody thickness & mass

This is “as manufactured” thickness of the material, usually specified at the 

stagnation point. Useful for a TPS of uniform thickness; less useful for a 

“tailored” TPS. The “as manufactured” thickness includes the “nominal design 

thickness” to which additional thickness is added (margin) to accommodate 

uncertainties in the entry environment and/or material performance.

Ablating / Ejected

Information on whether the forebody TPS material experienced or was 

designed for ablating. Ejected refers to whether the vehicle was designed to 

eject the forebody TPS. 

TPS integration method

Details on TPS integration method with the supporting substructure. Examples 

are single piece, direct bonded to substructure; cells filled with ablative 

compound.

Aftbody material designation

This can be a material trade name, defined by the manufacturer (e.g./ SLA561V) or a generic designator applied to a class of materials (e.g., carbon 

phenolic). It provides little useful information about the material other than a 

broad description of its constituents.

Aftbody thickness & mass

This is “as manufactured” or “as flown” thickness of the material. Descriptions 

vary by mission, since a common reference point for a thickness specification, 

such as the stagnation point, doesn’t exist.

Parachute: Type

General parachute design type descriptor for decent or recovery parachutes. 

Examples include Cross Type, Ringsail, or Disk-Gap-Band (DGB).

Deployment method

Method for releasing parachute system during descent/ entry. An example is 

mortar fired.

— 38 —
APPENDIX III Unit Conversion Factors

1 cm = 0.3937 in

1 m = 3.28084 ft

1 m2 = 10.764 ft2

1 km/s = 3280.8 ft/s

1 g =. 9.81 m/s2

1 kg = 2.20462 lb

1 kg/m2 = 0.2048 lb/ft2

APPENDIX IV References Mission: FIRE II

1. Cauchon, D.V.: Radiative Heating Results from the Fire II Flight Experiment at 

a Re-entry Velocity of 11.4 km/s. NASA-TM-X-1402, Jul 1967. 

2. Cornette, E.S.: Forebody Temperatures and Calorimeter Heating Rates 

Measured during Project Fire II Re-entry at 11.35 km/s. NASA-TM-X-1305, Nov 

1996. 

3. Slocumb, T.H.: Project Fire Flight II Afterbody Temperatures and Pressures at 

11.35 km/s. NASA-TM-X-1319, Dec 1966. 

4. Wright, M.; Loomis, M.; and Papadopoulos, P.: Aerothermal Analysis of the 

Project Fire II Afterbody Flow. AIAA-2011-3065, 35th Thermophysics Conference, 

Anaheim Ca, Jun 2011

Collected by: Trajectory Data Verified by: Traj Ver 17.06, Gary A Allen, Jr., 22 

Feb 2019

Mission: Apollo AS-201

1. Lee, D.B.; Bertin, J.J.; and Goodrich, W.D.: Heat Transfer and Pressure 

Measurements Obtained During Apollo Orbital Entries. NASA-TN-D-6028, Oct 

1970.

2.Erb, R.B.; Greeshields, D.H.; Lee, D.B.; and Weston, K.C.: 

Aerothermodynamics-Apollo Experience. Proceedings of the 1967 Heat Transfer 

and Fluid Mechanics Institute, San Diego CA, Jun 1967, edited by P.A Libby; D.B. 

Olfe; and C.W. Van Atta, Stanford University Press, 1967.

3. Lee D.B.; Apollo Experience Report: Aerothermodynamics Evaluation. NASATN-6843, Jun 1972.

4. North American Aviation Inc. Space and Information Systems Division, 

Aerodynamic Data Manuel for Project Apollo, NAS 9-150, Accession No. 07499-

65, Feb 1966.

5. NASA.TN-D-7437 Apollo Earth Landing System.

Mission: Apollo AS-202 (Apollo-3)

1. Lee, D.B.; Bertin, J.J.; and Goodrich, W.D.: Heat Transfer and Pressure 

Measurements Obtained During Apollo Orbital Entries. NASA-TN-D-6028, Oct 

1970. 

2. Hillje, E.R.: Entry Flight Aerodynamics from Apollo Mission AS-202. NASATN0#04185, Dec 1967. 

3. Erb, R.B.; Greenshields, D.H.; Lee, D.B.; and Weston, K.C.: 

Aerothermodynamics-Apollo Experience. Proceedings of the 1967 Heat Transfer 

and Fluid Mechanics Institute, San Diego CA, Jun 1967, edited by P.A. Libby; 

D.B. Olfe; and C.W. Van Atta, Stanford University Press, 1967. 

4. Crowder, R.S.; and Moote, J.D.: Apollo Entry Aerodynamics. Journal of 

Spacecraft and Rockets, Vol 6, No 3, Mar 1969.

5. NASA.TN-D-7437 Apollo Earth Landing System.

1 kg/m3 = 0.06243 lb/ft3

1 Joule = 0.9478x10–3 BTU

1 Watt = 1 J/s = 0.9478x10–3 BTU/s

1 J/cm2 = 0.88055 BTU/ft2

1 W/cm2 = 0.88055 BTU/s*ft2

1 atm = 1.01325x105 Pa

— 39 —

Reference diameter / area

Number to generate reference area, or calculated from reference area of the 

decelerator, used as a common size reference for interpreting decelerator 

aerodynamic properties. 

Deployment mach

Local Mach number at the time of parachute deployment.

Deployment dynamic pressure

The measured/ calculated dynamic pressure at the forebody at the time of 

deployment (similar to Mach number).

Parachute materials

Description of parachute materials used for construction, including canopy 

and suspension lines when possible. Examples include nylon, polyester, and 

Kevlar. 
Mission: Apollo AS-501 (Apollo-4)

1. Hillje, E.R.; and Savage, R.: Status of Aerodynamic Characteristics of the 

Apollo Entry Configuration. AIAA-1968-1143, Entry Vehicle Systems and 

Technology Meeting, Williamsburg VA, Dec 1968. 

2. Hillje, E.R.: Entry Aerodynamics at Lunar Return Conditions Obtained from the 

Flight of Apollo 4 (AS-501). NASA-TN-D-5399, Oct 1969. 

3. Reid, R.C., Jr; Rochelle, W.C.; and Milhoan, J.D.: Radiative Heating to the 

Apollo Command Module: Engineering Prediction and Flight Measurement. 

NASA-TM-X-58091, Apr 1972.

4. NASA.TN-D-7437 Apollo Earth Landing System.

Mission: Apollo AS-502 (Apollo-6)

1. Hillje, E.R.; and Savage, R.: Status of Aerodynamic Characteristics of the 

Apollo entry Configuration. AIAA-1968-1143, Entry Vehicle Systems and 

Technology Meeting, Williamsburg VA, Dec 1968. 

2. Lee, D.B.; and Goodrich, W.D.: The Aerothermodynamic Environment of the 

Apollo Command Module During Superorbital Entry. NASA-TN-D-6792, Apr 1972. 

3. Strouhal, G.; Curry, M.; and Janney, J.M.: Thermal Protection System 

Performance of the Apollo Command Module. AIAA/ASME 7th Structures and 

Materials Conference, Cocoa Beach FL, Apr 1966. 

4. Bartlett, E.P.; Nicolet, W.E.; Abbett, M.J.; and Moyer C.B.: Improved HeatShield Design Procedures for Manned Entry Systems: Part 2. Application to 

Apollo: Final Report. NASA-CR-108689, Jun 1970.

5. NASA.TN-D-7437 Apollo Earth Landing System.

Mission: Reentry F

1. Write, R, L.; and Zoby, E.V.: Flight Measurements of Boundary-Layer 

Transition on a 5 Degree Half-Angel Cone at a Free-Stream Mach Number of 20 

(Reentry F). NASA-TM-X-2253, May 1971.

Mission: PAET

1. Vojvodich, N.S.: PAET Entry Heat Protection Experiment. Journal of 

Spacecraft and Rockets, Vol 10, No 3, Mar 1973, pp 181-189. 

2. Seiff, A.; Reese, D.E.; Sommer, S.C.; Kirk, D.B.; Whiting, E.E.; and Niemann, 

H.B.: PAET, An Entry Probe Experiment in Earth's Atmosphere. ICARUS, Vol 18, 

Apr 1973, pp 525-563

Mission: Viking Lander 1

1. Inogoldby, R.N.; Michael, F.C.; Flaherty, T.M.; Doty, M.G.; Preston, B.; 

Villyard, K.W.; and Steele, R.D.: Entry Data Analysis of Viking Landers 1 and 2: 

Final Report. Martin-Marietta Co., TN-3770218, NASA-CR-159388, Nov 1976. 

2. Kirk, D.B.; Intrieri, P.F.; and Seiff, A.: Aerodynamics Behaviour of the Viking 

Entry Vehicle: Ground Test and Flight Results. Journal of Spacecraft and 

Rockets, Vol 156, No 4, Jul-Aug 1978. 

3. Viking '75 Project, Aerodynamics Data Book. NAS1-9000, Martin-Marietta, 

Revision E, 1974. 

4. Cooley, C.G: Viking 75 Project: Viking Lander System, Primary Mission 

Performance Report. NASA-CR-145148, Apr 1977. 4. Edquist, K.T.; Wright, M.J.; 

and Allen Jr., G.A.: Viking Afterbody Heating Computations and Comparisons to 

Flight Data. AIAA-2006-386, 44th Aerospace Sciences Meeting, Reno, NV, Jan 

2008.

5. NASA-CR-112288 Balloon Launched Viking Decelerator Test Program.

6. AIAA PAPER 73-458 The Viking mortar – Design, development, and flight 

qualification.

Mission: Viking Lander 2

1. Inogoldby, R.N.; Michael, F.C.; Flaherty, T.M.; Doty, M.G.; Preston, B.; 

Villyard, K.W.; and Steele, R.D.: Entry Data Analysis of Viking Landers 1 and 2: 

Final Report. Martin-Marietta Co., TN-3770218, NASA-CR-159388, Nov 1976. 

2. Kirk, D.B.; Intrieri, P.F.; and Seiff, A.: Aerodynamics Behaviour of the Viking 

Entry Vehicle: Ground Test and Flight Results. Journal of Spacecraft and 

Rockets, Vol 156, No 4, Jul-Aug 1978. 

3. Viking '75 Project, Aerodynamics Data Book. NAS1-9000, Martin-Marietta, 

Revision E, 1974. 

4. Cooley, C.G: Viking 75 Project: Viking Lander System, Primary Mission 

Performance Report. NASA-CR-145148, Apr 1977.

5. NASA-CR-112288 Balloon Launched Viking Decelerator Test Program.

6. AIAA PAPER 73-458 The Viking mortar – Design, development, and flight 

qualification.

— 40 —
Pioneer-Venus: Small ’North Probe’

1. Nolte, L.J.; and Sommer, S.C.: Probing a Planetary Atmosphere: PioneerVenus Spacecraft Description. AIAA-1975-1160, Conference on the Exploration 

of the Outer Planets, St. Louis Mo, Sep 1975. 

2. Pioneer-Venus Large and Small Probe Databook, Bendix, NAS2-830Q, Jun 

1976. 

3. Entry vector based upon a letter by Charles F. Hall, 21 Mar 1979.

4.Allen, G.: Trajectory and Heating calculated using TRAJ Code, private 

communication, Mar 2003.

Pioneer-Venus: Small ’Night Probe’

1. Nolte, L.J.; and Sommer, S.C.: Probing a Planetary Atmosphere: PioneerVenus Spacecraft Description. AIAA-1975-1160, Conference on the Exploration 

of the Outer Planets, St. Louis Mo, Sep 1975. 

2. Pioneer-Venus Large and Small Probe Databook, Bendix, NAS2-830Q, Jun 

1976. 

3. Allen, G.: Trajectory and Heating calculated using TRAJ Code, private 

communication, Mar 2003.

Pioneer-Venus: Small 'Day Probe' 

1. Nolte, L.J.; and Sommer, S.C.: Probing a Planetary Atmosphere: PioneerVenus Spacecraft Description. AIAA-1975-1160, Conference on the Exploration 

of the Outer Planets, St. Louis Mo, Sep 1975. 

2. Pioneer-Venus Large and Small Probe Databook, Bendix, NAS2-830Q, Jun 

1976. 

3. Entry vector based upon a letter by Charles F. Hall, 21 Mar 1979.

4. Allen, G.: Trajectory and Heating calculated using TRAJ Code, private 

communication, Mar 2003.

Pioneer-Venus: Large Probe 'Sounder' 

1. Nolte, L.J.; and Sommer, S.C.: Probing a Planetary Atmosphere: PioneerVenus Spacecraft Description. AIAA-1975-1160, Conference on the Exploration 

of the Outer Planets, St. Louis Mo, Sep 1975. 

2. Pioneer-Venus Large and Small Probe Databook, Bendix, NAS2-830Q, Jun 

1976. 

3. Entry vector based upon a letter by Charles F. Hall, 21 Mar 1979.

4. Allen, G.: Trajectory and Heating calculated using TRAJ Code, private 

communication, Mar 2003.

5. IPPW-7-529, Aerodynamic Decelerators for Modern Venus Probes/Landers, 7th

International Planetary Probe Workshop, Barcelona, Spain.

Galileo 

1. Givens, J.; Nolte,L.; and Pochettino, L.: Galileo Atmospheric Entry Probe 

System: Design, Development and Test. AIAA-1983-0098, 21st Aerosciences

Meeting, Jan 1983. 

2. Milos, F.; Chen, Y.K.; Squire, T.; and Brewer, R.: Analysis of Galileo Probe 

Heat Shield Ablation and Temperature Data. AIAA-1997-2480, 32nd 

Thermophysics Conference, Atlanta GA, Jun 1997. 

3. Milos, F.: Galileo Probe Heat Shield Ablation Experiment. AIAA Journal of 

Spacecraft and Rockets, Vol 34, No 6, Nov-Dec 1997.

4. AIAA-97-1510 Galileo Parachute System Performance.

Orex

1. Yamamoto, Y.; and Yoshioka, M.: CFD and FEM Coupling Analysis of OREX 

Aerothermodynamic Flight Data. AIAA-1995-2087, 30th Thermophysics

Conference, San Diego CA, Jun 1995. 

2.Gupta, R.N.; Moss, J.M.; and Price, J.M.: Assessment of Thermochemical 

Nonequilibrium and Slip Effects for Orbital Reentry Experiment (OREX). AIAA1996-1859, 31st Thermophysics Conference, New Orleans LA, Jun 1996.

Pathfinder 

1. Spencer, D.A.; Blanchard, R.C.; Braun, R.D.; Kallemeyn, P.H.; and Thurman, 

S.W.: Mars Pathfinder Entry, Descent, and Landing Reconstruction. Journal of 

Spacecraft and Rockets, Vol 36, No 3, May-Jun 1999. 

2. Milos, F.S.; Chen, Y.K.; Congdon, W.M.; and Thornton, J.M.: Mars Pathfinder 

Entry Temperature Data, Aerothermal Heating, and Heatshield Material 

Response. Journal of Spacecraft and Rockets, Vol 36, No 3, May-June 1999.

3. Michaeltree, R.A.; and Gnoffo, P.A.: Wake Flow About the Mars Pathfinder 

Entry Vehicle. Journal of Spacecraft and Rockets, Vol. 32, No. 5, Sep-Oct 1995. 

Witkowski, A.: Mars Pathfinder Parachute System Performance. AIAA-1999-

1701, 15th Aerodynamic Decelerator Systems Technology Conference, 

Toulouse, France, Jun 1999. 

Mirka

1. Schmitt, G.; Pfeuffer, H.; Kasper, R.; Kleppe, F.; Burkhardt, J.; and Schottle, 

U.M.: The MIRKA Re-entry Mission. IAF-98-V2.07, 49th Intl Astronautical 

Congress, Melbourne, Australia, Sep-Oct 1998. 

2. Schmitt, G.; and Kasper, R.: MIRKA Micro Re-entry Capsule. IAF-94-V.2.532, 

45th Intl Astronautical Congress, Jerusalem, Israel, Oct 1994.

— 41 —
3. Ocidation Behaviour of SiC-based Thermal Protection System Materials using 

newly developed Prode Techniques Journal of Spacecraft and Reockets, pg. 817-

824, Vol. 42, No. 5, Sept.-Okt. 2005. G. Herdrich, M. Auweter-Kurtz, M. Fertig, S. 

Löhle, S. Pidan, T. Laux.

4. Pyrometric Temperature Measurements on Thermal Protection Systems, 

GAMM 98 (Gesellschaft für angewandte Mathematik und Mechanik) Bremen, 

April 1998, edit. Zeitschrift f. angewandte Mathematik u. Mechanik (ZAMM) 79, 

Suppl, 3, pages S945-S946, 1999. G. Herdrich, M. Auweter-Kurtz, H. Habiger.

Huygens

1. Jones, J.C.; and Giovagnoli, F.: Huygens: Science, Payload and Mission. A 

Wilson, editor, ESA-SP-1177, Aug 1997. 

2. Baillion, M.; and Pallegoix, J.F.: Huygens Probe Aerothermodynamics, 

Aerospatiale Espace and Defense copyrighted document, AIAA-1997-2476, 32nd 

Thermophysics Conference, Atlanta GA, Jun 1997. 

https://doi.org/10.2514/6.1997-2476

3. Bouilly, J.M.; and Guerrier, D.: Entry Testing of AQ60 for Huygens. Presented 

at the First ESA/ESTEC Workshop on Thermal Protection Systems, Noordwidjk, 

The Netherlands, May 1993. 

4. Wright, M.J.; Olejniczak, J.; Walpot, L.; Raynaud, E.; Magin, T.; Callaut, L.; 

and Hollis, B.: A Code Calibration Study for Huygens Entry Aeroheating. AIAA2006-0382, 44th Aerospace Sciences Meeting, Reno NV, Jan 2006. 

https://doi.org/10.2514/6.2006-382

5. Lingard, J. and Underwood, J., "Huygens DCSS post flight reconstruction”, 

AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, 

Williamsburg, VA, 2007. (AIAA 2007-2579) https://doi.org/10.2514/6.2007-2579

6. Bouilly, J-M: Thermal Protection of the Huygens Probe during Titan Entry: Last 

Questions. Presented at 2nd International Planetary Probe Workshop, Moffett 

Field, CA, Aug. 2004. 

7. Huygens Mission Operation Report. HUY-OPS-RP-1006-OPS-OFH Iss. 1, Feb 

2005.

8. Colombatti et al., Reconstruction of the trajectory of the Huygens probe using 

the Huygens Atmospheric Structure Instrument (HASI). Planet. Space Sci. 56(5), 

586-600, 2008. https://doi.org/10.1016/j.pss.2007.11.017

9. Kazeminejad et al., Huygens’ entry and descent through Titan's atmosphere—

Methodology and results of the trajectory reconstruction. Planet. Space 

Sci. 55(13), 1845-1876, 2007. https://doi.org/10.1016/j.pss.2007.04.013

ARD 'Atmospheric Reentry Demonstrator’

1. The Atmospheric Reentry Demonstrator, ESA publication BR-138, Oct 1998. 

2. Johnston, I.A.; Weiland, M.; Schramm, J.M.; Hannemann, K.; and Longo, K.: 

Aerothermodynamics of the ARD: Postflight Numerics and Shock-Tunnel 

Experiments. AIAA-2002-0407, 40th Aerospace Sciences Meeting, Reno NV, Jan 

2002. 

3. Tran, P.; and Soler, J.: Atmospheric Reentry Demonstrator Post Flight 

analysis: Aerothermal Environment. Proceedings of the 2nd Intl Symposium on 

Atmospheric Reentry Vehicles, Arachon, France Mar 2001.

4. Tran, P.; Paulat J.C.; Boukhobza, P.; Re-entry Flight Experiments Lessons 

Learned – The Atmospheric Reentry Demonstrator ARD, RTO-EN-AVT-130.

Deep Space 2

1. Mitcheltree, R.A.; DiFulvio, M.; Horvath, T.J.; and Braun, R.D.: Aerothermal 

Heating Predictions for Mars Microprobe. AIAA-1998-0170, 36th Aerospace 

Sciences Meeting, Reno NV, Jan 1998. 

Stardust

1. Olynick, D.; Chen, Y.K.; and Tauber, M.E.: Aerothermodynamics of the 

Stardust Sample Return Capsule. Journal of Spacecraft and Rockets, Vol 36, No 

3, May-Jun 1999, pp 442-462. 

2. Mitcheltree, R.A.; Wilmoth, R.G.; Cheatwood, F.M.; Brauckmann, G.J.; and 

Greene, F.A.: Aerodynamics of Stardust Sample Return Capsule. AIAA-1997-

2304, 15th Applied Aerodynamics Conference, Atlanta GA, Jun 1997.

3. Brandis, A.M.; Johnston, C.O; Characterization of Stagnation-Point Heat Flux 

for Earth Entry. AIAA-2014-2374, 45th AIAA Plasmadynamics and Laser 

Conference, Atlanta GA, Jun 2014.

4. AIAA-2007-2546 Stardust Parachute Trajectory Performance Reconstruction.

Genesis

1. Lo, M.W.; Williams B.G.; Bollman E.; Han, D.; Hahn, Y.; Bell, J.L.; Hirst, E.A.; 

Corwin, R.A., Hong, P.E.; and Howell, K.C.: Genesis Mission Design. AIAA-1998-

4468, AIAA/AAS Astrodynamics Specialist Conference and Exhibit, Boston MA, 

Aug 1998. 

2. Cheatwood, F.M.; Merski, N.R.; Riley, C.J.; and Mitcheltree, R.A.: 

Aerothermodynamic Environment Definition for the Genesis Sample Return 

Capsule. AIAA-2001-2889, 35th AIAA Thermophysics Conference, Anaheim CA, 

Jun 2001

3. Tang, C. and Wright, M., "Analysis of the Forebody Aeroheating Environment 

during Genesis Sample Return Capsule Reentry”, AIAA Aerospace Sciences 

Meeting and Exhibit, Reno, NV, 2007. (AIAA 2007-1207)

— 42 —
4. Jenniskens, P., et al, “Surface Heating from Remote Sensing of the 

Hypervelocity Entry of the NASA GENESIS Sample Return Capsule”, AIAA 

Aerospace Sciences Meeting and Exhibit, Reno, NV, 2006. (AIAA 2006-381)

5. AIAA-2001-20017 Parafoil Recovery Subsystem for the Genesis Space Return 

Capsule.

Hayabusa

1. Ishii, N.; Hiraki, K.; Yamada, T.; Inatani, Y.; and Honda, M.: System 

Description and Reentry Operation Scenario of MUSES-C Reentry Capsule. 

Institute of Space and Astronautical Science Report, SP No 17, edited by Y. 

Inatani, Mar 2003, pp 389-400.

Mars Exploration Rovers "Spirit" and "Opportunity" 

1. Roncoli, R.B.; and Ludwinski, J.M.: Mission Design Overview for the Mars 

Exploration Rover Mission. AIAA-2002-4823, AIAA/AAAS Astrodynamics 

Specialist Conference and Exhibit, Monterey CA, Aug 2002. 

2. Szalai, C.; Chen Y.; Loomis, M; and Hui, F.: Mars Exploration Rover TIRS 

Cover Thermal Protection System Design Verification. AIAA-2003-3767, 36th 

AIAA Thermophysics Conference, Orlando FL, Jun 2003. 

3. “Mars Exploration Rover Aeroshell, Critical Design Review” document, 30-31 

May 2001 by JPL, NASA and Lockheed Martin 

4. AIAA-2005-1605 "Mars Exploration Rover Parachute System Performance" 

18th AIAA Aerodynamic Decelerator Systems Technology Conference, May 

2005, Munich Germany

Beagle 2

1. Burnell, S.; Liever P.; Smith, A.; and Parnaby G.: Prediction of the Beagle 2 

Static Aerodynamic Coefficients. Atmospheric Reentry Vehicles and Systems, 

Arcachon 2000. 

2. Smith A.; Parnaby, G.; Jones, T.V.; and Buttsworth, D.: Aerothermal 

Environment of the Beagle 2 Probe. Fourth European Symposium on 

Aerothermodynamics, Capua, Italy, Oct 2001. 

3. Liever, P.A.; Habchi, S.D.; Burnell, S.I.; and Lingard, J.S.: Computational Fluid 

Dynamics Predictions of the Beagle 2 Aerodynamics Database. Journal of 

Spacecraft and Rockets, Vol 40, No 5, Sep-Oct 2003. 

4. C.T. Pillinger with M.R. Sims and J. Clemmet, The Guide to Beagle 2. © C.T. 

Pillinger in association with Faber and Faber, 2003.

5. Bertrand, J., Labaste, V. and Mignot, Y., The Beagle 2 Thermal Protection 

System. 3rd International Symposium on ARVS, Arcachon, France, March 2003.

6. Clemmet, J., Sims, M.R., Higgett, N., Kuzmanova, T., Tatham, E., Pullan, D., 

Muller, J.-P., Putri, A., Bridges, J., Pillinger, J.M., Beagle 2 on Mars – The 

Discovery Assessed. J. Brit. Interplanet. Soc. 70, pp.261-304, 2017.

7. AIAA-2003-2153 Design and Development of the Main Parachute for the 

Beagle 2 Mars Lander.

Phoenix

1.Ryan McDaniel, Michael Wright, and Jarvis Songer. “Aeroheating Predictions 

for Phoenix Entry Vehicle”, AIAA 2008-1279, 46th AIAA Aerospace Sciences 

Meeting and Exhibit. January 2008.

2. Allen Witkowski, Mike Kandis, and Douglas Adams. “Mars Scout Phoenix 

Parachute System Performance”, AIAA 2009-2907. 20th AIAA Aerodynamic 

Decelerator Systems Technology Conference and Seminar. May 2009.

3. Karl T. Edquist, Prasun N. Desai, and Mark Schoenenberger. “Aerodynamics 

for Mars Phoenix Entry Capsule. “Journal of Spacecraft and Rockets 2011 48:5, 

713-726.

4. Myron Grover, Prasun Desai, and Benjamin Cichy. “Overview of the Phoenix 

Entry, Descent and Landing System Architecture,” AIAA 2008-7218. AIAA/AAS 

Astrodynamics Specialist Conference and Exhibit. August 2008.

Mars Science Laboratory

1. Dutta, S.; Braun, R.D.: Statistical Entry, Descent, and Landing Performance 

Reconstruction of the Mars Science Laboratory. AIAA-2014-0385, AIAA 

Atmospheric Flight Mechanics Conference, National Harbour, MD, January 2014.

2. Karlgaard, C.; Kutty, P.; Schoenenberger, M.; Munk, M.M.; Little, A.; Kuhl, 

C.A.; and Shidner, J.: Mars Science Entry Atmospheric Data System Trajectory 

and Atmospheric Reconstruction. Journal of Spacecrafts and Rockets, Vol. 51, 

No.4, July-August 2014, pp. 1029-1047.

3. Beck, R.A.S.; Driver, D.A.; Wright, M.J.; Hwang, H.H.; Edquist, K.T.; and 

Sepka, S.A.: Development of the Mars Science Laboratory Heatshield Thermal 

Protection System. Journal of Spacecrafts and Rockets, Vol. 51, No.4, JulyAugust 2014, pp. 1139-1150. Wright, M.J.; Beck, R.A.S.; Edquist, K.T.; Driver, D.; 

4. Sepka, S; Slimko, S.A.; and Willcockson, W.H.: Sizing and Margins 

Assessment of Mars Science Laboratory Aeroshell Thermal Protection System. 

5. Bose, D.; White, T.; Mahzari, M.; and Edquist, K,: Reconstruction of 

Aerothermal Environment and Heat Shield Response of Mars Science 

Laboratory. Journal of Spacecrafts and Rockets, Vol. 51, No.4 , July-August 

2014, pp. 1174-1184. 

— 43 —
6. Cruz, J.R.; Way, D.W.; Shidner, J.D.; Davis, J.L.; Adams, D.S.; and Kipp, 

D.M.: Reconstruction of the Mars Science Laboratory Parachute Performance. 

Journal of Spacecrafts and Rockets, Vol. 51, No.4, July-August 2014, pp. 1185-

1196.

7. AIAA-2013-1277 “Mars Science Laboratory Parachute System Performance”.

InSight

1. Robin A.S. Beck, et al., “InSight Aerotherml Environment Assessment.” Journal 

of Spacecraft and Rockets, Vol. 58, No. 6, November-December, 2021. DOI 

number 10.2514/1.A35078.

2. Ian G. Clark, et al., ”Reconstructed Performance of the Supersonic Parachute 

of the Mars InSight Lander,” Journal of Spacecraft and Rockets, Vol. 58, No. 6, 

November-December 2021. DOI number 10.2514/1.A35180.

3. AIAA-2019-3481 “Mars InSight Parachute System Performance”.

Inflatable Reentry Vehicle Experiment-3 (IRVE-3) 

1.Olds, A.D.; Beck, R.E.; Bose, D.M.; White J.P.; Edquist, K.T.; Hollis, B.R.; 

Lindell, M.C.; Cheatwood, F.M.; Gsell, V.T.; and Bowden, E.L.: IRVE-3 PostFlight Reconstruction, Aerodynamic Decelerator Systems Technology AIAA 2013-

1390

Hayabusa 2 

1. Yamada, T.; Yoshihara, K.; Yamada, K.; and Shimoda, T.: Development of a 

Hayabusa-2 Sample Return Capsule, Presented July 10, 2015, 2015-k-62

Exploration Flight Test-1

1. Details on where data can be found in a public location: Woollard, B.A.; Braun, 

R.D.; and Bose, D.: Aerothermodynamic and Thermal Protection System 

Instrumentation Reference Guide. IEEE-2016-2104, 2016 IEEE Aerospace 

Conference, Big Sky, MT, Mar 2016. 

2. Sinclair, R.: A Status Report on the Parachute Development for NASA's Next 

Manned Spacecraft. NTRS 20080015895, Safety and Flight Equipment 

Symposium, Genena, Jan 2008.

3. Bibb, K., Robinson, P., and Cassady, L., “Reconstruction of the Orion Crew 

Module Aerodynamics for theft-1 Flight Test”, AIAA Aviation Forum 2016, 

Washington, D.C., June 2016. 

— 44 —

ExoMars 2016

1. Tolker-Nielsen, T.: Exomars 2016 - Schiaparelli Anomaly Inquiry. DGI/2017/546/TTN, May 2017. 

https://exploration.esa.int/web/mars/-/47852-entry-descent-and-landingdemonstrator-module

https://spaceflight101.com/exomars/Schiaparelli-edm/

2. Gülhan, A.; Thiele, T.; Siebe, F.; Aerothermal Measurements from the 

Exomars Schiaparelli Capsule Entry. Journal of Spacecraft and Rockets. Vol. 56, 

No. 1, January-February 2019, pp. 68-81. https://doi.org/10.2514/1.A34228

3. Pinaud, G.; Bertrand, J.; Soler, J.; Tran, P.; Ritter, H.; and Bayle, O.: Exomars

mission 2016: A preliminary post-flight analysis of the heat shield during entry on 

Mars Atmosphere. AIAA SciTech Forum, San Diego CA, Jan 2019. 

https://doi.org/10.2514/6.2019-0244

4. Brandis, A.M.; White, T.R.; Saunders, D.A.; Hill, J.P.; and Johnston, C.O.; 

Simulation of the Schiaparelli Entry and Comparison to Aerothermal Flight Data. 

AIAA 2019-3260, AIAA Aviation Forum, Dallas TX, Jun 2019. 

https://doi.org/10.2514/6.2019-3260

5. Bonetti, D.; De Zaiacomo, G.; Blanco, G.; Pontijas Fuentes, I.; Portigliotti, S.; 

Bayle, O.; Lorenzoni, L.; ExoMars 2016: Schiaparelli coasting, entry and descent 

post flight mission analysis. Acta Astronautica 149 (2018) 93-105. May 2018. 

https://doi.org/10.1016/j.actaastro.2018.05.029

6. Ball, A.J., Blancquaert, T., Bayle, O., Lorenzoni, L.V., Haldemann, A.F.C. and 

the Schiaparelli EDM Team: The ExoMars Schiaparelli Entry, Descent and 

Landing Demonstrator Module (EDM) System Design. Space Sci. Rev. 218:44, 

2022. https://doi.org/10.1007/s11214-022-00898-z

OSIRIS-REx

1. Beshore, Edward, et al., The OSIRIS-REx Asteroid Sample Return Mission, 

2015 IEEE Aerospace Conference, 7-14 March 2015.

2. Lauretta, D.S.; Balram-Kutson, S.S.; OSIRIS-Rex: Sample Return from 

Asteroid (101955) Bennu, Space Sci Rev 212, 925-984 (2017).

3. Olynick, D.; Chen, Y.K.; and Tauber, M.E.: Aerothermodynamics of the 

Stardust Sample Return Capsule. Journal of Spacecraft and Rockets, Vol. 36, No 

3, May-Jun 1999, pp 442-462.

4. AIAA-2007-2546 Stardust Parachute Trajectory Performance Reconstruction 

(for PRS system sizing only).
MARS 2020

1. Karlgaard, C., et al, "Mars Entry, Descent, and Landings Instrumentation 2 

Trajectory, Aerodynamics, and Atmosphere Reconstruction," AIAA SciTech 2022, 

AIAA Atmospheric Flight Mechanics Conference, San Diego, CA, 2022. (AIAA 

2022-0423)

2. White, T., et al, "Mars Entry Instrumentation Flight Data and Mars 2020 Entry 

Environments" AIAA SciTech 2022, AIAA Atmospheric Flight Mechanics 

Conference, San Diego, CA, 2022. (AIAA 2022-0011)

3. Way, D., et al, "Assessment of the Mars 2020 Entry, Descent, and Landing 

Simulation" AIAA SciTech 2022, AIAA Atmospheric Flight Mechanics Conference, 

San Diego, CA, 2022. (AIAA 2022-0421)

4. Dutta, S., et al, "Post-flight Analysis of Atmospheric Properties from Mars 2020 

Entry, Descent, and Landing", AIAA SciTech 2022, AIAA Atmospheric Flight 

Mechanics Conference, San Diego, CA, 2022. (AIAA 2022-0422)

LOFTID

1. Dillman, R.A., DiNonno, J.M., Bodkin, R.J., Hughes, S.J., Cheatwood, F.M., 

Blakeley, H., Akamine, R.L., and Bowes, A., "Planned Orbital Flight Test of a 6m 

HIAD," 15th International Planetary Probe Workshop, Boulder, CO, 2018.

2. Swanson, G., Smith, B., Akamine, R., Bodkin, R.J., Cheatwood, N., Hughes, 

S., Gaddy, D., Chan, P., and Parker, A., "The HIAD Orbital Flight Demonstration 

Instrumentation Suite," 15th International Planetary Probe Workshop, Boulder, 

CO, 2018.

3. Hughes, S., Cheatwood, N., Swanson, G., Kazemba, C., Lindell, M., Tobin, S., 

Bodkin, R.J., Dillman, R., DiNonno, J., and Johnson, K., "Low-Earth Orbit Flight 

Test of an Inflatable Decelerator (LOFTID) - Aeroshell System Overview," 16th 

International Planetary Probe Workshop, Oxford, UK, 2019.

4. Swanson, G., Kazemba, C., Hughes, S., Williams, J., Johnson, K., Lindell, M., 

Bodkin, R.J., DiNonno, J., and Cheatwood, N., "Low-Earth Orbit Flight Test of an 

Inflatable Decelerator (LOFTID) - Aeroshell Engineering Development Unit 

Structural Testing," 16th International Planetary Probe Workshop, Oxford, UK, 

2019.

ARTEMIS I

References will be supplied in a future release.

— 45 —