planetary edl overview
TRANSCRIPT
Planetary Entry, Descent and Landing
Dr. Robert D. BraunGeorgia Institute of Technology
Scope
• Overview of aeroassist technology – First principles– Review of community accomplishments– Predictions for technology development
• Designed for program managers, systems engineers and disciplinary specialists interested in gaining a working knowledge (technical overview) of planetary atmospheric flight
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Seminar Format
• 15 hours• Highly interactive• Lecture material prepared to allow ample
time for questions and discussion• Large bibliography provided for further study
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Outline
• Introduction and Definitions– Aeroassist (Aerocapture, Aerobraking, Entry)– Typical sequence of events– Entry velocity (Vatm)– Entry flight path angle (γ)– Ballistic coefficient (β)– Reynolds, Mach and Knudsen numbers– Aerodynamic regimes– Aerodynamic coefficients (L/D)– Heat rate, heat load, dynamic pressure– Terminal velocity
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Outline• Back of the envelope calculations (Earth & Mars
examples)– Entry velocity from Vinf– Newtonian aerodynamics– Equations of motion– Terminal descent– Heating– Landing accuracy
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Outline• Key technologies and trades
– Approach Navigation– Thermal Protection System– Deployable Systems– Atmospheric GN&C – Terminal Descent System– Landing Systems
• Breaking out of the Viking technology box– Large mass robotic landers and human exploration
• Summary of aeroassist technology readiness• Simulations – inputs/outputs & methods• Test facilities
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Outline
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• Historical experience/case studies/project comparisons– Mars Landers
• Viking• MPF• MPL• MER• Phoenix
– Entry Probes• Pioneer Venus• Galileo• MSR EEV • DS-2• Huygens• Stardust• Genesis
– Aerobraking Spacecraft• Magellan• MGS• Odyssey• MRO
Outline• Future expectations
– Technologies ready for flight– Project needs
• Reference List• Aeroassist contacts
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Contact Information
http://www.ae.gatech.edu/~rbraun
http://www.ssdl.gatech.edu
http://www.ae.gatech.edu/~rbraun/PlanetaryEDL.pdf
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Acknowledgement• This material has been compiled by the author, based largely on
accomplishments of the Mars Pathfinder, Mars Polar Lander, Mars Microprobe, Mars Sample Return and Mars Surveyor 2001 flight project teams. As a member of these teams, the author has had the privilege of working with some of the most talented engineers in the country over the past decade. Team members include personnel at the Jet Propulsion Laboratory, NASA Langley Research Center, NASA Ames Research Center, Lockheed-Martin Astronautics, and Pioneer Aerospace.
• In addition, advances made by members of the Mars Exploration Rover and Mars Science Laboratory EDL teams, by NASA Capability Roadmap teams, by the author in recent studies of human exploration systems and by personnel at Ball Aerospace and Marshall Space Flight Center in regard to inflatable entry systems are included.
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Aeroassist Technology: Mission Classes
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Aeroassist Technology• Aeroassist systems span a wide range of
applications in which aerodynamic forces are used to improve or enable a mission concept that includes flight through a planetary atmosphere– Deceleration, acceleration, improved control
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Aeroassist Technology Applications• Entry (Landers)
– Entry into a planetary atmosphere from hyperbolic approach or planetary orbit
• Aerobraking (Orbiters)– Used after orbit insertion to trim science orbit– Multiple passes through the high atmosphere– Performed at sufficiently low density to eliminate heatshield
• Aerocapture (Orbiters)– Decelerates from hyperbolic approach to orbital velocity in a single
pass– Orbit control by aerodynamic lift/drag modulation
• Aerogravity Assist (Transfer Vehicles)– Used during interplanetary transfer to reduce trip time or propulsive
requirements– Similar to gravity assist, except dips into sensible atmosphere
providing larger change in velocity– Requires high lift-drag ratio vehicle capable of withstanding high
heatingRDB Aug 2005
13
Aeroassist Technology Readiness• Applications listed in order of technology readiness
• Entry and Aerobraking missions have both been accomplished multiple times– Technology investment required to enhance their proven capability
• Strictly speaking, Aerocapture has not been performed– Technology elements have been demonstrated and in place since ‘60s– Planned for Mars Surveyor 2001 Orbiter. Descoped after 18 months.– Significant mission benefits for some lunar-return/Mars mission
architectures and for outer planet exploration– Possible New Millennium technology demonstration project
• Aerogravity assist is most distant application and requires substantial investment in technology to realize its potential
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Direct EntryDirect entry is flight into the planet’s atmosphere from hyperbolic approach or orbit. The entry vehicle can be passive (ballistic) or actively controlled. The passive vehicle is guided prior to atmospheric entry and proceeds into the planet’s atmosphere as dictated by the vehicle shape and the atmosphere. An actively controlled direct entry vehicle may maneuver autonomously while in the atmosphere to improve landed location, or modify the flight environment.
Successfully performed on Viking, Apollo, Shuttle, Pioneer-Venus, Galileo, Mars Pathfinder and MER
MER Entry System
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Example Mars Ballistic Entry Corridor
100 km
Parachute Deploy Altitude 8 km
“Earth’s Mount Everest” (8.5 km)
< 20 kmSafety Corridor
MPF Landing Ellipse300 km by 50 km (For MER it was more like 80 km x 20 km)
The so-called “entry flight path angle”11.5 ° +/- 0.75° (MER)
Picture is to scale!
Peak Deceleration 8 Earth g’s “Entry Point”
128 km above the surface
Courtesy Rob Manning, JPL
Terminal Descent
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Lander Separation
Heatshield Separation
Parachute Deployment & Inflation
Bridle Deployment
Terminal Descent and Landing
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Bridle Deployment
Radar Acquisition of the Surface
Airbag Inflation & RAD firing
Airbag Bounces
Terminal descent imagery
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MER Entry, Descent and Landing
Mars EDL HistoryThere have only been five successful landings on Mars
– 2 Vikings in ‘76, Mars Pathfinder in ‘97, 2 MERs in ‘04– There have been at least as many failures
All five of these systems– Had touchdown masses < 0.6 MT– Landed at low elevation sites, below -1 km MOLA– Had landed footprints on the order of 100s of kms (unguided)
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Current EDL Technology Landed Elevation Capability
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-4
-3
-2
-1
0
1
2
3
4
500 700 900 1100 1300 1500
Delivered Mass (kg)
Optimized EDL
• The landing elevation capability of our current EDL systems drops by approximately 1 km for every 100 kg of added useful landed payload.– This has large implications for future plans for both robotic and human
exploration• An L/D on the order of 0.25 may be used to provide as much as a 3 km
surface elevation advantage.
Assumes large aeroshell (5 m) and use of Viking heritage parachute
Surface Elevation
(km)
L/D = 0.25
Ballistic
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Aerobraking• Aerobraking is a relatively low risk
maneuver that consists of repeated dips into an atmosphere to generate drag and lower velocity. This is done AFTER a vehicle has been inserted into an initial orbit to adjust the eccentricity of the orbit or to simply lower the orbit. Large performance margins are maintained to accommodate significant atmospheric variability. Generally, the total heat flux and peak temperatures are low enough to fly without a thermal protection system.
• Magellan was the first planetary spacecraft to use this technique. Also successfully employed on MGS (even with an anomaly) and Odyssey.
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Aerobraking Systems• Aerobraking employs atmospheric drag to reduce orbit energy (apoapsis) in
repeated passes through the upper atmosphere (near periapsis).– Originally demonstrated by Atmospheric Explorer-C (Earth) and later by Magellan (Venus).
• Significantly reduces the necessary propellant for orbit insertion, thus allowing a reduction in launch mass and potential launch cost savings.
• The primary drag surface for aerobraking is typically the orbiter solar panel(s).– Maximum allowable heat rate is constrained by solar panel thermal limitations (Example:
Odyssey not-to-exceed temperature on the solar panel was 175°C, which translated to a max heating rate of about 0.6 W/cm2 during aerobraking main phase).
• Atmospheric density uncertainty is a major risk factor.– At Mars, heat rate margins of 100% are used to accommodate large orbit-to-orbit density
variations.• Despite advancements in aerobraking automation, aerobraking remains a human-
intensive process.– 24 hr/day operations for weeks or months– Up to 4 sequence uploads per day– Detailed interaction between navigation, spacecraft team, sequencing, atmosphere
advisory group, and mission management.
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Aerobraking Design
Past aerobraking missions at Mars have divided aerobraking into three distinct phases: Walkin, Main Phase, and Walkout (aka “Endgame”).
Walkin: Orbit periapsis altitude is gradually reduced until (a) contactwith the atmosphere is established, and (b) Periapsis heating rates are within the desired heat rate corridor.
Main Phase: The majority of orbit period reduction is obtained during main phase. Target heating rates are at their highest level.
Walkout: As the orbit period is reduced, drag pass durations become longer. During walkout, the periapsis altitude is gradually increased, in order to maintain a desired orbit lifetime (typically, ~24-48 hrs) in the event ground communication is lost. Walkout completes the aerobraking phase with a maneuver to raise periapsis out of the atmosphere.
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Rate DampThrough drag pass on
Loose Deadbands
5 minuteGuardband
Accel Bias Calc@ Drag start Š 30 min
TelemetryPlayback (2)
RWAs to Tach Profile ŅFree DesatÓ
Start PTEPower 2ndary Gimbals
Transition to Thruster Control
Reconfigure TelecomLGA, Carrier only
@ Drag start Š 15 min
Slew to Drag Attitude@ Drag start Š 10 min
5 minuteGuardband
Stop PTETurn Off 2ndary Gimbals
Back to RWA ControlSlew to Vacuum Attitude
Back to Earthpoint@ Drag End + 10 min
Reconfigure Telecom back to HGAAccel Bias Calc
TelemetryPlayback (3)
TelemetryPlayback (1)
Drag Pass Overview
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Mars Odyssey Periapsis Altitude During Aerobraking
90
100
110
120
130
140
150
160
24-Oct 31-Oct 7-Nov 14-Nov 21-Nov 28-Nov 5-Dec 12-Dec 19-Dec 26-Dec 2-Jan 9-Jan 16-JanDate
hp(km)
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Atmospheric Density Variations During Odyssey Aerobraking
0
0.25
0.5
0.75
1
1.25
1.5
1.75
2
2.25
24-Oct 31-Oct 7-Nov 14-Nov 21-Nov 28-Nov 5-Dec 12-Dec 19-Dec 26-Dec 2-Jan 9-Jan 16-JanDate
ρ/ρ0
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Mars Odyssey Orbit Period During Aerobraking
0
2
4
6
8
10
12
14
16
18
20
24-Oct 31-Oct 7-Nov 14-Nov 21-Nov 28-Nov 5-Dec 12-Dec 19-Dec 26-Dec 2-Jan 9-Jan 16-Jan 23-JanDate
Actual Plan
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Aerocapture• Aerocapture is a maneuver designed to take advantage of a planet’s
atmosphere to slow a spacecraft to orbital capture velocities and results in a substantial propellant reduction. This mass savings generally translates into smaller launch vehicles.
• The maneuver begins with a shallow approach angle to the planet. An autonomous guidance and control system modulates the vehicle’s aerodynamics to mitigate off-nominal atmospheric conditions. Descent into the relatively dense atmosphere causes sufficient deceleration and heating to require a heatshield.
• Upon atmospheric exit, the heat shield is jettisoned and a propulsive maneuver is performed to raise the periapsis. The entire operation is short-lived and requires the spacecraft to operate autonomously while in the planets atmosphere.
• Demands placed on the vehicle depend greatly on the specifics of the planet being approached and the mission. Key variables includeatmospheric properties, desired orbit insertion geometry, interplanetary approach accuracy, entry velocity, and vehicle geometry.
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Often Proposed, Yet to Be Implemented• Aerocapture systems have been proposed, planned and developed many
times. To date, no flight system has implemented this aeroassist maneuver.
• Notes:– Apollo entry guidance was implemented with an aerocapture logic branch.
However, this guidance mode was never executed in-flight.– Hypersonic aeromaneuvering is a common subset of both aerocapture and
precision/pinpoint landing. This guidance mode was successfully demonstrated by the Apollo program.
Mission Timeframe Completed Mission Development
Termination Cause
New Millennium ST-7 1999-2002 Phase A Not selected
AFE 1984-1989 Phase D Mass, Cost, STS Use
MSP’01 Orbiter 1996 - 2000 Phase B Perceived riskMars Sample Return 1998 - 2000 Phase A MSR delayed
New Millennium ST-9 2004-present Phase A Ongoing
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Aerocapture Case Study: MSP’01 Orbiter
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• Mission planning initiated in 1996.• Aerocapture baselined at MSP’01 project start (1997).• Aerocapture selected for MSP’01:
– Higher launch mass margin (relative to aerobraking or prop capture)– Reduced launch vehicle cost (relative to aerobraking or prop capture)– Improved science return (relative to aerobraking)– Technology feed-forward (MSR and other planets)
• At time of MCO and MPL failures, development of MSP’01 aerocapture orbiter was on schedule. Phase B complete.
• Switch made to propulsive capture/aerobraking (Mars Odyssey) in 1999 based on:– Perceived risk in hypersonic aeromaneuvering and subsequent
autonomous sequences– Schedule and spacecraft development risk concerns– Thinking that MSR would go to propulsive capture/aerobraking
MSP’01 Aerocapture Orbiter
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Aerocapture Case Study: MSP’01 Orbiter Timeline
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Periapsis Raise (Burn)
Exit Orbit 500 x -100 km
Entry Interface
Intermediate Orbit 500 x 200 km
Crz Stage Jett
E - 10 min
Entry Interface (125 km)
E - 0 min
Exit (125 km)
E+11.2 m
Max g-Load
E+2.4 m
Periapsis Raise
E + 45 min
Atmospheric Flight
AeroshellJettison
E+11.7 m
Note: Times are Representative
MSP’01 Orbiter Aerocapture Trajectory Data
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MSP’01 Aerocapture Configuration – End of Phase B
• Launch mass (wet), CBE + contingency 647 kg• Cruise stage, CBE + contingency 71 kg• Earth-to-Mars cruise propellant 32 kg
• Aerocapture entry mass, CBE + growth 544 kg
• Heatshield mass, CBE + growth (2.65 m diam) 122 kg• Backshell mass, CBE + growth 75 kg• Total entry system mass 197 kg
(includes all structures and mechanisms)
• Post-Aerocapture periapsis-raise maneuver 20 kg(400 km circular orbit)
• “Payload” mass at Mars 327 kg• Payload Mass/Launch Mass 0.51
RDB Aug 200540
Science Data Volume Comparison
MSP’01 Mission: Aerobraking vs. Aerocapture
1 Mars-Year Science Mission Lifetime (687 days)
Total DataVolume
GRS
THEMIS
Aerobraking Aerocapture PIP Values
65 Gbits
350 Gbits
40 Gbits 40 Gbits
280 Gbits 370 Gbits
New Baseline Old Baseline
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Mission Modes: Aerocapture vs. Propulsive Capture/Aerobraking
Capture
Orbit Trim
Deployments
Direct-Entry- Pathfinder, MSP’98 Heritage
Autonomous Aeromaneuvering- Apollo, STS Heritage
Heatshield, Backshell Jettison- Pathfinder, MSP’98 Heritage
Solar Array DeploymentHGA Deployment
Autonomous Periapsis-RaiseManeuver
1700 M/S MOI Burn, Bi-Prop System
- MGS, MSP’98 Heritage
70 Days of Aerobraking- MGS, MSP’98 Heritage
Solar Array Deployment
HGA Deployment
Aerocapture Propulsive Capture/Aerobraking
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Return Orbiter AerocaptureReturn Orbiter Aerocapture
Aerocapture Corridor• Flyable corridor (shaded) for deceleration defined by:
– Lift down trajectory– Trajectory flying lift up until reaching the deceleration limit (5 g),
then banking to achieve the desired exit energy
Aerocapture Trajectories(100 T, 10 m, L/D = 0.3, ventry = 7.5 km/s)
0
20
40
60
80
100
120
140
0 1 2 3 4 5 6 7 8
Velocity (km / s)
Alti
tude
(km
)
Lift Down
Lift Up to 5 g Limit,Bank to OrbitLift Up
PhysicalCorridor
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• Physical Corridor (shaded) showing entry flight path angle range for:– 100 T– 10 m aeroshell– L/D = 0.3
• Boundaries:– Lift-up– Lift-down– 5 g limit, 0º bank angle– Velocity– Heat rate constraints
Aerocapture Corridor Width
5 6 7 8 9 10 11 12-20
-15
-10
-5
0100 T 10 m L / D = 0.3
Entry Velocity (km / s)
Flig
ht P
ath
Ang
le (º
)
Lift DownLift Up5 g (lift up)Velocity100 W / cm2
500 W / cm2
Robotic Mission Benefits of Aerocapture
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0
5 0
10 0
15 0
20 0
25 0
30 0
M a rs1 2 .4 k m /s
V e n u s2 3 .3k m /s
T i ta n 1 4 .4k m /s
U ra n u s2 4 .5k m /s
V e n u s1 4 .6k m /s
N e p tu n e 26 .0 km /s
S a tu rn 1 8 .0k m /s
J u p ite r11 7 .0 k m /s
% S
avin
gs
% m as s s aving s% c o s t s a vin gs
AerocaptureEnabled
AerocaptureEnhanced
Ref: Cost-Benefit Analysis of the Aerocapture Mission Set;Hall, J.L.; Noca, M.A.; Bailey, R.W.; Journal of Spacecraft & Rockets, Vol. 42, No. 2, 2005, pp. 309-320.
The Value of Aerocapture and Other Technology Investments for Human Mars Exploration
0.0
1.0
2.0
3.0
4.0
5.0
6.0
7.0
8.0
Mas
s Sav
ings
Nor
mal
ized
to IS
S M
ass
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Advanced Propulsion
Closed Loop Life Support
Advanced Materials
Maintenance & Spares
Advanced Avionics
Aerocapture
All Propulsive Chemical
Today
NOTES:• Results are cumulative and thus trends will be different
for different technology combinations/sequences• The change between points shows the relative mass
savings for that particular technology• 2018 One-Year Round-Trip Mission, Crew of 4,
Lander pre-deployed
Courtesy K. Joosten, Johnson Space Center
Aerocapture Also Enables Shorter Transit TimesFor 200-Kg Arrival Mass at Neptune
0
50
100
150
200
4567891011
Use
ful I
nser
ted
Mas
s (k
g)
Trip Time to Neptune (Years)
Target
Aerocapture
Propulsion
Non-Deceleration System Mass into Orbit
*Courtesy of Paul Wercinski, NASA Ames Research Center
RDB Aug 200548
Deployable Entry Systems: When Size Matters
Trailing ballute concept(Courtesy Ball Aerospace)
Attached ballute concept
Generally considered for direct entry or aerocapture missions with high mass payloads
RDB Aug 200549
Aerogravity AssistHigh lift (L/D = 3 ) configuration
Payload volume
TPS
Sharp Leading Edge
Aero Control Surfaces
• Never performed• No currently planned missions• Provides
• Shorter trip times• V for accelerating s/c
• Requires • Low drag vehicle (wave-rider configuration)• Ultra-high performance TPS• Efficient packaging of s/c
RDB Aug 200550
Aerogravity assist is an extension of gravity assist. It differs from conventional gravity assist in that the spacecraft performs part of its flyby within the planetary atmosphere. While in the atmosphere, the vehicle’s aerodynamic forces are used to further rotate the heliocentric velocity vector, resulting in a potentially large ∆V. To perform aerogravityassist without a large drag penalty, a high L/D vehicle with sharp, non-eroding leading edges is required. The sharp leading edge requirement necessitates the development of new materials that can be manufactured with a very small radius of curvature and are resistant to extremely high heating rates. Aerogravity assist also requires micro-spacecraft that can be packaged within the low available volume. Aerogravity assist would significantly reduce the trip time to the outer solar system planets.
Benefit of Aerogravity Assist
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The synergistic use of both gravity and aerodynamics can significantly increase the heliocentric velocity turn angle, resulting in a larger ∆V
Aeroassist Mission Summary
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Mission Type Launch Aeroassist CommentApollo E 65-69 65-69
7678
81-pres
95
9797-98
01
04
Cassini Huygens Probe E 97 0406040507
Active control; Aerocapture logic
Viking Landers E 75 Entry from orbit with active controlPioneer-Venus Probes E 78Space Shuttle E 81-pres Landing and crossrange
requirements drove geometry
Magellan AB AB performed after science mission
Mars Global Surveyor AB 96 Success despite damaged array
MRO AB 05
Mars Odyssey AB* 01 *Originally planed as AC
Galileo Probe E 89 Highest entry of all time; 60 km/s
Mars Pathfinder E 96 First direct EDL
Mars Exploration Rovers
E 03 Much improved EDL reliability and landed mass ratio
Stardust E 99 Highest speed Earth entry;12.8 km/sGenesis E 01
Phoenix Mars Lander E 07 Active control planned
Aeroassist Terms and DefinitionsEmpirical Trades
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Commonly Used Terms• Entry velocity• Entry flight path angle• Angle of attack• Ballistic coefficient• Reynolds, Mach and Knudsen numbers• Aerodynamic regimes • Aerodynamic coefficients• Aeroshell geometry• Dynamic pressure• Heat rate• Heat load• Terminal velocity
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Velocity and Flight Path Angle
γi Vi
local horizontal
atmospheric interface
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• Inertial velocity, Vi: Vehicle’s velocity wrt inertial coordinate system• Relative velocity, Vr: Vehicle’s atmospheric relative velocity vector• FPA, γi: The angle between the local horizontal (defined perpendicular to the
vehicle radius vector) and the vehicle’s velocity vector. Can be specified as inertial or atmospheric-relative. Defined positive above horizontal.
• By convention, inertial entry conditions are generally specified• Steeper (more negative) γi implies:
– Deceleration lower in the atmosphere (higher peak deceleration and heat rate)– Shorter range and timeline
Mars Entry Trajectory Entry FPA Variations
β = 90 kg/m2
Vi = 5.5 km/s
0
20000
40000
60000
80000
100000
120000
140000
0 1000 2000 3000 4000 5000 6000
Rel Velocity (m/s)
Alti
tude
(m)
10 deg12 deg14 deg
0
100
200
300
400
500
600
700
800
900
1000
0 50 100 150 200 250 300
Time (s)
Rang
e (k
m)
10 deg12 deg14 deg
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Main Forces During Ballistic Blunt Body Hypersonic Entry
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Vr
gravity
local horizontal
drag
Ballistic Coefficientβ = m/(CDA)
• Typically specified in kg/m2 wherem = vehicle massCD = vehicle drag coefficientA = reference area, typically defined by maximum diameter
• Ballistic coefficient is measure of (inertial/aerodynamic) forces• High β implies
– Deceleration, heating, parachute deployment and subsequent events occur lower in the atmosphere
– Longer range and timeline– Higher peak dynamic pressure, heat rate – Lower peak deceleration (small effect)
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Mars Entry Trajectory Ballistic Coefficient Variations
γi = -12 deg Vi = 5.5 km/s
0
20000
40000
60000
80000
100000
120000
140000
0 1000 2000 3000 4000 5000 6000
Rel Velocity (m/s)
Alti
tude
(m) BC=40
BC=65BC=90BC=140
0
100
200
300
400
500
600
700
800
0 50 100 150 200 250
Time (s)
Rang
e (k
m) BC=40
BC=65BC=90BC=140
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Effect of Ballistic Coefficient on Downrange
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Aerodynamic Terms and Definitions
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Reynolds, Mach and Knudsen NumbersMach number: ratio of atmospheric relative velocity to the local speed of sound
M = Vr/aM < 1 subsonic flowM > 1 supersonic flowM > 5 hypersonic flow
Reynolds number: ratio of inertial to viscous fluid dynamic forcesRe = [(ρVL)/µ]∗(φ/L) where φ is theoretical boundary layer thicknessRe/M > 300 turbulent flowRe/M < 300 laminar flow
Knudsen number: ratio of gas’ mean free path to vehicle characteristic lengthKn = λ/dKn > 10 rarefied flowKn < .001 continuum flow
RDB Aug 200562
Vrbody axesα
lift normal
axial
Body and Aerodynamic Axes
αdrag
Body Axes: Specified by designer, generally based on vehicle symmetry. Most flight systems have multiple sets of body axes. Body axes are typically located at center of mass with x axis running through the nose along axis of symmetry. For an axisymmetric body, Y and Z axis stations are arbitrary but must be uniquely prescribed.
Axial force: integrated force in the +X direction, ANormal force: integrated force in the -Z direction, NSide force: integrated force in the +Y direction, Y
Aerodynamic Axes: Specified by atmospheric relative velocity vector (including wind if present). By convention, aero axes are located at center of mass with x axis parallel to atmospheric relative velocity vector.
Drag force: integrated force along the wind axis, DLift force: integrated force perpendicular to the wind axis in the X-Z plane, LSide force: integrated force perpendicular to the wind axis in the X-Y plane, Y
X
-Z
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Vrbody axesα
lift normal
axial
Body and Aerodynamic Axes
αdrag
• Body and Aerodynamic Axes Related By Angle of Velocity Vector Relative to Body
L = Ncosα - AsinαD = Nsinα + Acosα
For blunt bodies at small α, A >> N:
L/D = -Asinα/Acosα = -tan(α)
L = 1/2ρV2CLSref
D = 1/2ρV2CDSref
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45-deg Sphere Cone Drag Coefficient as a Function of Kn and Mach numbers
RDB Aug 200566*Regimes depicted for Earth
Dynamic Pressure and Deceleration• Dynamic pressure, Q = (1/2)ρV2, N/m2
• Specifies aerodynamic environment (drag, stability, deceleration)
• Heatshield spallation above certain values of stagnation point pressure may provide another limit on entry system design– Generally, not a driver
• Deceleration generally specified in Earth g’s• Peak deceleration and angular rates generally coincide with
peak dynamic pressure• As entry FPA (γi) steepens, peak deceleration and dynamic
pressure increase• As ballistic coefficient (β) increases, peak dynamic pressure
increases and peak deceleration decreases (small effect)
RDB Aug 200567
Mars Entry Trajectory Entry FPA Variations
β = 90 kg/m2
Vi = 5.5 km/s
0
5
10
15
20
25
30
35
0 50 100 150 200 250 300
Time (s)
Acce
l (g'
s) 10 deg12 deg14 deg
0
20000
40000
60000
80000
100000
120000
140000
0.00E+00 2.00E+03 4.00E+03 6.00E+03 8.00E+03 1.00E+04 1.20E+04
Dyn Press (Pa)
Alti
tude
(m)
10 deg12 deg14 deg
RDB Aug 200568
Mars Entry Trajectory Ballistic Coefficient Variations
γi = -12 deg Vi = 5.5 km/s
0
20000
40000
60000
80000
100000
120000
140000
0.00E+00 2.00E+03 4.00E+03 6.00E+03 8.00E+03 1.00E+04 1.20E+04
Dyn Press (Pa)Al
titud
e (m
) BC=40BC=65BC=90BC=140
0
5
10
15
20
25
30
0 50 100 150 200 250
Time (s)
Acce
l (g'
s) BC=40BC=65BC=90BC=140
RDB Aug 200569
Entry Configuration Design
RDB Aug 200570
Vehicle Geometry• Entry System or Aeroshell:
Complete entry package typically composed of a forebody (heatshield) and a aftbody (backshell) that generally meet just behind the maximum diameter station
• Forebody: Generally a sphere-cone geometry consisting of a constant angle conical flank (θ1) with a hemispherical nose (rn)
• Aftbody: Generally a single cone angle. Sometimes packaging issues require two cone angles (biconicaftbody), θ2 and θ3. Aftbodyterminates with generally flat backshell interface plate (BIP) for spacecraft mating.
rn
θ1
θ2
θ3
BIP
D
Viking Entry System
RDB Aug 200571
D = 3.54 mrn = 0.88R = 1.56 mθ1 = 70 degθ2 = 40 degθ3 = 62 deg
Affect of Forebody Cone Angle on Entry System Drag and Stability
Cone Angle
Drag
Stability
RDB Aug 200572
Aeroshell Configuration Selection• Choice of forebody cone angle requires a design compromise between
drag, stability and packaging– Blunter cones exhibit more drag per surface area– Sharper cones exhibit better stability characteristics– Angle-of-attack considerations push forebody towards lower cone angles
• Nose bluntness selected largely from a heating rationale– Little effect on drag– Larger nose radius decreases static stability– Larger nose radius decreases stagnation point heat rate
• Shoulder radius is largely determined by local heating effects– Blunting the shoulders decreases local heat rate, drag and stability
• Afterbody geometry selection based largely on supersonic/subsonic flow considerations and other mission requirements– Examples: Mars Microprobe and MSR EEV designs had hemispherical
afterbody to induce hypersonic reorientation in event of backwards entry
RDB Aug 200573
Mars Microprobe and Mars Pathfinder Aeroshell Geometry
Microprobe Design Drivers: Backwards instability, Forwards stability, Low drag Pathfinder Design Drivers: High drag, Viking heritage
RDB Aug 200574
Aerothermodynamics: Terms, Definitions and Empirical Trends
RDB Aug 200575
Entry Heating• The kinetic energy of an entry vehicle is dissipated by
transformation into thermal energy (heat) as the entry system decelerates.
• The magnitude of this thermal energy is so large that if all of this energy were transferred to the entry system it would be severely damaged and likely vaporize
• Only a small fraction of this thermal energy is transferred to the entry system – most of this energy is carried away by the flowfield– The thermal transfer fraction is dependant on vehicle shape, size,
aerodynamic regime and velocity– Near peak heating, 1% to 5% of the total thermal energy is
transferred to the entry system
RDB Aug 200576
Energy Loss Over TimeAssume the following approximation:
E = 1/2mV2 + mgh
Energy (MJ)MER Genesis Galileo Probe
Atmospheric Interface
Parachute Deploy
End
1260 1414 1.07 x 106
105(92%)
84(94%)
1.28 x 105
(88%)0.2
(99.98%)0.9
(99.94%)18
(99.998%)
RDB Aug 200577
Note that:• Water vaporizes at approximately 2.3 MJ/kg• Carbon vaporizes at approximately 60.5 MJ/kg
Heat Rate and Heat Load• Blunt body planetary entry heating is generally comprised of convective and
radiative components• Heat rate is the instantaneous heating at a point on the vehicle, typically
expressed at the stagnation point, W/m2
– Specifies the type of heatshield material appropriate• Stagnation point defined as point where velocity vector intersects forebody• Heat load is the integration of heat rate over the trajectory, J/m2
– Specifies the heatshield thickness• Convective heat rate varies with V3 and (ρ/rn)1/2
• Peak heat rate occurs prior to peak dynamic pressure• Radiation becomes a significant contributor as the entry system diameter
and/or entry speed is increased– Generally neglected for robotic Mars missions– Of significance (order 5-20%) for robotic Earth return missions – Of greater significance (order 30-60%) for robotic missions to Venus or human
return from the Moon or Mars, human entry at Mars– Dominant heat transfer mechanism for Galileo entry probe (99%)
• As entry FPA (γi) steepens, peak heat rate increases and heat load decreases• As ballistic coefficient (β) increases, peak heat rate and heat load increase
RDB Aug 200578
Mars Entry Heating Entry FPA Variations
β = 90 kg/m2
Vi = 5.5 km/s
0
50000
100000
150000
200000
250000
300000
350000
400000
450000
500000
0 50 100 150 200 250 300
Time (s)
Heat
Rat
e (W
/m2)
10 deg12 deg14 deg
0
20000
40000
60000
80000
100000
120000
140000
0 5000000 10000000 15000000 20000000 25000000 30000000
Heat Load (J)
Alti
tude
(m)
10 deg12 deg14 deg
RDB Aug 200579
Mars Entry Heating Ballistic Coefficient Variations
γi = -12 deg
Vi = 5.5 km/s
0
50000
100000
150000
200000
250000
300000
350000
400000
450000
500000
0 50 100 150 200 250
Time (s)
Heat
Rat
e (W
/m2)
BC=40BC=65BC=90BC=140
0
20000
40000
60000
80000
100000
120000
140000
0 5000000 10000000 15000000 20000000 25000000 30000000
Heat Load (J)
Alti
tude
(m) BC=40
BC=65BC=90BC=140
RDB Aug 200580
Effect of Turbulence Increases with Entry System SizeNASA Ames Result for Mobile Science Laboratory
RDB Aug 200581
Heating Experience
Peak Heat Rate of Past and Planned Entry Systems
Approximate Entry Heating Design Margins for the Stardust Entry System
RDB Aug 200582*Figures from, “Computational Aerothermodynamic Design Issues for Hypersonic Vehicles,” Gnoffo, et. al., AIAA 97-2473.
Back of the Envelope Calculations
RDB Aug 200583
Approximate Relations
• Entry Velocity from Vinf• Modified Newtonian aerodynamics - drag coefficient• Equations of motion for ballistic entry• Equations of motion for lifting entry• Terminal descent• Heating• Ballistic entry landing accuracy
RDB Aug 200584
Inertial Entry Velocity
• Energy equation: E = v2/2 – µ/r• At SOI, E = Vinf
2/2• At atmospheric interface, E = Vatm
2/2 – µ/ratm
• 2µ/ratm = 24.6 (km/s)2 at Mars, 122.6 (km/s)2 at Earth• Vatm > 4.96 km/sec (Mars)• Vatm > 11.1 km/sec (Earth)
Vatm = sqrt[Vinf2 + 2µ/ratm]
RDB Aug 200585
Newtonian Aerodynamics• 3 centuries ago, Newton postulated a physical
model to describe fluid flow over a body.– Fluid is assumed to have low density so that interactions
among the particles is neglected. This has led to the term impact aerodynamics as the presence of the body is not transmitted upstream.
– After impinging on the body, the normal momentum of the particles is entirely lost.
– After impinging on the body, the tangential momentum of the particles is entirely conserved.
• In general, and for the application of interest to Newton, this theory has been shown to be inaccurate– With the exception of the hypersonic aerodynamics
RDB Aug 200586
RDB Aug 200587
Newtonian (Impact) Aerodynamics
A
VcosδV δ
A
Normal velocity = VsinδTangential velocity = Vcosδ
Normal velocity = 0Tangential velocity = Vcosδ
Change in normal velocity = VsinδMass flux incident on surface A = ρVAsinδd/dt(mV) = (ρVAsinδ)(Vsinδ-0) = ρV2Asin2δ
F = d/dt(mV) = ρV2Asin2δ
F/A = P – Pinf = ρV2sin2δ
(P – Pinf)/(1/2)ρV2 = 2sin2δ
Cp = 2 sin2δ
Bottom line: For impact flows, pressure coefficient is solely a function of vehicle geometry, δ. Not a function of Mach, Re or ρ. This theory is generally applicable to continuum hypersonic flow.
Newtonian Aerodynamics for a Cone
RDB Aug 200588
z
xδ
CP
CPcosδ
CN = cos2δsin2α
CA = 2sin2δcos2α + cos2δsin2α
For α = 0 deg,
CN = CL = 0
CA = CD = 2sin2δ
α
V
CN
CA
CPsinδ
Modified Newtonian Flow
Lees proposed the following modification to Newtonian theory:
CP = Cpmaxsin2δ
Where Cpmax is a function of γ, the ratio of specific heats
Cpmax ~ (γ+3)/(γ+1)
For Earth, γ = 1.4 and Cpmax = 1.833For Mars, γ = 1.3 and Cpmax = 1.869
Note that Cpmax 2 as γ 1
RDB Aug 200589
Modified Newtonian Drag Coefficient
• CD = 1.869sin2δ (γ = 1.3)– δ = 20 deg, α = 0, CD = 0.22– δ = 30 deg, α = 0, CD = 0.47 – δ = 45 deg, α = 0, CD = 0.93 – δ = 60 deg, α = 0, CD = 1.40– δ = 70 deg, α = 0, CD = 1.65
RDB Aug 200590
Main Forces During Hypersonic Entry
RDB Aug 200591
Vrlocal horizontal
gravity
lift
γr
Drag force opposes the vehicle’s atmospheric velocity vector
Lift acts in a direction normal to the atmospheric velocity vector
Gravitational force is directed towards the central body along the radius vector
Centrifugal force is directed away from the central body along the radius vector
centrifugal forcedrag
Newtonian Equations of Motion• For a point mass in an atmosphere, with the origin
of inertial space located at the center of mass of the planet in atmospheric axes:
-D + mgsinγ - m(V2/r)sinγ = mdV/dt (1)-L + mgcosγ - m(V2/r)cosγ = mVdγ/dt (2)
Two kinematic relations can also be derived:dr/dt = dh/dt = Vsinγ (3)ds/dt = rdφ/dt = Vcosγ (4)
where,s: distance over the planetary surface (downrange)dφ/dt: angular velocity of the radius vector, r, wrt inertial space
RDB Aug 200592
Simplifying Assumptions for Ballistic EntryAllen and Eggers solution:
Ballistic entry, L = 0Concerned with region of flight where γ = constant
From (2): mg = mV2/r(1) becomes: -D = mdV/dt
dV/dt = -ρV2CDS/(2m)dV/dt = -ρV2/(2β)
dV/V2 = -[ρ/(2β)]dtFrom (3): dt = dh/Vsinγ
dV/V = -[1/(2βsinγ)]ρdh (5)
RDB Aug 200593
Simplifying Assumptions for Ballistic EntryAssuming exponential atmospheric density, where:
ρ = ρ0e-xh (6)
1/x is commonly referred to the scale height
Planet “Sea-level” density, ρ0
(kg/m3)
x(m-1)
ρ0/x (kg/m2)
Venus 16.02 1.606 x 10-4
1.378 x 10-4
1.275 x 10-4
9.975 x 104
Earth 1.226 8.897 x 103
Mars 0.057 4.471 x 102
RDB Aug 200594
Simplifying Assumptions for Ballistic EntryCombining (5) and (6) yields:
dV/V = -[ρ0/(2βsinγ)] e-xh dhIntegrating from the atmospheric interface to
any other altitude, h, yields:ln(V/Vatm) = [ρ0/(2βxsinγ)] (e-xh − e-xhatm)
V/Vatm = exp{[ρ0/(2βxsinγ)] (e-xh − e-xhatm)}At the atmospheric interface, ρ = ρ0e-xhatm = 0So,
V = Vatm exp{[ρ0/(2βxsinγ)]e-xh}V = Vatm exp{Ce-xh} (7)
RDB Aug 200595
Assumptions and Region of Validity
• Allen and Eggers assumptions:– No lift– Constant γ– Constant β– Exponential atmospheric density
• These conditions generally hold in the region of peak deceleration and heating
RDB Aug 200596
Altitude
Relative Velocity
Region of validity
Constant, CC = ρ0/(2βxsinγ)
• C is dimensionless and negative (since γ is negative and all other terms are positive)
• The planet specifies ρ0 and x– Decreasing ρ0/x decreases C, thereby increasing the
value of V at a specified altitude (Recall Venus-Earth-Mars comparison earlier).
• The magnitude of C is inversely proportional to both β and γ– Increasing β or γ decreases C, thereby increasing the
value of V at a specified altitude (same as result empirically derived earlier)
RDB Aug 200597
Deceleration
RDB Aug 200598
Differentiation (7) wrt time yields,dV/dt = Vatm Cexp{Ce-xh}(-xe-xh)dh/dt (8)
Where from (3),dh/dt = Vsinγ = Vatmsinγ(exp{Ce-xh})
Such that (8) becomes,dV/dt = -xCVatm
2sinγexp{2Ce-xh}(e-xh)
• Both C and sinγ are negative, so dV/dt will be negative (the vehicle is decelerating)
• Deceleration is generally provided in Earth g’s, n = (dV/dt)/9.806
(9)
Altitude of Peak Deceleration• To determine the magnitude of the maximum
deceleration and the altitude at which it occurs, d/dh[dV/dt] = 0
• Which from (9) becomes:x2CVatm
2sinγexp{2Ce-xh}(e-xh)(2Ce-xh+1) = 0(2Ce-xh+1) = 0
e-xh = -1/2Chn,max = (1/x)ln(-2C)
• From (10) we see that the altitude of peak deceleration is independent of entry velocity, and for a given planetary atmosphere is only dependant on β and γ
(10)(11)
RDB Aug 200599
Magnitude of Peak Deceleration• Substituting (10) into (9) yields the following
expression for nmax
nmax = (dV/dt|max)/9.806nmax = xVatm
2sinγ/(2*9.806*e)nmax = xVatm
2sinγ/(53.3)• From (12) we see that the peak deceleration
magnitude is independent of β and, for a given planetary atmosphere, is purely a function of the entry angle and square of the entry velocity (recall earlier empirical results showing weak dependence on β)
(12)
RDB Aug 2005100
Velocity of Peak Deceleration• Substituting (10) into (7) yields the following
expression for velocity of nmax
V = Vatm exp{Ce-xh}Vn, max = Vatm e-1/2
Vn, max = 0.606Vatm
• Which shows that at nmax the velocity is a function only of entry velocity.
(13)
RDB Aug 2005101
Range• The approximate distance traversed during the entry
can be obtained by combining (3) and (4) as follows:ds = Vcosγdt = Vcosγ(dh/Vsinγ)
ds = cotγdh∆s = (cotγ)∆h (14)
RDB Aug 2005102
Practical Considerations
• In practice, selection of the atmospheric density model (scale height and ρ0) and γrequire care.
• Note that these relations were derived assuming an exponential density profile such that a closed-form solution could be obtained. One could use any other analytic expression for density or simply integrate.
RDB Aug 2005103
Assumptions and Region of Validityβ = 90 kg/m2, γ = -12 deg
-25
-20
-15
-10
-5
00 1000 2000 3000 4000 5000 6000
Atmospheric Relative Velocity (m/s)
Atm
osph
eric
Rel
ativ
e FP
A (d
eg)
Series1
0
20000
40000
60000
80000
100000
120000
140000
0 2000 4000 6000
At mosphe r i c Re l a t i v e Ve l oc i t y ( m/ s)
What γ should be used?RDB Aug 2005
104
Numerical Example• A 70-deg sphere cone with 2.65 m diameter is used
to perform a direct entry at Mars. This system enters the Mars atmosphere with a mass of 830 kg, an atmospheric-relative entry velocity of 5.45 km/s and an atmospheric-relative entry flight path angle of -5.0 deg in the region of peak deceleration. Assume the COSPAR Mars atmospheric model. Calculate the ballistic coefficient, then determine the altitude and velocity of peak deceleration as well as the peak deceleration magnitude. Approximate the parachute deployment altitude and downrange distance, knowing that chute deploy occurs at Mach 1.8.
RDB Aug 2005105
Numerical Example• Ballistic coefficient, β = m/(CDS)
CD = 1.869sin2(70 deg) = 1.65β = (830)/(1.65*pi*2.65*2.65/4) = 91.2 kg/m2
• For the COSPAR Mars atmospheric model, x = 1.275 x 10-4 m-1
ρ0 = 5.70 x 10-2 kg/m3
• C=ρ0/(2βxsinγ) = -28.12
• hn,max = (1/x)ln(-2C) = 31.61 km • Vn, max = 0.606Vatm = 0.606(5.45) = 3.30 km/s• nmax = xVatm
2sinγ/(53.3) = 6.19 g’s
RDB Aug 2005106
Numerical Example• Parachute deploy Mach number = 1.8• Parachute deploy velocity = 1.8*(222) = 400 m/s• V = Vatm exp{Ce-xh}• 400 = 5450exp{-28.12exp-(1.275 x 10-4)h}• Chute deploy altitude = 18.6 km
• For an atmospheric interface altitude of 125 km, ∆h = 125 – 18.6 = 106.4 km
• ∆s = 106.4*cot(5.0) = 1216 km
RDB Aug 2005107
Comparison with Numerical Integration
-2.50E+01
-2.00E+01
-1.50E+01
-1.00E+01
-5.00E+00
0.00E+000.00E+00 2.00E+03 4.00E+03 6.00E+03
Velocity (m/s)
FPA
(de
Series1
• 3D POST used to numerically integrate equations of motion• Spherical planet• COSPAR exponential density model• -11 deg FPA at atmospheric interface
RDB Aug 2005108
Comparison with Numerical Integration
0.00E+002.00E+044.00E+046.00E+048.00E+041.00E+051.20E+051.40E+05
0.00E+00
1.00E+03
2.00E+03
3.00E+03
4.00E+03
5.00E+03
6.00E+03
Velocity (m/s)
Alti
tude
(
RDB Aug 2005109
Numerical ExampleParameter Approximate
Solution3-DOF
numerical integration
% error
Peak deceleration, Earth g’s
6.19 6.28
32.65
3.50
Chute deploy downrange (km)
1216 848 43%
13.4
1%
Altitude of peak deceleration (km)
31.61 3%
Velocity of peak deceleration (km/sec)
3.30 6%
Chute deploy altitude (km)
18.6 39%
RDB Aug 2005110
Numerical Example
RDB Aug 2005111
• Clearly, this approximation works well in predicting peak deceleration conditions.
• We will see later that since the peak heat rate occurs in this constant-FPA regime, this approximation provides an accurate estimate of these conditions as well.
• By the time of parachute deployment, we are long past the point of constant flight-path angle. As such, we should not expect this approximation to produce accurate results. Since the flight-path angle is increasing rapidly in this region, this approximation will always predict a parachute deployment condition of longer range and higher altitude than reality.
MER Numerical Example
• The MER entry system is a 70-deg sphere cone with 2.65 m diameter. This system enters the Mars atmosphere* with a mass of 830 kg, an atmospheric-relative entry velocity of 5.45 km/s and an atmospheric-relative entry flight path angle of -11.0 deg. Calculate the ballistic coefficient, then determine the altitude and velocity of peak deceleration as well as the peak deceleration magnitude. Approximate the parachute deployment altitude and downrange distance, knowing that chute deploy occurs at Mach 1.8.
*Atmospheric interface is defined at 125 km altitude
RDB Aug 2005112
MER Numerical Example
RDB Aug 2005113
• Ballistic coefficient, β = m/(CDS)CD = 1.869sin2(70 deg) = 1.65β = (830)/(1.65*pi*2.65*2.65/4) = 91.2 kg/m2
• For Mars, x = 1.275 x 10-4 m-1
ρ0 = 5.70 x 10-2 kg/m3
• C=ρ0/(2βxsinγ) = -30.67• Using the same flight-path angle ratio as earlier, I
estimate the constant-altitude FPA to be:-11*(5/12) = -4.583 deg
• hn,max = (1/x)ln(-2C) = 32.3 km • Vn, max = 0.606Vatm = 0.606(5.45) = 3.30 km/s• nmax = xVatm
2sinγ/(53.3) = 5.68 g’s
MER Numerical Example• Parachute deploy Mach number = 1.8• Parachute deploy velocity = 1.8*(222) = 400 m/s• V = Vatm exp{Ce-xh}• 400 = 5450exp{-30.674exp-(1.280 x 10-4)h}• Chute deploy altitude = 19.25 km
• For an atmospheric interface altitude of 125 km, ∆h = 125 – 19.25 = 105.75 km
• ∆s = 120.42 cot(4.583) = 1319.25 km
RDB Aug 2005114
MER Numerical ExampleParameter Approximate
SolutionMER 6-DOF prediction1
% error
Peak deceleration, Earth g’s
5.68 5.93
32.9
3.6
Chute deploy downrange (km)
1319 780 70%
8.66
4%
Altitude of peak deceleration (km)
32.3 2%
Velocity of peak deceleration (km/sec)
3.3 8%
Chute deploy altitude (km)
19.25 122%
1Desai, P.N.; and Lee, W.J.,” Entry, Descent and Landing Scenario for the Mars Exploration Rover Mission,” October 2003, Lisbon, Portugal.
RDB Aug 2005115
Note that, in this case, good agreement of the peak deceleration conditions is obtained eventhough the MER 6-DOF prediction does not rely on the COSPAR exponential atmospheric density model.
Reasons to Fly a Lifting Entry
• Lift can be judiciously used to:– Increase the allowable approach navigation
uncertainty– Reduce the deceleration environment– Mitigate atmospheric density and wind uncertainty– Enable higher surface elevation landing sites or
more mass to the surface (e.g., slide 23) – Improve the landed accuracy (e.g., slide 24)– Execute a plane change without propulsion
RDB Aug 2005116
Simplifying Assumptions for Lifting EntryEquilibrium glide: Relatively shallow glide in
which gravitational force is balanced by the combination of lift and centrifugal forces.
Assumptions:– γ is small, such that sinγ = 0 and cosγ = 1– γ is changing slowly, such that dγ/dt = 0– L/D > 0.5– Lift is in the orbital plane, vertical direction
For this case, equations (1) and (2) become:-D/m = dV/dt (15)-L/m + g = V2/r (16)
RDB Aug 2005117
Simplifying Assumptions for Lifting Entry• Substituting for L and D,
-ρV2CDA/2m = dV/dt (17)-ρV2CLA/2m + g = V2/r (18)
• Equation (18) can be rewritten as:g = V2{1/r + ρCLA/2m}
g = V2{1/r + ρ(L/D)CDA/2m}gr = V2{1 + ρ(L/D)r/2β}
V = sqrt[gr/{1 + (ρ0e-xh)(L/D)r/2β}] (19)• Increasing L/D or decreasing β shifts the
deceleration higher in altitude
RDB Aug 2005118
Deceleration• From equation (15),
n = (1/g)dV/dt = -D/mg = -(L/m)/{(L/D)g}• From equation (16),
L/m = g - V2/r n = -(g - V2/r)/{(L/D)g}n = -(1 - V2/rg)/(L/D)
• Where V2 can be obtained from (19). Since V2
continuously decreases over the entry trajectory, n continuously increases, reaching a maximum of,
nmax = -1/(L/D)
(20)
(21)
RDB Aug 2005119
Deceleration• From equation (15),
n = (1/g)dV/dt = -D/mg = -(L/m)/{(L/D)g}• From equation (16),
L/m = g - V2/r n = -(g - V2/r)/{(L/D)g}n = -(1 - V2/rg)/(L/D)
• Where V2 can be obtained from (19). Since V2
continuously decreases over the entry trajectory, n continuously increases, reaching a maximum of,
nmax = -1/(L/D)
(20)
(21)
RDB Aug 2005120
Deceleration
• From equation (21), we see that for an equilibrium glide trajectory.
– Peak deceleration is independent of entry velocity, the ballistic coefficient or the planetary atmosphere
– A small amount of lift can significantly reduce the peak deceleration
RDB Aug 2005121
Time to Landing
• The relationship between time and velocity for an equilibrium glide entry is obtained as follows:
dt = -[(L/D)/(L/m)]dVdt = -(L/D)/[g(1 - V2/rg)]dV
• Integrating from the present time to the time of landing (where V = 0) yields,
∆t = 0.5(r/g)1/2(L/D)ln{(1+V2/rg)/(1-V2/rg)}• Equation (22) demonstrates that the time in the
entry trajectory is proportional to L/D ratio and entry velocity
(22)
RDB Aug 2005122
Downrange• With the small angle assumption associated with
equilibrium glide, the range can be obtained as: ds = Vdt = -(L/D)V/[g(1 - V2/rg)]dV
• Integrating from the present time to the time of landing (where V = 0) yields,
∆s = -r/2(L/D)ln[1-(V2/rg)]• As with time of flight, downrange is proportional to L/D and
entry velocity• The downrange achievable with a lifting vehicle is
significantly greater than that achievable in ballistic flight• Modulating vertical L/D directly controls downrange (range
to the target) – hence, the use of bank-angle modulation in precision landing
(23)
RDB Aug 2005123
Crossrange
RDB Aug 2005124
• An additional benefit of lift is the ability to make turns within the atmosphere and reach landing sites not in the orbital (vertical) plane.
• This out-of-plane distance is termed crossrange.• From the lateral equations of motion, it can be shown that
maximum crossrange is:Ymax = (r/5.2)(L/D)2[1/sqrt(1+0.106(L/D)2)]
• For global access, Ymax/r = pi/2 and L/D = 3.5
L/D Ymax/r0.5 0.0471.0 0.1832.0 0.6983.0 1.2383.5 1.554
(24)
Shuttle Numerical Example• A Shuttle-like vehicle is entering the Earth’s
atmosphere using an equilibrium glide trajectory. It has a velocity of 7.5 km/sec at an altitude of 85 km. β = 500 kg/m2 and the L/D is 2.0. What is the velocity and deceleration at 60 km? What is the peak deceleration, downrange and crossrange capability? What is the time of flight?
• At 60 km on Earth, r = 6378+60 km = 6.438e+06 mρ = 3.1459e-04 kg/m3
• V = sqrt[gr/{1 + (ρ0e-xh)(L/D)r/2β}]• V = 3535 m/s
RDB Aug 2005125
Shuttle Numerical Example
• n = -(1 - V2/rg)/(L/D)• n = 0.4 Earth g’s• nmax = -1/(L/D) = 0.5 Earth g’s• Downrange capability = 14269 km• Time of flight = 38.5 min• Crossrange capability = 4451 km
RDB Aug 2005126
Shuttle Numerical Example
0
20
40
60
80
100
0 2000 4000 6000 8000
Velocity (m/s)
Alti
tude
(km
)
Series1
RDB Aug 2005127
00.10.20.30.40.50.6
0 2000 4000 6000 8000
Velocity (m/s)
Dec
eler
atio
n (E
arth
g's
)
Series1
Terminal Velocity• Consider a body in a planetary atmosphere under the
influence of two forces, gravity and drag. If the velocity vector aligns with the radius vector:
ΣF = maD – mg = 0
1/2ρV2CDS – mg = 0V = srqt{(2mg)/(ρCDS)}
For a fixed configuration with known aerodynamics,
V = constant/sqrt{ρ}
For a fixed mass with unknown (but constant) aerodynamics,
∆V = constant/sqrt{∆ρ} orV – Vref = constant/sqrt{ρ – ρref}
Drag
mg
RDB Aug 2005128
With radar altimeter data to measure V directly, ρ or dρ/dh may be calculated
Terminal Descent Numerical Example• During descent, the MER radar altimeter
measured an altitude time history approximated by the following expression:
h = 114.36-70.298*t+0.09256*t2
• Can the near-surface density be determined?
RDB Aug 2005129
Terminal Descent Numerical Example• The density can not be determined from this
data without either a single density measurement or assuming the aerodynamic of the parachute configuration are known precisely. However, the rate of change of density (or scale height) can be determined as follows:
• V = hdot = -70.298+2*(0.09256)t• Vref = V(t=0) and ρref = ρ(t=0)• ρ/ρref = (Vref*Vref)/(V*V)
RDB Aug 2005130
Terminal Descent Numerical Example
0
1000
2000
3000
1.1500E-02
1.2000E-02
1.2500E-02
1.3000E-02
1.3500E-02
1.4000E-02
1.4500E-02
Density (kg/m3)
Alti
tude
(m)
Series1
Assuming density at surface = 1.415e-02 kg/m3
RDB Aug 2005131
Gravity TurnInitially developed for the lunar Surveyor landings, the “gravity turn” control law aligns the thrust vector opposite to the velocity direction. The presence of a gravitational force tends to orient the flight path toward vertical over time.
Equations of Motion:
dv = -aT + g cos ψdt
v dψ = -g sin ψdt
ψ
T
mg v
RDB Aug 2005132
aT = acceleration due to thrustv = inertial velocityg = gravitational accelerationψ = off-nadir anglem = massT = thrust
Gravity Turn (cont.)
RDB Aug 2005133
Assuming a small off-nadir angle, these equations of motion may be solved yielding the following quadratic closed-form solution for a constant acceleration due to thrust, given the initial and final state constraints.
(aT/g)2 - [2(v2 - vt2)/(4g*(h-ht))] (aT/g) - [1+(2v2 - 2vt
2)/(4g*(h-ht))] = 0
where: aT = T / m = acceleration due to thrustg = gravitational accelerationh = instantaneous altitude ht = final altitudev = instantaneous velocity vt = final velocity
Note: when solving quadratic, take positive root!
This equation can be solved iteratively at each time step during unpowereddescent. As altitude decreases, the required thrust level (T) will increase. When the required thrust level reaches the flight system’s design thrust level, the gravity turn is initiated. Upon gravity turn completion, altitude and velocity will equal ht, vt.
In flight missions, the gravity turn segment is generally followed by a constant velocity vertical descent segment until h = 0.
Gravity Turn (cont.)To compute the propellant mass expended during the gravity turn, we can approximate the ∆V accrued over each time step as:
∆V = aT * ∆T
Propellant mass is then calculated via the rocket equation:
∆mprop = mi [1 - (1/exp(∆V/gEISP))]
Where ∆mprop = propellant mass expendedmi = initial mass∆V = change in velocitygE = gravitational acceleration at Earth (9.8 m/s2)ISP = specific impulse of descent propulsion system
RDB Aug 2005134
Gravity Turn Numerical ExampleFollowing parachute descent, the Phoenix lander (wet mass = 380 kg) will perform terminal descent using its twelve 289 N (65 lbf) thrusters at a 75% duty cycle, for a total thrust level of 2600 N. Given the following altitude/velocity history as computed using radar altimeter data, with target values of h=12 m, v = 2.4 m/s at the completion of the gravity turn, at what altitude does thrusting begin to start the gravity turn? Assume g = 3.7 m/s2.
t h v(s) (m) (m/s)0 1500 781 1422 77.92 1344.1 77.83 1266.3 77.74 1188.6 77.65 1111 77.56 1033.5 77.47 956.1 77.38 878.8 77.29 801.6 77.110 724.5 7711 647.5 76.912 570.6 76.813 493.8 76.714 417.1 76.615 340.5 76.516 264 76.417 187.6 76.318 111.3 76.219 35.1 76.1
RDB Aug 2005135
Gravity Turn Numerical Example (Solution)At each time step, we solve the quadratic for aT/g. Then, for each timestep we can calculate aT and required thrust, T. As expected, the required thrust increases as altitude decreases. When the required thrust equals the thrust capability of our system (2600 N), the gravity turn is initiated. In this example, descent engine turn-on begins at an altitude of about 966 m (interpolated).
t h v aT/g aT T(s) (m) (m/s) (m/s^2) (N)0 1500 78 1.55 5.75 2185.621 1422 77.9 1.58 5.86 2226.452 1344.1 77.8 1.61 5.98 2272.013 1266.3 77.7 1.65 6.11 2323.154 1188.6 77.6 1.69 6.27 2380.985 1111 77.5 1.74 6.44 2446.896 1033.5 77.4 1.79 6.64 2522.717 956.1 77.3 1.85 6.87 2610.878 878.8 77.2 1.93 7.14 2714.629 801.6 77.1 2.01 7.47 2838.5110 724.5 77 2.12 7.87 2989.0311 647.5 76.9 2.25 8.36 3175.8112 570.6 76.8 2.42 8.98 3413.7413 493.8 76.7 2.64 9.81 3727.1714 417.1 76.6 2.95 10.94 4158.8015 340.5 76.5 3.40 12.61 4791.0316 264 76.4 4.12 15.28 5806.0417 187.6 76.3 5.46 20.27 7702.3618 111.3 76.2 8.87 32.92 12508.4819 35.1 76.1 34.76 128.94 48995.45
RDB Aug 2005136
Convective Stagnation-Point Heat Rate• General expression for stagnation-point heating (convective or
radiative) is:q ~ VN∗ (ρ)Μ∗(rn)R
• Where,rn is the vehicle nose radiusV is the atmospheric relative velocity magnitudeρ is the atmospheric densityq is the convective stagnation-point heat rate
• Several useful approximations have been developed to estimate the convective stagnation-point heat rate (e.g., Chapman’s equation, Sutton-Graves equation, and Fay-Riddell equation)
• These equations denote the following general dependency:
q ~ V3*(ρ/rn)1/2
N ~ 3, M ~ 0.5, R ~ -0.5
• Stagnation-point convective heat rate increases with flight speed and density and decreasing nose radius
RDB Aug 2005137
Radiative Stagnation-Point Heat Rate• General expression for stagnation-point heating (convective or
radiative) is:q ~ VN∗(ρ)Μ∗(rn)R
• Where,rn is the vehicle nose radiusV is the atmospheric relative velocity magnitudeρ is the atmospheric densityq is the convective stagnation-point heat rate
• Several useful approximations have been developed to estimate the radiative stagnation-point heat rate (e.g., Tauber-Sutton equation)
• This equation denotes the following general dependency (for Earth):
q ~ V8.5∗(ρ)1.6∗(rn)
N ~ 8.5, M ~ 1.6, R ~ 1.0
• Stagnation-point radiative heat rate increases rapidly with flight speed and with density and nose radius – (clear implications for human return entry systems from the Moon or Mars)
RDB Aug 2005138
Approximate Stagnation-Point Convective Heat Rate Equations
• Chapman’s equation (for Earth) can be expressed as:
q = 1.83e-04(ρ/rn)1/2V3(1-hw/ht)
– hw/ht is the ratio of wall enthalpy to total flow enthalpy, which is typically < 0.1
• Sutton-Graves equation can be expressed as:
q = k(ρ/rn)1/2V3
(25a)
(25b)
RDB Aug 2005139
Planet kEarth 1.74153e-04Mars 1.90270e-04
Entry Heating (Ballistic Entry)• For a ballistic entry, it can be shown that
qmax ~ V3(βsinγ)1/2
Q ~ V2(β/sinγ)1/2
where Vq,max = e-1/6Vatm = 0.846Vatm
• Since Vn,max occurs at 0.606Vatm, we see that peak heating occurs earlier (at higher velocity and altitude) than peak deceleration. It can be shown that hq, max > hn, max
• From these relations, we see that increasing β or the entry velocity, increases the peak heat rate and heat load; whereas increasing the entry angle decreases the total heat load but increases the peak heat rate (recall earlier empirical results)
(26)
RDB Aug 2005140
Entry Heating (Lifting Entry)• These relations are modified for lifting entry as follows:
qmax ~ V3[βsinγ/(L/D)]1/2
Q ~ V2[(L/D)β/sinγ]1/2
where
Vq,max = 0.816Vatm
• From these relations, we see that increasing L/D, decreases the peak heat rate and increases the total heat load
(26)
RDB Aug 2005141
Radiative Equilibrium Temperature• A body radiates heat at a rate proportional to the 4th
power of its temperature• Stephen-Boltzman law:
qrad = σεTw4
qrad = 5.67ε(Tw/1000 K)4 W/cm2
Where ε = emissivity, ε = 1 for blackbody
In equilibrium, qin = qout
qin = 5.67ε(Tw/1000 K)4 W/cm2
Can solve for Tw
(27)
RDB Aug 2005142
Shuttle Numerical Example
RDB Aug 2005143
• For the Shuttle-like entry previously studied, estimate the stagnation-point convective heat rate and temperature at 60 km altitude. Assume a 1 m nose radius and an emissivity of 0.8. Estimate these quantities at the peak heating condition.
• At 60-km altitude,ρ = 3.1459e-04 kg/m3
V = 3535 m/sq = 1.83e-04(3.1459e-04)1/2(3535)3(1) = 14.3 W/cm2 (Chapman’s eq)q = 1.74153e-04(3.1459e-04)1/2(3535)3 = 13.6 W/cm2 (Sutton-Graves eq)
14.3 = 5.67(0.8)(Tw/1000 K)4
Tw = 1332 K = 1938 F
• At the peak heating condition, Vqmax = 0.816*7.5 = 6.12 km/sech = 72.875 km (from solution to eq 19)ρ = 5.3362e-05 kg/m3
q = 1.83e-04(5.3362e-05)1/2(6120)3(1) = 30.6 W/cm2
Tw = 1612 K = 2442 F
MER Numerical Example• What is the peak stagnation-point heat rate and
temperature for the MER example previously examined? The MER entry system has a nose radius = 0.5 base radius.
• At the peak heating condition, Vqmax = 0.846*5.45 = 4.61 km/secrn = (2.65/2)/2 = 0.6625 mh = 40.87 km (from solution of eq 7)ρ = 3.110e-04 kg/m3
q = 1.9027e-04(3.110e-04 /0.6625)1/2(4610)3 = 40.4 W/cm2
Tw = 1727 K = 2616 F (assuming an emissivity of 0.8)
39.9 W/cm2 (a 1.2% difference) is cited in Desai, P.N.; and Lee, W.J., ”Entry, Descent and Landing Scenario for the Mars Exploration Rover Mission,” October 2003, Lisbon, Portugal.
RDB Aug 2005144
Ballistic Entry Landed Accuracy• For Mars ballistic entry, the major uncertainties that affect
landing accuracy are entry flight path angle uncertainty and atmospheric density
• For Mars entry flight path angle uncertainties greater than 0.1 deg, this variable dominates and the following approximation can be made
a = 150*∆γ, km wherea: 3-σ semi-major axis of landed footprint∆γ: one-sided FPA uncertainty
• This approximation breaks down gradually as the target γapproaches skipout or becomes excessively steep
• For entry γ uncertainties less than 0.1 deg, the landed dispersion is largely determined by the atmospheric density uncertainty assumed. Unfortunately, there is little data to substantiate these assumptions.
RDB Aug 2005145
Key Aeroassist Technologies
RDB Aug 2005146
Key Technologies
• Approach Navigation• Thermal Protection System• Deployable Systems• Atmospheric GN&C • Terminal Descent System• Landing Systems
RDB Aug 2005147
Approach Navigation• Sets initial conditions: single most important
driver for aeroassist performance• For ballistic entries at Mars, end-to-end
landing ellipse major axis is approx 300 km per deg of γ uncertainty, for γ errors > 0.1 deg
• ∆DOR and optical navigation provide a significant impact on landed accuracy and also reduce the γ dependence on latitude (next slide)
RDB Aug 2005148
Navigation Data Types• During interplanetary cruise, the JPL navigation team uses several different
techniques to track the spacecraft’s position and speed through the DSN.– Doppler and ranging are the two most common techniques– Mars Odyssey and MER utilized ∆DOR to improve arrival nav accuracy– MRO plans to use optical nav as another means of improving arrival nav accuracy
• In ranging, a signal is sent from Earth to the s/c and the s/c sends a signal back to Earth. By measuring precisely how long the signal takes to make the round trip at the speed of light, the spacecraft’s distance from Earth along the line of sight can be determined.
• In two-way Doppler tracking, a ground station sends a signal to the s/c and the s/c sends a signal back to Earth. By looking for small changes in the frequency of the spacecraft’s signal, the s/c velocity along the Earth line of sight can be determined.
– The signal’s frequency changes with the spacecraft’s speed, much like the rising and falling of the siren of a fire truck or train as it passes by.
• For the Odyssey and MER missions, an additional technique, “delta differential one-way range,” or ∆DOR was employed. In this technique, two different ground stations on Earth simultaneously measure signals from the s/c and from one of several distant quasars in space. Like beacons in the cosmos, quasars provide very stable radio signals. By combining the measured signals using interferometry (VLBI), navigators measure the s/c’s angular motion relative to Earth. These measurements provide insight into the “plane of sky” s/c motion.
RDB Aug 2005149
Ranging
• Spacecraft range is measured by the round-trip transit time of a ranging signal generated at one of the DSN stations, to the spacecraft, and returned to Earth.
– A ranging signal consists of a sequence of sinusoidal tones phase-modulated onto a carrier signal.
– The spacecraft receiver locks on to the ranging signal and turnsaround a downlink signal.
– The received downlink signal at the DSN is demodulated, and the received “range code” is compared with the uplinked range code to compute round-trip transit time. The round-trip transit time, τ, can be divided by two times the speed of light, c, to find the one-way slant range, ρ:
ρ = τ / 2c
– In practice, τ is computed from t (the measured roundtrip time) by subtracting a known and calibrated ε (spacecraft process time)
RDB Aug 2005150
• An expression for the received frequency of a signal sent from areceding spacecraft to Earth is:
fR = (1 - ρ / c) fT
where fT is the frequency transmitted by the spacecraft and ρ is the spacecraft instantaneous slant range rate.– The quantity (ρ / c) fT is referred to as the Doppler shift.– The Doppler measurement provides information on the spacecraft
slant range rate.
Doppler
.
.
.
RDB Aug 2005151
Doppler (cont.)• For a receding spacecraft, the slant range rate calculated by the
equation on the previous page is a sinusoid superimposed upon a ramp function representing the spacecraft geocentric velocity.
– The diurnal sinusoid behavior is the result of the rotation of the tracking station about the Earth’s spin axis.
– The amplitude and phase of this sinusoid provide information about the spacecraft declination and right ascension.
– From a single pass of Doppler data, it is possible to determine the spacecraft radial velocity, right ascension and declination.
– Velocities normal to the line of sight must be inferred from several days of Doppler data.
RDB Aug 2005152
Angular Measurements: VLBI Example• For most interplanetary missions, spacecraft position uncertainty is much
smaller in the Earth-to-spacecraft “radial” direction than in the perpendicular “plane-of-sky” direction.
– Radial components of position and velocity are directly measured by range and Doppler observations.
– Plane-of-sky errors are more than 1000 x radial errors, even under the most favorable conditions, using only ranging and Doppler.
• In general, angular measurements can be made using multiple ground stations to simultaneously receive spacecraft transmissions during DSN view period overlaps.
– From an accurately known baseline, B, and a calculated delta slant range distance, ρ2 - ρ1, one can compute the spacecraft declination, δ. (VLBI)
RDB Aug 2005153
∆DOR• Delta Differential One-Way Range (∆DOR) is a VLBI measurement
technique that utilizes two ground stations to simultaneously view the spacecraft and a known radio source (quasar or another spacecraft) to provide an angular position determination.
– Two stations receiving the same ranging signal allows a geometric plane-of-sky angular position measurement (Differential), as shown on the previous slide.
– By receiving signals from two sources, common errors can be canceled out, allowing a precise measurement of the angular separation of the two radio sources (Delta).
– Since the plane-of-sky angular position of the quasar is well known, the plane of sky angular position of the s/c can be determined
– ∆DOR is a particular type of Very Long Baseline Interferometry (VLBI), that has been used for several decades on deep space missions.
– Recent application of ∆DOR with upgraded equipment at the DSN has enabled unprecedented navigation accuracy on the Mars Odyssey and Mars Exploration Rover missions.
RDB Aug 2005154
Error Sources
A number of errors in the Doppler and range observations limit the accuracy of the orbit determination process:
Oscillator Instability An error in the transmitted frequency of the spacecraft radio signal due to oscillator instability translates into range rate measurement errors.
Instrumental Effects Delays in station and spacecraft electronics represent the major source of error in the ranging system. Thermal noise and instabilities in the signal path through the receiver and telecom subsystem introduce Doppler measurement errors.
Transmission Media Charged particles in the interplanetary medium and Earth’s ionosphere cause propagation delays in radio signals.
Station Locations A longitude error in the station location maps into spacecraft right ascension error, while a station latitude error maps into spacecraft declination error.
Earth Orientation Unmodelled changes in the Earth’s rotation rate, precession and nutation of the spin axis translate into angular position errors.
RDB Aug 2005155
Small Forces Have a Large Effect
Navigation accuracy is also negatively impacted by the mismodelling of “small forces” that impact the spacecraft trajectory.
Solar Radiation Pressure Solar radiation pressure exerts a small but significant force on the spacecraft. If left uncorrected, solar pressure would alter the course of a typical spacecraft on an Earth-to-Mars trajectory by tens of thousands of kilometers.
Reaction Wheel Desaturations For three-axis stabilized spacecraft using reaction wheels, momentum builds up in the wheels over time, requiring periodic thrusting events to despin, or “desaturate” the wheels. These thrusting events must be carefully modeled in the orbit determination process.
Outgassing In the extreme temperatures and vacuum of space, materials on the spacecraft that had accumulated moisture on Earth will outgas. Outgassing is quite noticeable in the navigation solutions for the initial 1-2 weeks following launch.
RDB Aug 2005156
The B-Plane
RDB Aug 2005157
• Navigators often use the B-Plane to describe the arrival trajectory relative to the target body.– The B-Plane is defined perpendicular to the incoming asymptote of the
trajectory.– The “B-Vector” extends from the center of the target body to the point where the
incoming asymptote intersects the B-Plane. Offset distance (∆) is the two-dimensional depiction of the B-Vector (along the T-axis).
– The S-direction is defined // to Vinf– The T-direction is often defined in the mean equatorial plane of the target body– The R-direction is down, such that R x S = T– Arrival conditions are expressed as: B•R, B•T, and TOF.– The B-Plane angle is the angle from the +T direction to the B-vector.
Effect of Arrival Geometry on Entry FPA
MSP’01 Lander navigation based on Doppler and range data types only∆DOR tends to circularize these uncertaintiesFPA Corridor shown is +/- 0.27 degRef: Mase, et. al. “Navigation Strategy for the Mars 2001 Lander Mission, AAS 99-441.
RDB Aug 2005158
Effect of Approach Geometry on Footprint
Ref: Mase, et. al. “Navigation Strategy for the Mars 2001 Lander Mission, AAS 99-441.
RDB Aug 2005159
Thermal Protection System• Required to protect vehicle from intense heating
during atmospheric flight– Peak heat rate dictates material selection– Heat load dictates TPS thickness
• At Mars, SLA-561V is only demonstrated forebody material
• Carbon-phenolic, carbon-carbon, PICA, SIRCA are other commonly used materials for Earth, Venus and Jovian entries
• TPS mass is typically developed with significant margin and has highest aeroshell subsystem mass fraction
RDB Aug 2005160
The Ablative TPS Problem and Potential Material Properties
Material Name Manufacturer Density
(kg/m3)Limit
(W/cm2)
SLA-561V Lockheed-Martin 256 ~ 200
FM 5055 Carbon Phenolic
Fibercote (formerly US Polymeric), Hitco Inc.
1450 > 10,000
PhenCarb-20,24,32
Applied Research Associates (ARA) 320-512 ~ 750
PICA (PhenolicImpregnated Carbon Ablator)
Fiber Materials, Inc. (FMI) 240 > 2500
Avcoat 5026 (Apollo)
AVCO Corp (out of business) 513 > 2500
RDB Aug 2005161
Ablative TPS History of Success, Little Recent Development
No Human Rated Ablative TPS Available Today!
RDB Aug 2005162
Courtesy Bernie Laub, NASA Ames Research Center, 2004.
TPS Mission SummaryMission TPS
MaterialThickness
(cm)TPS mass fraction
Apollo AVCO 4.32
1.38***
1.2***
1.6***
14.6***
1.9
1.0
1.57
5.82
6.0
13.7%
Viking Landers SLA-561V 2.8%
Pioneer-Venus Small Probes
Carbon Phenolic
12.9%
Pioneer-Venus Large Probe
Carbon Phenolic
10.35%
Mars Microprobe, DS-2 SIRCA-SPLIT
Galileo Probe Carbon Phenolic
50.5%, Highest entry of all time; 60 km/s
Mars Pathfinder SLA-561V* 8.2%, First direct EDL
Mars Exploration Rovers SLA-561V* 5.6%
Cassini Huygens Probe AQ60 (Silica) 30.1%
Stardust PICA 22%, Highest speed Earth entry;12.8 km/s
Genesis Carbon-carbon**
18%, Spacecraft mounting required forebody penetrations
Phoenix Mars Lander SLA-561V*
*SLA-561S and SIRCA on backshell
**SLA-561V on backshell
***at the stagnation
point
RDB Aug 2005163Ref: Planetary Mission Entry Vehicles Reference Guide, Version 2, NASA Ames Research Center
TPS Mission Environments and Mass Fractions
RDB Aug 2005164
Courtesy Bernie Laub, NASA Ames Research Center, 2004.
Deployable Systems: Fly Higher, Fly Lighter
RDB Aug 2005165
What Is A Ballute?• BALLoon + ParachUTE = Ballute
– An inflatable drag device– Used for aerocapture or entry
• Low ballistic coefficient provide means to decelerate high in atmosphere with negligible heating
• Technology promises packaging, modularity, and mass advantages
Clamped Ballute Trailing Ballute
RDB Aug 2005166
Recent Deployable System Concepts
Attached AfterbodyInflatable Decelerator
Trailing Ballute
RDB Aug 2005167Inflatable AeroshellClamped Ballute
Ballute Development History• Development began in early ’60’s under contract to
Goodyear to Develop Expandable Terminal Decelerators for Mars Missions-
– Ballute flown on Gemini for high altitude crew escape system
• Ballutes have been and are still used extensively as decelerators for military applications
• Low-level development activity for space applications into early 80’s
– AOTV studies included ballutes for aerocapture & entry– Hampered by analytical and manufacturing limitations,
but potential performance benefit maintained luster• Ballute entry system developed, flight qualified, and
launched by Soviets for Mars 96 Mission– Ballute system would have been used for Mars landing,
but mission was lost due to launch vehicle failure– Follow-on German/Soviet development - Earth return
from orbit (IRDT, 2000-2002) Illustrates key features of technology, including some of the performance benefit (both flights failed due to launch vehicle problems)
RDB Aug 2005168
Ballutes for munitions deceleration
OTV with Ballute Aerocapture into LEO
IRDT-1, IRDT-2 Flight Systems
Ballute Flight Regime
RDB Aug 2005169
Altitude 300 km
Velocity 11 km/s
Heat Rate 0.05 W/cm2
Knudsen # ~100
Altitude 125 km
Velocity 9.5 km/s
Heat Rate 0.8 W/cm2
Knudsen # ~0.1
Altitude 200 km
Velocity 7.6 km/s
Heat Rate 0.05 W/cm2
Knudsen # ~100
Aerocapture at Earth
Lunar Return Trajectory Comparison
0
50
100
150
200
250
300
0 2 4 6 8 10 12
Relative Velocity (km/s)
Alti
tude
(km
)
BalluteCapsule
Capsule Heat Rate
0
50
100
150
200
250
300
350
400
0 2 4 6 8 10 12
Relative Velocity (km/s)
Heat
Rat
e (W
/cm
2)
RDB Aug 2005170
Ballute Heat Rate
0.0
0.2
0.4
0.6
0.8
1.0
1.2
0 2 4 6 8 10 12
Relative Velocity (km/s)
Hea
t Rat
e (W
/cm
2)
Titan Aerocapture Trajectory Comparison
Representative convective heating rate calculated with Sutton-Graves Equation for 1 m reference sphere.
TrailingClampedRigid
Clamped ballute flies slightly steeper trajectory than trailing ballute, but both ballute trajectories flight much higher than traditional rigid aeroshells, resulting in much lower heating rates and dynamic pressures. Deceleration is slightly higher for ballute aerocapture due to its very low ballistic coefficient. (Courtesy Ball Aerospace)
RDB Aug 2005171
Ballute Technical Issues Presently Under StudyIssue Description
Optimal ballute shape Which shape (sphere, disk, toroid, ?) for which missions? Description
Survivability of the ballute Can the membrane material survive the heating and drag?
Flow stability If the flow is not stable, can its effects be tolerated?
System mass Will ballutes be low mass enough to be competitive?
Trajectory robustness How to compensate for atmospheric uncertainties?
Structural integrity How to ensure structural integrity?
Tether design Can the tether(s) be designed to take the heating and stress?
Parent spacecraft deployment and inflation protection
What auxiliary thermal and aerodynamic protection does the spacecraft require? Can this tech. be borrowed from other inflatable efforts? Can its mass be tolerated?
Experimental verification What is a good ground vs space testing mix?
RDB Aug 2005172
Jeffery Hall, “A Review of Ballute Technology for Planetary Aerocapture,” International Conference on Low-Cost Planetary Missions, 00-0382, 2000
Ballute Technology Development Focuses on 5 Risk Areas
Aeroelastic StabilityMaterials and Seaming Flow Stability
Delivery AccuracyPackaging and Deployment
RDB Aug 2005173
(Courtesy Ball Aerospace)
• Ballutes are flexible– Large deformations– Low stiffness– Low natural frequencies
• The flowfield is unsteady– The spacecraft wake interacts
with the ballute• Ballute survivability
– Material tears due to surface flutter
– Tether failure due to pogo between the spacecraft and ballute
– What is the drag of the deformed shape? Is it the same as predicted for the undeformed shape at each point in the trajectory?
0
100
200
300
400
500
600
0 100 200 300 400 500 600 700
Key Issue: Aeroelastic Analysis of Ballutes
RDB Aug 2005174
Unsteady Flow• Flow near peak
dynamic pressure
• Solver: LAURA• Temp. profiles• Continuum flow
Credit: Peter Gnoffo, NASA Langley Research Center
RDB Aug 2005175
Trailing Ballute Configuration & Mass Estimate• Toroid with 5:1 Major/Minor Diameter Ratio at a
42° Trailing Angle Selected as “Lower Bound” Geometry to Minimize Shock/Shock Interaction and Towing Spacecraft Wake Flow Aero Design Issues
– High Temperature Tensile Members React Drag Loads
– Inflatable Columns Provide Path for ToroidInflation and Reacting Compressive Loads
• Upilex film of 0.6 mil thick chosen for ballute material based on test results (factor of safety of 2).
• CAD model was used to compute ballute system mass properties. Seams, and other features add an estimated 25%.
• Mass estimates for inflation, deployment, and separation systems included.
• Total spacecraft mass allocation is 1000 kg. With ballute system mass of 97 kg, the ballute mass fraction is <10% (vs. 30% for aeroshell).
Element Mass (kg)Ballute 59.0Tension cords 6.5Compression members 1.3Seams, etc. (25%) 16.7Inflation gas (N2) 7.9N2 tank 2.0Tether cutter 1.0Packing Box 1.5Plumbing 0.75Total 96.65
RDB Aug 2005176
(Courtesy Ball Aerospace)
Current Ballute Efforts• In-Space Propulsion Ballute Analysis and Test Studies
– Program aimed at increasing the TRL of ballute technologies– Ground-based testing and detailed analyses
• NASA JSC/LaRC ESR&T Inflatable Reentry Vehicle Experiment Technology Maturation Project– Planned flight test series of attached ballutes using a sounding rocket– Rapid development flight tests
• Ball Aerospace ESR&T Ultra-Lightweight Inflatable Thin-Film Ballutes for Return from the Moon– Analysis and ground-based test program focused on increasing thin-
film ballute technology to TRL 6 within 4 years.• Vorticity and Vertigo
– Static coupled analysis of clamped ballutes– Proposed supersonic test program
• Satellite recovery with ballutes
RDB Aug 2005177
Guidance, Navigation, and Control
The GN&C subsystem in an aeroassist vehicle is responsible for inflight solution of the following:• Where am I? (Navigation)• How do I get there from here? (Guidance)• How is the guidance solution implemented? (Control)The answers are obtained with use of on-board computers, software, sensors, and actuators.
RDB Aug 2005178
Guidance, Navigation, and Control– Navigation system – determines current vehicle state– Guidance system – determines required orientation to achieve
desired vehicle state subject to constraints– Control system – responsible for generating forces/moments
used to achieve desired vehicle orientationSystem Drivers– Flight computer– Sensors, flight data– Actuators (aero surfaces, reaction control system)– Vehicle aerodynamics, mass properties– Atmospheric properties, delivery accuracy
RDB Aug 2005179
Aerodynamic Control of Atmospheric Flight Path
Lift
Drag
Relative�velocity
Low density,�low deceleration,�
low heat rate
High density,�high deceleration,�
high heat rate
α
�
φ
RDB Aug 2005180
L/D0
7
6
5
4
3
2
1
0
1.0
2.0
Zcg,�cm
αtrim�
αtrim,�deg
αtrim
Zcg�
Zcg
Vr
�Xcg��—— = 0.27�� D
�
.05� .10� .15�
L/D vs CG Offset for a 70-deg Sphere Cone
RDB Aug 2005181
Precision Landing Miss Distance vs CG OffsetMSP’01 Lander
RDB Aug 2005182
Aerodynamic Control of Atmospheric Flight Path• While in a planetary atmosphere, the controlled use of aerodynamic forces provides an effective means of accomplishing a mission maneuver. Drag is typically used to decrease the vehicle’s energy (aerobraking, direct entry, aerocapture); whereas, lift is typically used to adjust a vehicle’s current trajectory relative to the desired flight path (aerobraking, direct entry, aerocapture, aerogravity assist).• For a fixed center-of-gravity position, bank-angle modulation is used to orient the vehicle’s lift vector about the velocity vector. Additionally, the magnitude of the lift-vector may be altered by allowing for a variable center-of-gravity position or the use of aerodynamic control surfaces.• As the aeroassist vehicle flies through the atmosphere, the vehicle lift-vector may be oriented up or down to place the vehicle in a lower or higher density flight environment. In this manner, control of the energy profile is maintained. This allows satisfaction of mission objectives in the presence of off-nominal flight conditions (e.g., density, aerodynamic, or mass property misprediction, and position and velocity uncertainty). Furthermore, satisfaction of inflight constraint (e.g., heat-rate or deceleration) is ensured.
RDB Aug 2005183
Guidance, Navigation, and Control Schematic
Navigation
MeasuredConditions
Guidance
CommandedOrientation
Control
CommandedMoments
Actuator
ActualMoments
FlightDynamics
RDB Aug 2005184
Guidance, Navigation, and Control Schematic
Based on initial measurement and inflight sensor data, the navigationsystem is responsible for determining the vehicle’s current flight conditions. The guidance system then determines how the vehicle should respond (if at all) such that the inflight and terminal mission constraints are satisfied based on a prediction of the impending flight environment. To adjust the flight path, this system may command a change in vehicle orientation, thereby changing the aerodynamic forces. The control system determines how to best achieve this desired orientation change. Typically, this change would be accomplished propulsively, with use of a reaction-control system (RCS). However, for some missions, the use of aerodynamic surfaces is also a viable option. The orientation change is then effected through the response of the actuators (RCS or control surfaces) which alter the vehicle’s dynamic flight. This process is repeated numerous times throughout the atmospheric portion of flight.
RDB Aug 2005185
Guidance and Control Definition
Guidance– Open Loop: follows predetermined command (e.g. pitch angle vs time)– Closed Loop: uses inflight data to determine command sequence
• Reference Approach – commands generated to follow predetermined profile (e.g. γ vs energy)
• Adaptive Approach – commands generated based on inflight calculations (e.g. predictor corrector)
• Hybrid Approach – uses adaptive updates to adjust reference approach solution
Controls– Passive: no explicit force/moment commands (e.g. vehicle spin)– Active: command the application of auxiliary forces/moments (e.g.
RCS, aero surfaces)
RDB Aug 2005186
Guidance and Control DefinitionOver the years, numerous guidance and control systems have been designed. These systems may be loosely categorized in the manner illustrated on the accompanying chart.
Guidance systems are typically referred to as open-loop or closed-loop depending on whether or not they utilize inflight data to adjust the vehicle’s flight path. Open-loop systems are simpler to design, develop, and test; however; closed-loop systems provide better performance. Many launch vehicle’s use open-loop systems to provide first-stage guidance. Closed-loop systems may be further divided by their level of complexity. The simplest closed-loop system is often referred to as a reference trajectory approach. In such an approach, typified by the Space Shuttle entry guidance, a reference flight path is loaded into the onboard computer prior to flight. During flight, the system is continually attempting to remain on this reference path. A common consequence is the performance of S-turns about the reference path. Another approach, typified by a predictor-corrector system, is to adaptively generate guidance commands based on inflight measurements and a prediction of the impending flight environment. This process of predicting ahead and then correcting for the actual conditions is performed repeatedly during flight. While such a system should yield better performance than a reference trajectory approach, it requires more on-board computational resources. In between these extremes, numerous hybrid approaches have been designed to take advantage of this trade-off between performance and required computation.
Control systems are typically referred to as passive or active depending on their means of providing control. Passive systems like the Mars Pathfinder spacecraft, rely on an initial vehicle spin aerodynamic stability to maintain flight at a given orientation. Active systems like the Viking and Mars 98 Landers rely on auxiliary forces (in these cases reaction-control forces) to provide a desired orientation. Passive systems are clearly simpler, but can not enforce as strict tolerances as active systems.
RDB Aug 2005187
Guidance and Control Aeroassist ApplicationsMission Guidance ControlMercury .................................................................. — RCSGemini ................................................................... ref. traj. RCSApollo ..................................................................... ref. traj. RCSShuttle entry ........................................................... ref. traj. RCS, aero surfaces
Pioneer-Venus ....................................................... — passiveViking ..................................................................... — RCSGalileo .................................................................... — passiveHuygens ................................................................. — passiveMars Pathfinder ...................................................... — passiveMars Microprobe (DS-2) ......................................... — passiveMars Polar Lander ...................................................... — RCSStardust/Genesis .......................................................... — passiveMER ...................................................... — passive
Study Guidance Control
Aeroassist Flight Experiment ................................. hybrid RCSMars Atmospheric Knowledge Working Group ...... adaptive RCSMars Aerocapture ......................................... hybrid RCSMars Precision Landing ......................................... hybrid RCS, aero surfacesNeptune Aerocapture ............................................. adaptive RCS, aero surfaces Manned Mars Mission ............................................. adaptive RCS, aero surfaces
RDB Aug 2005188
Guidance and Control Aeroassist ApplicationsPast, present, and future examples of aeroassist guidance and control are listed on the accompanying chart. It is interesting to note that while the manned missions of the Mercury-Shuttle era successfully demonstrated the use of closed-loop reference trajectory guidance and active reaction-control system control, the robotic Mars missions are using simplified approaches. These systems, like the Viking, Galileo, and Pioneer-Venus systems before them do not employ a guidance system. Instead, event timers are used in conjunction with sensor data to accomplish the mission. Additionally, the Mars Polar Lander is the first robotic flight since Viking to use an active control system.
Beyond this current set of robotic missions, advances are required in guidance and control technology to ensure mission success. The advances are required as a result of the more elaborate use of aeroassist technology in the accommodation of precision landing and aerocapture goals.
To support future robotic missions, hybrid (or perhaps completely adaptive) guidance systems will be required with reaction-control systems similar in complexity to that of the 98 Lander. The use of aerodynamic control surfaces may be needed to satisfy the precision landing requirements for both the Mars sample return and human exploration mission plans.
Fully adaptive guidance and control systems may be required to support future outer-planet robotic missions, where atmospheric and aerodynamic uncertainty is larger. These systems could also be used to reduce the risk associated with piloted missions.
RDB Aug 2005189
Terminal Descent• The terminal descent phase typically begins at supersonic
speeds with parachute deployment– Also governed by dynamic pressure limits
• On a planet with sufficiently atmospheric density, a supersonic parachute is generally followed by larger subsonic parachute(s) to reduce the descent speed to that required for safe landing
• On Mars, the size of a supersonic decelerator is generally so large that it is the same system is typically used for subsonic deceleration– Once subsonic speeds are reached entry system deployments and
extractions are performed to prepare the system for landing– Propulsion is typically used to augment a Mars parachute decelerator
RDB Aug 2005190
Parachute Decelerator System Basics• First-order parachute system objective is to quickly
decelerate vehicle from supersonic to low subsonic speeds
• Parachute deceleration success requires:– Successful parachute inflation– Successful parachute strength (loads)
• Relevant loading tests can be achieved in several ground-based facilities, even at full-scale
• Inflation relevant conditions (supersonic speeds, low Q) can only be replicated at high altitude – costly development program that has not been advanced since Viking program
• No credible parachute inflation physics theory
RDB Aug 2005191
Purposes of Parachute DeceleratorsParachute decelerators typically provide one or more of the following functions:
• Deceleration• Control acceleration• Minimize descent rate• Provide specified descent rate• Provide stability (drogue function)• System deployment (pilot function)• Provide difference in ballistic coefficient for separation events• Provide height• Provide timeline• Provide specific state (e.g., altitude, location, speed for
precision landing)
RDB Aug 2005192
Historical Review
RDB Aug 2005193
Planetary Exploration Missions Using ParachutesVenera 5-14, USSR Venus, 1969-1982Luna 16, 20, and 24, USSR Earth Sample Return from Moon, 1970-1976Mars 2 & 3, USSR Mars, 1971Mars 6, USSR Mars, 1974Viking 1 & 2, US Mars, 1976Pioneer Venus, US Venus, 1978Vega 1 & 2, USSR Venus, 1985Galileo, US Jupiter, 1995Mars Pathfinder (MPF), US Mars, 1997Mars Polar Lander (MPL), US Mars, 1999Beagle 2, UK Mars, 2003Mars Exploration Rovers (MER), US Mars, 2004Huygens, Europe Titan, 2004Genesis, US Earth Sample Return from Space, 2004Stardust, US Earth Sample Return from Comet, 2006
Mars 2 & 3Entry Heatshield
ReleaseRocket-DeployedPilot Parachute
TerminalDescentPilot-Deployed
Main Parachute
Reefed MainParachute Retro-Rocket
Firing
Landing
Full-OpenMain Parachute
Graphic Source: Perminov, V. G: The Difficult Road to Mars - A Brief History of Mars Explorationin the Soviet Union, NASA Monographs in Aerospace History Number 15, 1999.
RDB Aug 2005194
Pioneer VenusEntry
Mortar-Deployed Pilot Parachute at M ~ 0.8, H ~ 67 km
Pilot-Deployed Main Parachute
Heatshield Release
Probe Releaseat H ~ 47 km
RDB Aug 2005195
19 min
3.25 s
~ 1 s
Pilot Parachute: Guide Surface, D0 = 0.76 mMain Parachute: 20° Conical Ribbon, D0 = 4.9 m
Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
Mars Pathfinder
RDB Aug 2005196
EntryMortar-Deployed Parachute at M = 1.7, q = 590 Pa
Heatshield Separation
Lander Separation
Airbag Inflation
Retro-Rocket Firing
Bridle Cut
Bouncing
Rover Deployment
Disk-Gap-Band (DGB) ParachuteD0 = 12.7 m
GenesisMortar-Deployed Drogue/Pilot Parachute at M ~ 1.4, H ~ 33 km
Descent Under Drogue/Pilot Parachute
Drogue/Pilot-Deployed Parafoil
Descent Under Parafoil
Mid-Air Retrieval
Drogue/Pilot Parachute: DGB, D0 = 2.03 mParafoil: S0 = 39 m2
RDB Aug 2005197
Graphic Source: Genesis Sample Return Press Kit, NASA, September 2004.
HuygensEntry
Mortar-Deployed Pilot Parachute at M ~ 1.5Pilot-Deployed Main Parachute
Heatshield SeparationDescent Under Main Parachute
Main Parachute-DeployedDrogue Parachute
Descent UnderDrogue Parachute
Touchdown
~ 2 hr
15 min
30 s
2.5 s
Graphic Source: Cassini-Huygens Saturn Arrival Press Kit, NASA, June 2004.
ParachutesPilot: DGB, D0 = 2.59 mMain: DGB, D0 = 8.30 mDrogue: DGB, D0 = 3.03 m
RDB Aug 2005198
Canopies for Planetary Exploration MissionsThe most commonly used canopies in planetary exploration missions are:
GuideSurface
ConicalRibbon
Disk-Gap-Band
Ringsail
RDB Aug 2005199
Parachute System Components
MortarTubeSabotGas GeneratorMortar Cover
ParachuteDisk-Gap-Band CanopySuspension LinesRisers & BridleDeployment Bag
RDB Aug 2005200
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Parachute System Components
Disk-Gap-Band CanopyDisk
GapBand{
Suspension Lines
Riser
Bridle
VentDeployment Bag
RDB Aug 2005201
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Mortar Components
RDB Aug 2005202
Cover
Sabot
Tube
Gas Generator
AttachmentLugs (3)
Rails (3)
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Parachute System Design and Qualification
RDB Aug 2005203
–Analysis
–Testing
• Drop testing
• Wind tunnel testing– High speed conditions
– Low speed conditions
First-Order Parachute Design
DiskGap
Band
Vent• To first order,
• Relative disk area is primary drag control
• Relative band area is primary stability control
• Vent and gap area balance disk loading and inflation rate
RDB Aug 2005204
Drag vs Stability Trade Space I
RDB Aug 2005205
0
5
10
15
20
25
30
0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0
CD0
Solid Textile Parachutes
Slotted Textile Parachutes
Guide Surface
Ringsail
Disk-Gap-Band
Conical Ribbon
Ave
rage
Ang
le o
f Osc
illat
ion
(AA
O),
deg.
Guide Surface (Ribless) Parachutes
Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
• Low drag (CD0 ~ 0.3) with good stability (0° to ±3° AAO)
• Used in situations where stability is principal consideration (drogue, pilot)
• Abrupt transition at maximum projected diameter and subsequent flow separation is reason for stability characteristics
• Appropriate for subsonic applications
• Difficult to manufacture
• Used by Pioneer Venus (pilot)
RDB Aug 2005206
Conical Ribbon Parachutes
RDB Aug 2005207
• Moderate drag (CD0 ~ 0.5) with good stability (0° to ±3° AAO)
• Appropriate for subsonic andsupersonic applications
• Can be made very strong (especially if manufactured from Kevlar) and deployed at high dynamic pressure
• Relatively high weight per unit drag area
• Used by:Pioneer VenusGalileo
Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
Disk-Gap-Band Parachutes
RDB Aug 2005208
• Low-to-moderate drag (CD0 ~ 0.4 to 0.7) with good-to-moderate stability (±5° to ±15° AAO)
• Drag can be traded for stability by changing the gap and band heights
• Appropriate for subsonic andsupersonic applications
• Strong heritage data at supersonic speeds in low density atmospheres key to its continued use
• Used by:Viking MPF MPL Beagle 2MER Huygens GenesisStardust
Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
Ringsail Parachutes
• High drag (CD0 ~ 0.8) with good-to-moderate stability (±5° to ±10° AAO)
• Design tailored for optimum performance by varying canopy shape and distribution of geometric porosity throughout canopy
• Currently limited to subsonic applications
• Time consuming fabrication
• Relatively light weight per unit drag area
• Used by Beagle 2 and proposed for other missions
Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
RDB Aug 2005209
Recovery System Design Guide
RDB Aug 2005210Graphic Source: Ewing, E. G., Bixby, H. W., and Knacke, T. W.: Recovery System Design Guide, AFFDL-TR-78-151, 1978.
Parachute Aerodynamics
MER Drag Coefficient Estimate:
Wind tunnel testNASA LaRC TDT
Mars Pathfinderflight reconstruction
RDB Aug 2005211
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
CD0 vs M
RDB Aug 2005212
0.35
0.40
0.45
0.50
0.55
0.60
0.65
CD0
0.0 0.5 1.0 1.5 2.0 2.5 3.0M
Viking Parachute Wind Tunnel Test Results in Wake of Aeroshell
Sources: Jaremenko, I., Steinberg, S., and Faye-Petersen, R.: Scale Model Test Results of the Viking Parachute System at Mach Numbers from 0.1 Through 2.6, NASA CR-149377, 1971.Moog, R. D. and Michel, F. C.: Balloon Launched Viking Decelerator Test Program Summary Report, NASA CR-112288, 1973.
Design Effects on CD0 IHow does parachute design affect CD0?
CD0 Comparison
Canopy Type• Example: Ringsail parachutes have higher >
CD0 than Guide Surface parachutes
Geometric Porosity• Parachutes with smaller geometric porosity >
have a higher CD0• Example: Increasing gap size on a DGB
parachute decreases CD0
Fabric Permeability• Reducing fabric permeability increases CD0
RDB Aug 2005213
0.40
0.45
0.50
0.55
0.60
0.1 0.2
CD0
0.3 0.4 0.5
Error Bars at3-Sigma Level
M
1.6 Viking Parachute (Permeable Fabric)
1.6 Viking Parachute (Impermeable Fabric)
Design Effects on CD0 II
RDB Aug 2005214
How does parachute design affect CD0?CD0 Comparison
Suspension Lines Length• Increasing suspension line length >
increases CD0
Trailing Distance*• Increasing trailing distance increases CD0 >
Forebody-to-Parachute Diameter Ratio*• Reducing forebody-to-parachute ratio >
increases CD0
*Due to wake effects of forebody on parachute
Wake Effects on CD0
0.35
0.40
0.45
0.50
0.55
0.60
0.65
0.70
CD0
0.0 0.5 1.0 1.5 2.0 2.5 3.0
Viking ParachuteWind Tunnel Test DataIn Wake of Aeroshell
M
Viking ParachuteWind Tunnel Test DataNo Aeroshell
Sources: Moog, R. D. and Michel, F. C.: Balloon Launched Viking Decelerator Test Program Summary Report, NASA CR-112288, 1973.
RDB Aug 2005215
Dynamics - Importance to Planetary Missions
Dynamic behavior of the entry system during the parachute phase of descent and landing is important for numerous reasons, for example:
• Scientific observations (imaging)
• Sensor performance (radar)
• Separation events (heatshield)
• Initial conditions for propulsiveterminal descent
• Attitude at rocket firing events
• Control of horizontal velocity
RDB Aug 2005216
Design Effects on Stability
Parachute choice and design can be used toaffect stability:
• Guide surface parachute is more stable than a Ringsail parachute
• Increasing band height on DGB parachutes improves stability
• Increasing geometric porosity improves stability
• Increasing fabric permeability improves stability
Stability considerations may drive choice of parachute and its design
RDB Aug 2005217
Stability Comparison of Viking and MPF Designs
RDB Aug 2005218
Inflation Qualification of DGB Parachutes
200
300
400
500
600
700
800
900
1000
Dyn
amic
Pre
ssur
e (P
a)
1.0 1.4 1.8 2.2 2.6 3.0
VikingAV-1
Mach Number
VikingAV-4
VikingAV-2
VikingLanders
MPFLander
Viking Requirement
MPF Requirement
MER Requirement
MER Operations
DGBParachute
FlightHistory
RDB Aug 2005219
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Opening Loads
-10000
-5000
0
5000
10000
15000
20000
0 1 2 3 4 5 6 7
Start of deployment (mortar firing)Mortar recoil force
Snatch load
End of deployment& start of inflation
Peak opening load
Time from Mortar Firing (s)
Load
(lb)
RDB Aug 2005220
Parachute Aerodynamics and Terminal Descent V&V
Multi-bodydynamic analyses
Wind tunnel testNASA LaRC TDT
Multi-bodydynamic test
RDB Aug 2005221
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Froude Number and Dynamic Similarity
RDB Aug 2005222
• Just as Reynolds number is used as a similarity parameter between test and flight when the dominant forces are inertial and viscous, Froudenumber can be used when the dominant forces are inertial and gravitational
Fr = V/sqrt(gL)• Here, dynamic similarity infers that two points in
corresponding positions at corresponding times will have proportional velocities and accelerations
• As such, Froude scaling is generally used when it is important to model the appropriate flight dynamics (e.g., deployments)
MER Froude Scaling Example: Validation of Terminal Descent Algorithms
• Algorithms are designed using simulations of Descent• Simulations are validated against each other but must also be
validated by test. • Ideally this would be a system drop test which includes all flight-like
EM hardware (I.e structures, motors, avionics, FSW, etc). • Perform this Ideal test and confirm that the velocity at ground
impact is within the airbags capabilities. • Unfortunately, this Earth test of Mars Flight Hardware would not
validate performance. (Pendulum Dynamics at incorrect freq and damping, descent rate a factor of 5 to slow so RAD motors would be over-powered, airbags would over-pressurize during ambient impacts.
• Fortunately, dynamic scaling laws exist that define an Earth scaled test for validating aspects of the Mars 3- body descent.
RDB Aug 2005223Courtesy R. Mitcheltree, MER EDL V&V Review, May 2003
Terminal Descent Dynamic Similarity Derivation
tV
LmV
FroudegLVwhere
tV
gLVV
VLtwhere
tttlet
ACmgVwhere
VVVlet
tVmACVmg
maF
D
D
∂∂
⎟⎟⎠
⎞⎜⎜⎝
⎛=−
=⎟⎟⎠
⎞⎜⎜⎝
⎛∂∂
⎟⎟⎠
⎞⎜⎜⎝
⎛=−
==
==
∂∂
=−
=
32
22*2
**
*
**
2
)1(
#)1(
221
ρ
ρ
ρ
mg
1/2ρV2CDA
mg
RDB Aug 2005224
*Assumes CD constant Courtesy R. Mitcheltree, MER EDL V&V Review, May 2003
Froude Scaling: Governing Relation and Options
RDB Aug 2005225
• Define:– Nm = masstest/massflight
– NL = Lengthtest/Lengthflight
– Nρ = densitytest/densityflight
• For –1.3 km Mars (ρ=0.014 kg/m3) and 2.5 km Earth (ρ=0.962 kg/m3),Nρ = 68.7
13 =L
M
NNN
ρ
OPTIONS
FlightTest Lm
Lm
⎟⎟⎠
⎞⎜⎜⎝
⎛=⎟⎟
⎠
⎞⎜⎜⎝
⎛33 ρρ
NM NL NT NV Nt NMach NRe
0.50 7.7412.5103447
0.611.212.0
0.260.3040.621.0
NI Nk
0.38 0.18 1.0 0.67 0.0120.0668.78810
5.61.0 0.244 2.64 0.8 10.868.7 1.0 181 1.6 1811264 2.64 3337 2.64 1264
Mass Lengths Forces Velocity Time Mach Reynolds Inertias Stiffness
Courtesy R. Mitcheltree, MER EDL V&V Review, May 2003
Further Study: Parachute Decelerators
Bixby, H. W., Ewing, E. G., and Knacke, T. W.: Recovery Systems Design Guide, AFFDL-TR-78-151, 1978.
• Comprehensive (458 pages)• Extensive bibliography (> 500) referenced through text• Published 27 years ago - some sections (e.g., materials) are outdated• As with all documents, watch out for typos and incorrect information• Required reading for engineers involved in the development and
qualification of aerodynamic decelerators for planetary entry systems
Knacke, T. W.: Parachute Recovery Systems Design Manual, Para Publishing, Santa Barbara, California, 1992.
• Comprehensive (~250 pages)• Extensive bibliography referenced through text• Similar to Recovery Systems Design Guide - not as comprehensive but
more up-to-date• Required reading for engineers involved in the development and
qualification of aerodynamic decelerators for planetary entry systems
RDB Aug 2005226
Landing Systems• Typically less than 1% of energy remains at landing;
however 10s of Earth gs are possible at impact• Surface unknowns (e.g., surface, rocks, slope) greatly
complicate this event• Multiple (mission unique) systems under development:
– Airbags: MPF, MER– Propulsive Touchdown: Viking, MPL, Apollo, Phoenix– Sky Crane: MSL– Hard Landing:
• Penetrators: DS-2• Energy absorption: MSR EEV, Pallet Landers
• MER team greatly expanded airbag capability
RDB Aug 2005227
Landing System: Legs or Airbags?Landed Mass Delivery Capability Advantage: Legs• Lander legs can be scaled to accommodate significantly greater landed mass than
airbags. Three- and four-legged designs have been developed and used successfully on Mars (Viking) and lunar (Apollo) missions. Landed mass capability of 15,000 kg has been demonstrated on the Moon.
• The MER landed mass of 550 kg required a new dual-bladder design to succeed. A landed mass approaching 600 kg may be the limit for airbag technology.
Robustness to Surface Hazards Advantage: Airbags• Legged landing systems typically have low ground clearance (22 cm for Viking, 33 cm for
Mars ‘01/Phoenix), and are susceptible to rock impacts on leg stabilizers and lander undercarriage. This liability drives legged landers to low rock abundance sites.
• Surface slopes pose tipover risk for legged landers. Mars’01/Phoenix surface slope capability of 10° drove a requirement for an RMS surface slope within the landing ellipse of no greater than 4°.
• Airbag systems can land on rocks exceeding 0.5 m, and are robust to surface slopes exceeding 20°. These capabilities allow airbag landing systems to be targeted to a much more diverse range of landing sites than legged landers.
RDB Aug 2005228
Landing System: Legs or Airbags? (cont.)Surface Contamination Advantage: Airbags• Descent engine plume interaction with the surface is a major concern with legged
landing system. Propellants can contaminate the soil and atmosphere (atmospheric contamination will disperse with time). Plume impingement on the surface can kick up significant dust, sand, and small rocks, and can excavate holes in the surface up to 0.5 m deep1.
• Surface contamination from airbags and gas generation system may be detectable. Interior scrapings of airbags have shown organic material.
Rover Deployment Advantage: Toss-Up• Rover deployment from top deck of a legged lander may involve ramps or robotic
arm (Mars ‘01 design). Robotic arm approach allows a broader range of deployment azimuths than egress via ramps.
• Rover deployment from an airbag landing system can be impeded by partially-deflated or partially-retracted airbags (Ex: Spirit).
1Ref: Albert Haldemann, JPL
RDB Aug 2005229
RDB Aug 2005230
MER Airbag Capability Map Update
0
4
8
12
16
20
0 4 8 12 16 20 24 28
Tangential Impact Velocity (m/s)
Nor
mal
Impa
ct V
eloc
ity (m
/s)
Radar Bracket Damage
Verified In-Spec. Capability as of Apr. 03
Capability Added by Qual Drop Tests in Aug. 03
Out of Spec. Region
Qual Drop Passed
Qual Drop Resulted in Bag Modifications
45° Tests
30° Tests
18° Tests
90° Tests
Development Tests
Vertical Limit Varies Between 10 – 17 m/s Depending on Rock Size and Shape, and
Bag Impact Orientation
All new tests show very good
performance
Grazing angle impacts (yellow) are now considered safe.
Ref: Rob Manning MER presentation, Oct 2003
RDB Aug 2005231
Landing System Evolution: From Legs to Wheels
Mars 98 & MSR
•The failure of the M98 lander mission during MSR’sphase A, led to a change in risk posture on landing robustness.
•Several review boards and tiger teams were assembled to redirect MSR’slanding/EDL architecture.
•Robust rover egress for MSR was never addressed.
Mars Smart Lander
•Extensive evaluation of many different EDL and Landing architectures suitable for MSR were studied.
•Pallet style landing system with a large rover was selected based on expectation of a 2005 launch.
•Pallet greatly improved egress and landing safety.
Mars Expl. Rover
•MSL mission was delayed to 2007 and then 2009, resulting in more time to develop technologies.
•MER made a large investment in developing multi-body control dynamics.
•MER discovered the hidden challenges and costs ofegressing a rover.
Mars Science Lab.•EDL architecture given one last “fresh” look, focused on:
Cost ReductionPerformance IncreaseEDL Feed Forward
•Desire to incorporate best lessons and technologies from MER; multi-body control, DRL…
•Further advancement of sensor technologies & HDA
• Sky Crane invented
Courtesy: T, Rivellini, JPL
MSL Sky Crane Nominal Timeline
RDB Aug 2005232
Entry Interface
Deploy Supersonic Chute
Jettison Heatshield, Activate Radar, and Deploy Mobility
M = 2.2
*L/D = 0.18*Hypersonic Aeromaneuver Guidance
Jettison Chute and Backshell, Begin Powered Descent
Velocity Altitude AGL
Timeline:E + 0 s
h = 8.0 kmγ = -15.0 degV = 491 m/s
r = 3522.2 km
Begin Sky Crane Maneuver
2500 m above MOLA areoid
Flight Path Angle
Flyaway
h = 5.7 kmγ = -31.3 degV = 179 m/s
h = 2.0 km
h = 1.0 kmγ = -89.7 degVv = 95 m/s, VH = 30 m/s
337.3 s323 s304.0 s233.0 s 255.4 s
Mach
M = 0.8
Rover Touchdown
Sense Velocity with Radar
h = 28 mVv = 3 m/s, VH = 0 m/s
Rover Touchdown
71 s
48.6 s
33.3 s
Courtesy: T, Rivellini, JPL
MSL Sky Crane Maneuver Description
RDB Aug 2005233
One Body Vertical Descent Phase
Vehicle has just transitioned from approach phase to sky crane by achieving its altitude velocity way point.
Once the preset altitude and velocity targets are achieved GNC switches the radar off and navigates by IMU propagation & the rover is commanded to be released from the descent stage
Deployment Phase
The DRL allows the rover and descent stage to separate with a predetermined separation profile.
The data and communications umbilical is deployed.
The vehicle has undergone major changes in configuration, mass properties, modal properties, and has introduced 2-body pendulum dynamics.
Two Body Vertical Descent Phase/Constant Velocity
The system will enter this phase while the rover is still 2-3 meters above the surface under 3 sigma worst case.
After the rover has been fully deployed the system enables the touchdown logic to start.
The TD logic looks for a persistent reduction in the averaged throttle settings.
Touchdown Phase
As soon as the rover begins to contact the surface the DS will throttle down to maintain its .75 m/s downward velocity.
Continued DS downward motion causes bridle tension to reduce/disappear which minimizes/eliminates rover-terrain interaction disturbances to the DS
When the TD logic determines that touchdown is complete, the bridle and umbilical are released.
Fly-Away Phase
Just prior to separation, the DS micro-controller is initialized and handed control of the DS ascent guidance.
Flyaway is performed via an open loop vertical ascent followed by a turn and burn profile optimized to maximize DS flyaway distance.
Courtesy: T, Rivellini, JPL
MSL Entry, Descent and LandingThe Next Big Thing
RDB Aug 2005234
Rover
Descent Stage
Aeroshell Comparison
MSL: 4.7 m dia.
MER/MPF: 2.7 m dia.
Viking: 3.5 m dia.
MSL Entry, Descent and LandingThe Next Big Thing
RDB Aug 2005235
Rover
Descent Stage
MSL MER
Entry Mass, kg 2000 832Descent Mass, kg 1700 743Delivered Mass, kg 725 420
Breaking Out of the Viking Box:Current Landed Mass and Elevation Limits• All five successful landers
– Had touchdown masses < 0.6 MT– Landed at low elevation sites, below -1 km MOLA– Had landed footprints on the order of 100s of kms– Based on large technology investment made in the late
1960s and early 1970s as part of the Viking program• Aerodynamic characterization of 70-degree sphere cone• Lifting entry• SLA-561V forebody TPS• 16 m diameter supersonic DGB parachute• Autonomous terminal descent propulsion
RDB Aug 2005236
Mars above 2.0 km in Black
MOLA Topography ±90º Lat, 180º to -180°W Lon
RDB Aug 2005237
Black area is topography > 2.0 kmLines at ±50º and ±60º latitude
Mars above 1.0 km in Black
RDB Aug 2005238
Black area is topography > 1.0 kmLines at ±60º latitude
Mars above 0 km in Black
RDB Aug 2005239
Black area is topography > 0.0 kmLines at ±60º latitude
RDB Aug 2005240
But -1.0 km is the BEST we have been able to do with our Heritage Viking-Technology
• So far the Southern hemisphere has been largely out of reach.
Mars Elevation Variation• The highest landing to-date is Opportunity at Meridiani
Planum (-1 km MOLA).• We are still 2 km below the flanks of the Highlands.
RDB Aug 2005241
Pathfinder
Gusev
MeridianiAncientHighlands
NorthernLowlands
Viking
Courtesy Rob Manning, JPL
What Limits Our Landed Mass, Landed Elevation Capability
• Entry Vehicle– Larger diameter lowers β (or allow more mass for the same β) providing
higher altitude deceleration. 5 m diameter is the largest we can fit in today’s launch vehicles. Need new launch vehicle or wider fairing
– Lift (L/D on order of 0.25) can gain as much as 3 km compared with ballistic (MPF/MER) entry
• Supersonic Parachute– 16.15 m diameter is the largest qualified chute (Viking)– Parachute inflation dynamic pressure limit as high as 800 Pa– Parachute inflation Mach limit as high as 2.2
• Atmospheric Density Variability and Dust Effects– Significant atmospheric variability across a Mars year (+/-3 km
elevation impact) limits our ability to develop a common EDL system– Variability within an opportunity can cause +/-1.5 km elevation impact– Significant dust content can cause loss of approx. 3 km elevation
RDB Aug 2005242
Environmental Effects: Density cycle• Atmosphere Variation
– Density drops in the winter by 30% (moves to poles)
– Imagine if in the winter the Shuttle had to land at 10,000 ft!
• EDL performance varies as much as +/-3 km– Performance is latitude
dependant and decade dependent
20092018 20202011 20162013
Northern
Southern Winter Summer
Summer Winter
RDB Aug 2005243Courtesy Rob Manning, JPL
Impact of Atmospheric Variability on Landed Elevation
RDB Aug 2005244
-8
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Delivered Mass (kg)
Optimized EDL
• Performance for the 2013 opportunity• L/D = 0.24• Two parachute system used, Viking supersonic and 110 ft. subsonic• Maximum aeroshell diameter used (5.0 m)
No Dust Performance
Tau = 3 Performance
OpportunityVariation
Courtesy Rob Manning, JPL
RDB Aug 2005245
-6
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Delivered Mass (kg)
Deli
very
Alt
itude (
MO
LA
, km
Common+PDS, Dust Common, Dust
propellant if pinpoint landing is desired
Courtesy Rob Manning, JPL
Altitude Capability with Large, High Mach Supersonic Parachute
• 30 m, Mach 2.7 chute• Dust degraded performance assumed
– τ = 3.0 (common in the next decade)• Subtract 150 kg from delivered mass for additional
Breaking the Viking Box
RDB Aug 2005246
• Next Generation Supersonic Parachute– The lowest risk technology would be to re-qualify a 30 m diameter
Mach 2.7 parachute– Performance gains of 5-6 km in altitude– Of order $100M / 3-4 year investment!
• Larger launch vehicle fairing and aeroshell diameter– 6.5 m LV fairing would allow for ~6.0 m aero-shell– Performance gains of ~1-2km in altitude possible in conjunction with
larger parachute– Unknown impact on launch vehicle cost & performance
• Other technology– Inflatable aerodynamic decelerators (ballutes)– Large subsonic chutes– Significant (perhaps even supersonic) propulsion
Courtesy Rob Manning, JPL
Where Do We Need To Go By 2030?The robotic program is presently attempting to develop systems that deliver 0.8 MT for MSL by 2009.
With a new supersonic decelerator we can get to 1.3 or 1.5 MT.Maybe by 2020.
But the next step is across an ocean!We need to develop EDL systems that can get 30-60 MT of payload
down to surface per landing (60-120 MT entry mass).
Will these human-scale EDL systems look anything like today’s robotic landers?
Probably not.
RDB Aug 2005247Courtesy Rob Manning, JPL
Humans on MarsHuman Mars Missions
RDB Aug 2005248
Apollo landed 10,000 kg
will require > 3X this mass: 30,000 kg - 60,000 kg> 50x larger than MER!
Can we land something this big?
Human-Scale Entry Systems Must Be of Large Diameter
20 m Diameter
-10-505
1015202530354045
0 10 20 30 40 50 60 70 80 90 100Mass (T)
Fina
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(km
)M = 4 UM = 3 UM = 2 UM = 4 OM = 3 OM = 2 O
10 m Diameter
-10-505
1015202530354045
0 10 20 30 40 50 60 70 80 90 100Mass (T)
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(km
)
M = 4 UM = 3 UM = 2 UM = 4 OM = 3 OM = 2 O
5 m Diameter
-10-505
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0 10 20 30 40 50 60 70 80 90 100Mass (T)
Fina
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(km
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M = 4 UM = 3 UM = 2 UM = 4 OM = 3 OM = 2 O
15 m Diameter
-10-505
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0 10 20 30 40 50 60 70 80 90 100Mass (T)
Fina
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(km
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M = 4 UM = 3 UM = 2 UM = 4 OM = 3 OM = 2 O
RDB Aug 2005249
L/D = 0.5Entry from orbit corridorVatm = 4.63 km/s
The Mars Atmosphere is Too Thin To Slow This Much Mass Down….
• Initial descent conditions strongly dependant on β– For a 5 m diameter aeroshell and 10 MT entry mass, Mach 2 is
reached at 2 km altitude. 50 MT vehicle reaches 0 km altitude atMach 4.
– For a 10 m diameter aeroshell, Mach 2 is reached at 10 km altitude for entry masses below 20 MT. A Mach 4 decelerator would enable entry masses as high as 100 MT (above 12 km).
– For a 20 m diameter aeroshell, Mach 2 is reached at 10 km altitude for entry masses below 80 MT.
• As we will see, a Mach 2 parachute that decelerates 80 MT to theground must be quite large
RDB Aug 2005250
… Viewed Another Way
• Red areas are above Mach 1• On Earth, terminal velocity never gets above Mach 0.4• On Mars, as mass increases terminal velocity becomes supersonic
- Below 20 km on Mars, for β < than 100 kg/m3 (current experience), Vt is subsonic.- For 5m diameter aeroshell (robotic exploration limit), β = 100 kg/m3 for 3 MT- For 10m diameter aeroshell, β = 100 kg/m3 for only 12.5 MT- For a Mars entry mass of 100 MT and a 15m diameter aeroshell, β = 350 kg/m3
RDB Aug 2005251Courtesy Rob Manning, JPL
Will Human Scale Descent Systems Use a Parachute?
• Viking DGB parachute technology is limited by Earth-based testing (qualification) to Mach numbers below 2.2 and diameters below 15 m– Mach 1.8 is highest Mach deployment
known to be successful on Mars (MER).
– MPL chute deploy targeted Mach 2.1– 15 m diameter is largest parachute
deployed on Mars (Viking)• In an effort to improve landed mass,
robotic program may one day pursue a larger diameter supersonic chute– Parachute systems as large as 30 m
diameter are about as large as can be envisioned
– Deployment Mach number is not likely to increase above Mach 2.7 due to heating concerns
– Is a 30 m supersonic chute an efficient decelerator for human-scale EDL?
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
200
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Dyn
amic
Pre
ssur
e (P
a)1.0 1.4 1.8 2.2 2.6 3.0
VikingAV-1
Mach Number
VikingAV-4
VikingAV-2
VikingLanders
MPFLander
Viking Requirement
MPF Requirement
MER Requirement
Inflation Qualification of Mars Disk-Gap-Band Parachutes
RDB Aug 2005252
Parachute Sized to Reach 50 m/s at h = 0 m• 0.02 s/m inflation constant• Terminal velocity of 50 m/s at 0 m altitude• Assumes technology development to allow Mach 3.0 deployment • Required parachute diameters is outside reasonable range (55 – 185 m)
Required Parachute Diameter for M = 3.0 Deployment to 50 m/s at h = 0 m (Curve Fit)
50
70
90
110
130
150
170
190
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Mass (T)
Para
chut
e D
iam
eter
(m)
RDB Aug 2005253
Mach 3 Parachute Deployed to Reach h = 2 km, M = 0.8Required Parachute Diameter for M = 3.0 Deployment
to Reach M =0.8 at h = 2 km
10
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30
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Mass (T)
Para
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10 m12 m15 m20 m
Required Parachute Diameter for M = 3.0 Deployment to reach M = 0.8 at h = 2 km for Conservative Density
20
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Mass (T)
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10 m12 m15 m
• Inflation constant of 0.02 s/m.• 30 m diameter, Mach 3 parachute allows for subsonic gravity turn
maneuver if entry mass below 35-40 MT• Supersonic parachute diameter in the 45-65m range seems unlikely
RDB Aug 2005254
• 30 m chute only viable for 20 MT if conservative atmospheric density assumed (30% reduction). Chute diameters in the 60-80 range required for 80-120 MT.
30 m Parachute with Gravity Turn: ∆VPropulsive Descent ∆V Requirement
10 m Aeroshell, No Parachute
0
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∆V
Req
uire
men
t (m
/s)
Mach 3 InitiationMach 2 InitiationMach 1 InitiationMach 0.8 InitiationMach 0.5 Initiation
Propulsive Descent ∆V Requirement10 m Aeroshell, 30 m Parachute Deployed at Mach 3
0
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1400
0 20 40 60 80 100 120
Vehicle Entry Mass (T)
Mach 3 InitiationMach 2 InitiationMach 1 InitiationMach 0.8 InitiationMach 0.5 Initiation
Propulsive Descent ∆V Requirement10 m Aeroshell with Aeroshell Release and Gravity Turn Initiation
0
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∆V
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Propulsive Descent ∆V Requirement10 m Aeroshell, 30 m Parachute Deployed at Mach 3 with Aeroshell Release
0
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Vehicle Entry Mass (T)
Mach 3 InitiationMach 2 InitiationMach 1 InitiationMach 0.8 InitiationMach 0.5 Initiation
RDB Aug 2005255
30 m Parachute with Gravity Turn: ∆V
• A 30 m parachute allows burns to begin at Mach 0.8 for entry masses less than 30 MT.
• A 30 m parachute allows burns to begin at Mach 1.0 for entry masses less than 50 MT
• For entry masses over 50 MT, a larger chute is required or the propulsive maneuver must be initiated supersonically.
RDB Aug 2005256
The “Supersonic Transition Problem”• How do we get from here …
to here …
• How do we …– Slow from Mach 5 to subsonic– “Undress” and re-orient– Translate to the landing site– Before hitting the ground?
(at >100k ft Earth-density altitude)
in 90 s??RDB Aug 2005
257
EDL System OptionsOptions for HypersonicDecelerators
Options for SupersonicDecelerators
Options for SubsonicDecelerators
Options for Terminal Decent
RDB Aug 2005258
RDB Aug 2005259
14
Past mission experience: Viking, MPF, MER
CURRENT READINESS LEVEL ( Low Med. High)
Aeroassist Technology Readiness Mars Entry Missions
Assessment: Mars direct entrytechnology is at a moderate level of readiness for velocities below 7.5 km/s and low L/D configurations. Aerocapture capability requires demonstration. Larger entry vehicles (for the piloted missions) and/or alternate shapes lack technology foundation.
Viking
Aerothermodynamics3D Non-ablatingCO2 chemistry3D Ablating, radiatingTransition turbulenceDynamic stability
GN&CPassive systemsRef. traj. guidanceAdaptive guidanceRCS controlsAero control surfacesOptical Nav.
TPS1970’s ablatorsNew LCA’sMan-rated ablatorsShock tubes and ballistic rangesArc-jets (high enthalpy CO2)
Vehicle DesignLow L/D shapesMed L/D shapesHigh L/D shapes
Mars-Pathfinder
Aeroassist Technology ReadinessEarth Entry Missions
RDB Aug 2005260
Assessment: Earth entry aeroassist technology is at a moderate level of readiness due to past experience and ongoing mission studies. Experience is limited to velocities at or below 11 km/s. Past missions also relied on large safety margins.
Past mission experience: Mercury, Gemini,Apollo, Shuttle, Genesis, Stardust
CURRENT READINESS LEVEL ( Low Med. High)Aerothermodynamics3D Non-ablating3D Ablating, radiatingTransition turbulenceDynamic stability
GN&CPassive systems Ref. traj. guidanceAdaptive guidanceRCS controlsAero control surfaces
TPS1960’s ablatorsNew LCA’sMan-rated ablatorsShock tubes and ballistic rangesArc-jets (med. enthalpy air)
Vehicle DesignLow L/D shapesMed L/D shapesHigh L/D shapes
11 km/sec0.4 kW/cm2
14.1km/sec2 - 3 kW/cm2
- weak ionization- small ablation
- strong ionization- massive ablation
1960’s Capability Future Capability(Apollo) (Mars Return)
Piloted Mars Missions
Aeroassist Technology ReadinessOuter Planet Entry Missions
RDB Aug 2005261
Assessment: Outer planet aeroassist technology is at a low level of readiness. Aerothermodynamics and TPS models are highly uncertain. GN&C and design enhancements for aerocapture are required. Ballutetechnology is promising, but at a low level of readiness.
Galileo
Note: Galileo survived Jovian entry at the cost of:• launch on Shuttle-Centaur (6 yr trip)• 50% TPS mass fraction (small payload)• supporting capability has atrophied
Past mission experience: Galileo, Huygens
CURRENT READINESS LEVEL ( Low Med. High)AerothermodynamicH2/He chemistry3D Ablating, radiatingTransition turbulenceDynamic stability
GN&CPassive systemsRef. traj. guidanceAdaptive guidanceRCS controlsAero control surfacesOptical Nav.
TPSHighly ablating TPSShock tubes and ballistic rangesArc-jets (high enthalpy H2/He)
Vehicle DesignLow L/D shapesMed L/D shapesHigh L/D shapes
• Aerothermodynamics: Prediction of flowfield surrounding entry vehicle to determine aerodynamic forces and surface heating conditions.Impact: Reduce uncertainties -> smaller safety factors -> mass & cost decrease
• TPS: Protective material system surrounding entry vehicle, designed to maintain specified spacecraft structure and payload temperatures. Impact: Lightweight TPS -> Smaller launch vehicle & useful payload mass
increase• GN&C: Actively control vehicle attitude and trajectory during entry
Impact: Enables precision landing and aerocapture missions• Vehicle Design: Optimized integration of entry vehicle systems to meet mission
requirementsImpact: Drives technology focus & assures project goals are met. Allows design
problems to surface before Phase C/D
Aeroassist Technology investment will enable exciting planetary missions, allow for larger payloads, or use of smaller launch vehicles. Technology investment is required to enable advanced robotic missions, like MSR, and eventual human exploration.
Aeroassist TechnologyInvestment Returns
AeroassistTechnology
WorkshopJanuary
1997Pasadena, CA
1
RDB Aug 2005262
Aeroassist Systems Deliver Large Payload Mass Ratio
Entry Mass(kg)
Landed Mass(kg)
Payload Mass(kg)
Pathfinder 585.3 360.8
540.4
MSP’01 AerocaptureOrbiter
565 N/A 344(61%)
N/A
256.8(44%)
MER-B 832.3 420.8(51%)
WercinskiNeptune AerocaptureStudy
200 150(75%)
• Entry mass: total mass of entry system including its payload at atmospheric entry interface
• Payload mass:– Landed mass (lander): total
mass of system landed (generally excludes most entry system components)
– Payload mass: total mass of payload delivered into orbit (generally excludes all entry system components)
RDB Aug 2005263
Importance of Simulation
RDB Aug 2005264
Simulation Process Overview• EDL simulation is performed via multiple-stage Monte Carlo process that
starts at the time of cruise stage separation and ends with airbag roll stop.• Within a Monte Carlo run, individual trials that “pass” are those that do not
encounter a situation, typically with respect to its velocity or the terrain, that is considered to be beyond the validated capability of the system.– Cases that “fail” only result from an interaction of the vehicle with its
environment.– No provision is made to model the reliability of mechanical assemblies or
electronic components with respect to manufacturing quality.– No provision is made to model cases where the vehicle exceeds its
specifications, but may continue to operate in some degraded fashion.• Statistics are tracked regarding the types of situations that cause Monte
Carlo trials to fail.• Monte Carlo runs are landing site specific in order to capture the
convolution of hazards unique to a specific site.
Ref: Wayne Lee, MER presentation, March 2003.
RDB Aug 2005265
High Level Goals of EDL Simulation• Demonstrate that the vehicle’s performance under
nominal or expected conditions is adequate to ensure landing success.
• Enumerate the performance margins.• Determine whether the predicted probability of success,
given modeling and environmental uncertainties, is commensurate with the risk posture of the project.
• Identify areas of risk where the system is operating “close” to the performance bounds of one or more of the subsystems.
• Identify areas of risk where the system performance is sensitive to modeling or environmental assumptions.Ref: Wayne Lee, MER presentation, March 2003.
RDB Aug 2005266
End-to-End Systems Validation Approach• Computer simulation of EDL performance is utilized for system validation
because it is not possible to conduct a meaningful end-to-end systems test on Earth.
• In order for the simulation to produce meaningful results, its input models must be verified by completion of the following activities:Element Tests and Analysis
– Does each EDL element survive its interaction with the environment?– Does each element deliver its functional performance as advertised?– What forces are generated and how does it fly through the atmosphere?
Interaction Tests– How do two or more elements behave when working and interacting with each other?
Testbed and ATLO– Are the proper commands (with the proper timing) generated by avionics and flight
software when presented with a set of flight-like sensor data? Environmental Modeling
– What conditions (e.g., atmosphere, terrain) will the vehicle encounter during flight?
Ref: Wayne Lee, MER presentation, March 2003.
RDB Aug 2005267
Simulation Process Overview• Major stochastic variable
groups are:– Mass properties– Aerodynamics– Retrorocket performance– Environment (terrain,
atmosphere)• Vehicle state at the end of a
stage is saved and used as the input for the next stage.– Changes to the simulation
inputs only requires a re-run of downstream stages.
Cruise Stage JettisonSimulation
Hypersonic FlightSimulation
Terminal DescentSimulation
Terrain InteractionSimulation
Vehicle Mechanical ModelPost Vent Attitude Distribution
State at Chute Deploy(pos, vel, attitude & rate)
Post Separation State (attitude & rate)
Final Navigation StateVehicle Aerodynamic ModelAtmosphere ModelParachute Deploy Algorithm
Vehicle Aerodynamic ModelVehicle Mechanical ModelAtmosphere and Wind ModelRAD / TIRS AlgorithmsRadar / IMU / Rocket ModelsDigital Terrain Map
State at First Impact(pos, vel)
Airbag Survivability MapDigital Terrain MapsRock Distribution Model
Ref: Wayne Lee, MER presentation, March 2003.
RDB Aug 2005268
RDB Aug 2005269
MER End-to-End EDL V&V by Overlapping Simulations
Cruise Stage Separation
Exo-Atmospheric Coast
Atmospheric Interface
Parachute Mortar Fire
Heatshield Separation
Lander Deployment
RAD Fire
First Impact
Roll Stop
Interplanetary NAV ADAMS Cruise Stage Sep Analysis
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POST Mortar Cover Sep
POST H/S Sep
ADAMS H/S Sep
ADAMS Lander Deploy
Mul
t i-B
o un c
e
Primary Validation PathCourtesy R. Mitcheltree, MER EDL V&V Review, May 2003
MSP’01 Simulation
Guidance�Various
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RDB Aug 2005270
3-DOF vs 6-DOF Simulation
RDB Aug 2005271
• Degrees of freedom (DOF) refers to the number of equations of motion solved in a given trajectory simulation
• Typically, a 3-DOF trajectory simulation solves the translational equations of motion (F=ma); while a 6-DOF simulation solves both the translational and rotational equations of motion.– A static trim equation may be added to 3-DOF simulations to improve the
estimate by approximating vehicle attitude over time• Position, velocity, deceleration, heating and event timing are generally
well predicted by a good 3-DOF simulation, particularly if the vehicle attitude is “known” (controlled within a small uncertainty by spin, aerodynamic surfaces or a RCS system) or approximated by solving a static trim equation.– Ballistic, lifting and guided trajectories can be accurately modeled with 3-
DOF simulation• 6-DOF simulation is required when vehicle dynamics are significant,
attitude information is of concern (e.g., the angle-of-attack at parachute deployment), or control system performance is being assessed.
• Multi-body simulations solve these equations of motion for each modeled body
Simulation Approaches
RDB Aug 2005272
• Parametric studies– Best way to get a physical feel for the problem, quickly understand
trade-space and assess major drivers– In many cases, entry system trade-space can be defined with O(10)
simulations. Response surfaces can also be employed to improve approximation of trade-space
– Generally employed in pre-Phase A through Phase B, but also of use through Phase E.
– Slides 39 and 43 are examples
• Monte-Carlo studies– Statistical distributions for each input variable defined– Variable ranges are randomly sampled and a significantly large
number of simulations are performed (e.g., 2000)– Simulation outputs are represented statistically– Important insight can be gained by assessing tails of distributions &
statistical outliers– Slide 148 presents an example output
Atmospheric Environment
• Significant uncertainty remains in estimation of planetary atmospheric density and winds, even for Earth
• Different gas composition– Not a large driver at Mars– A radiation driver at Venus– A significant aerodynamic/aerothermodynamic
driver at the outer planets
RDB Aug 2005273
Challenges of Mars Atmospheric Flight
-10
0
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175 200 225 250 275 300 325 350
Speed of Sound (m/s)
Alti
tude
(km
)
EarthMars
-10
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10
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50
0.001 0.01 0.1 1 10
Density (kg/m3)
Alti
tude
(km
)
EarthMars
• Mars surface density equivalent to 34 km on Earth• Low density reduces deceleration effectiveness• Altitude-Velocity profile affected by 2/3 lower speed of sound• Large uncertainties in density and winds
RDB Aug 2005274
Mars Atmospheric Density Uncertainty
RDB Aug 2005275
.5 2.0 2.5
25
50
75
100
125
1.0 1.5ρ/ρnom
0
Altitude,�km
ρnom:τ =1.0
Dust Storm Impact on Atmosphere
RDB Aug 2005276
Typical Simulation
Inputs
RDB Aug 2005277
Entry Aerodynamics and Aerothermodynamics Characterization
• Detailed aerodynamic database are developed specific to each project using a combination of sophisticated computational methods and existing ground-based test facilities
• Aerothermodynamic analysis are performed in a similar fashion and integrated with detailed thermal models to evaluate TPS response to the heating environment. TPS validation is typically performed on small coupons in the Ames arc-jet complex.
RDB Aug 2005278
Typical Aerodynamic Inputs
RDB Aug 2005279
Test Facilities
RDB Aug 2005280
Aerodynamic/AerothermodynamicTest Facilities
• NASA has a significant investment in aerodynamic/ aerothermodynamic test facilities
• Most of the facilities are located at NASA Langley or NASA Ames
• Air Force/DoD has additional relevant facilities• Ground-based testing in relevant conditions• Test data for many entry configurations has been
validated in flight. As such, ground-based testing is typically used to validate computational analyses.
• As NASA reduces its infrastructure costs, many of these facilities are under consideration for closure, a potentially significant loss of capability and expertise.
RDB Aug 2005281
Ground-Based FacilitiesNo ground-based facility can completely match relevant flight environment
Wind-tunnels (e.g., LaRC Mach 6)• Good match for aerodynamic environment• Provide controlled environment for precise data analysis• Model-size constraints• Match aerothermodynamic environment, but not at proper enthalpy
Arc-jets (e.g., Ames IHF)• Good match of flow enthalpy• High cost, short test times• Poor characterization of free-stream conditions• Model-size and dimensionality constraints
Ballistic range (e.g., Eglin AFB range)• Quiescent free-stream, no sting• Model-size constraints• Uncertainty in data analysis
RDB Aug 2005282
NASA Langley Aerothermodynamic Facilities
RDB Aug 2005283
15-Inch Mach 6 Hi Temp. Air22-Inch Mach 15/20 He
31-Inch Mach 10 Air 20-Inch Mach 6 Air 20-Inch Mach 6 CF4
MSR EEV Hypersonic Reorientation Wind Tunnel TestsFacility: Mach-6 CF4 and Mach-20 Helium NASA LaRC FacilitiesTest Objective: Evaluate hypersonic stability of reference EEV design in backwards orientation. Assess geometric options to improve hypersonic reorientation capability. EEV reorientation capability provides risk mitigation for incorrect entry attitude (e.g., spin-eject failure).
Numerical SimulationFree molecular solutions predict baseline geometry is unstablebackwardsContinuum solutions predict baseline geometry is stablebackwards
Mach-6 CF4 Test DataTests completed. Preliminary results show the EEV is marginally stable backwards which supports the LaRC CFD analyses
Mach-20 Helium TestsTests completed. Alternate configurations demonstrated to possess improved reorientation capability.
Mach-6 CF4 EEV model mounted on sting and Schlieren photograph at α = 168 deg
RDB Aug 2005284
RDB Aug 2005285
Subsonic Aerodynamic Tests: 20-Ft Vertical Spin Tunnel
Ames Arc Jet Complex
• Testing of heat shield materials for planetary entry vehicles, planetary probes, or hypersonic flight vehicles in relevant aerothermodynamicconditions • 3 test bays contain operative Arc Jet units of differing configurations that are serviced by common support equipment. • Can deliver 75 MW for a 30 sec duration or 150 MW for a 15 sec duration.
RDB Aug 2005286
MSR EEV Forebody TPS Material Screening Arc Jet Test
• Five material candidates identified from Market Survey
Carbon Phenolic (FMI, Edler Industries)Genesis heritage material (LMA)PICA (FMI)3CF (FMI)Dual Layer (Textron Systems)
Facility: 60 MW NASA ARC Interactive Heating FacilityTest Objective: Evaluate ablation characteristics and thermal performance of candidate materials - results will aid in the material selection of the EEV forebody heat shield
PICA model in arc jet during test
RDB Aug 2005287
Typical TPS Qualification Matrix: Genesis SRC
Reference: Willcockson, W.H.; “Genesis Recovery System Design Review, February 2004.
RDB Aug 2005288
Parachute Testing In Relevant ρ and Q
LaRC Transonic Dynamics Tunnel
Ames 80 x 120 ft Tunnel
RDB Aug 2005289
MER Parachute Load Qualification Testing
Wind tunnel testFull-scale parachute
NASA Ames 80 x 120 ft Tunnel
RDB Aug 2005290
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.Video courtesy of Pioneer Aerospace
Historical Experience & Case Studies
RDB Aug 2005291
Entry System Design• Entry system design constraints vary widely from mission to mission• Driving constraints derived from both mission and flight systems• Typical requirements from the mission system:
– Entry velocity• TPS material selection & thickness
– Entry angle• TPS thickness, structure load cases
– Upper atmospheric density• Forebody: entry system hypersonic drag coefficient
– Lower atmospheric density & surface elevation• Forebody: entry system supersonic/transonic drag vs. stability• Aftbody: supersonic stability
– Atmospheric winds• Parachute drag vs stability
– Expected surface environment• Landing system (hazard tolerance) or hazard avoidance system
– Terminal footprint accuracy (science)• Navigation approach, need for L/D and atmospheric guidance• Parachute deployment algorithm
RDB Aug 2005292
Entry System Design• Typical requirements from the flight system:
– Landed mass• Sizes most entry system components• May requires additional systems
– Launch Vehicle• Maximum diameter (ballistic coefficient)
– Instrumentation• May require heatshield penetrations or windows
– Spacecraft/Cruise-Stage Separation• Attitude control during coast phase: passive roll rate or active ACS• Backshell geometry: hypersonic reorientation
– Telecom• Critical events fault reconstruction
RDB Aug 2005293
Example Set of Entry System Requirements: Genesis SRC
Reference: Willcockson, W.H.; “Genesis Recovery System Design Review, February 2004.
RDB Aug 2005294
Historical Experience – Case Studies• Mars Landers
– Viking– MPF– MPL– MER– Phoenix
RDB Aug 2005295
Viking Lander Design Features• Two landers, separate L/V
– System redundancy, mitigates large unknowns of first planetary landing– 3.5 m maximum diameter
• 70-degree sphere-cone– Chosen to maximize hypersonic drag coefficient (mitigate risk of low Mars
atmospheric density)– Maintains marginal supersonic stability (active RCS during entry an
additional mitigation)• 11-deg angle of attack (L/D = 0.18)
– Minimize real-gas uncertainty (aerodynamic coefficients in Earth and Mars atmospheres crossed at this orientation)
– Allowed aerodynamic confidence with existing wind-tunnel facilities (before advent of CFD).
– Large qualification test program employing Nation’s wind-tunnel facilities.• SLA-561V
– Mass-efficient thermal protection material. Specifically designed to low Mars heat rate, with significant margin (factor of 10). Large qualification test program employing Nation’s arc-jet facilities.
RDB Aug 2005296
Viking Lander Design Features• Disk-gap band parachute
– Viking design emphasized drag performance over stability. Qualification included large test program, including full-scale supersonic deployment tests in relevant environment
• Terminal Descent Propulsion – Throttleable hydrazine system design with significant redundancy and
reliability• Use of Lift During Entry
– Modify trajectory to better balance environmental loads (structural margin) and parachute deployment altitude (timeline margin)
– Did not choose to perform bank-angle modulation during entry to reduce landed footprint
Viking’s success became the basis for all future planetary landersand provides and useful benchmark from which to assess risk
RDB Aug 2005297
Viking LandersLanding Dates: July 20, 1976 (VL1)
Sept. 3, 1976 (VL2)Launch Mass: 3530 kgLanded Mass: 600 kg
Target Coordinates: 22°N, 48°W (Chryse, VL1)44°N, 226°W (Utopia, VL2)
Landing Ellipse: 100 x 300 km (3σ)Miss Distance: 20 km (VL1), ?? (VL2)MOLA Surface Elevation: -3.627 km (VL1), -4.505 km (VL2)Local Time at Landing: Afternoon (VL1), Morning (VL2)EDL Data Return: Orbiter Relay
RDB Aug 2005298
Viking Entry SystemEntry Mass: 980 kgAeroshell Diameter: 3.54 mBallistic Coefficient: 63 kg/m2
Aeroshell Forebody Shape: 70° Sphere-ConeTPS Material: SLA-561VTPS Stagnation Point Thickness: 1.38 cmCG Location (XCG/D): 0.219CG offset: 2.5 cm
Inertial Entry Velocity: 4.61 km/sInertial Entry Angle: -17.0 deg at 244 kmTrim Angle of Attack: -11.1 degEntry System L/D: 0.18Peak Heating Rate: 21 W/cm2
Total Integrated Heating: 1100 J/cm2
RDB Aug 2005299
Viking EDL Sequence of Events
Entry Altitude: 244 kmParachute Deployment Altitude: 6 kmParachute Deployment Velocity: 250 m/sParachute Diameter: 15 mHeat Shield Release: Chute Deploy + 7 sTime on Parachute: 60 sTerminal Velocity on Parachute: 60 m/sDescent Engine Ignition Altitude: 1.5 kmDescent Engine Burn Duration: 40 sTerminal Descent Propellant: Purified HydrazineRetrorocket Design: 3 Throttleable Engines (18 nozzles ea.)Impact Velocity: 2 m/sLanding System: 3 Lander Legs, 22 cm Ground Clearance
RDB Aug 2005300
Mars Pathfinder EDL Design Features• One lander, launched on a Delta II 7925
– Cost (about 1/10 cost of Viking project*)– First direct entry at Mars– 2.65 m maximum diameter– Entry allocation mass of 603 kg
• Maintain Viking heritage where possible (reliability, cost and schedule)– 70-degree sphere-cone geometry– SLA-561V thermal protection system– Disk-gap band parachute
• Entry modifications– 2 rpm roll rate sufficient to provide inertial stability during coast phase– Higher entry velocity required additional TPS thickness and qualification
• Terminal descent and landing modifications– Sacrificed 30% in parachute drag for greater stability (90% larger relative
band area)– Airbag landing system (early studies indicate lower mass solution)– RAD rockets (augmentation added as landed mass increased)
*Recall that Viking project developed and successfully operated 2 landers and 2 orbiters
RDB Aug 2005301
Mars Pathfinder
Landing Dates: July 4, 1997Launch Mass: 894 kgLanded Mass: 370 kg
Target Coordinates: 19.5°N, 32.8°W (Chryse)Landing Ellipse: 100 x 300 km (3σ)Miss Distance: 23 kmMOLA Surface Elevation: -3.682 kmLocal Time at Landing: Early Morning (4 AM)EDL Data Return: DTE X-Band Semaphores
RDB Aug 2005302
Pathfinder Entry SystemEntry Mass: 585 kgAeroshell Diameter: 2.65 mBallistic Coefficient: 63 kg/m2
Aeroshell Forebody Shape: 70° Sphere-ConeTPS Material: SLA-561V (Forebody)
SLA-561S (Aftbody)TPS Stagnation Point Thickness: 1.9 cmCG Location (XCG/D): 0.27
Inertial Entry Velocity: 7.26 km/sInertial Entry Angle: -14.06 deg at 125 kmTrim Angle of Attack: 0 degEntry System L/D: 0Peak Heating Rate: 106 W/cm2
Total Integrated Heating: 3865 J/cm2
RDB Aug 2005303
Pathfinder EDL Sequence of Events
Entry Altitude: 125 kmParachute Deployment Altitude: 8 kmParachute Deployment Velocity: 386 m/sParachute Diameter: 12.7 mHeat Shield Release: Chute Deploy + 20 sTime on Parachute: 128 sTerminal Velocity on Parachute: 63 m/sRAD Ignition Altitude: 88 mRAD Burn Duration: 2.2 sRAD Propellant: HTPBRetrorocket Design: 3 Solid Rockets on BackshellImpact Velocity: 14.7 m/sLanding System: 4 Airbags, 6 Lobes per Bag (Vectran)
RDB Aug 2005304
Mars Pathfinder Mass Growth
Data courtesy of M. TauberRDB Aug 2005
305
Mars Pathfinder EDL System MEL
Subsystem Flight Mass (kg) Flight Mass (%)Forebody heatshield 73.9
94.017.57.01.430.7
104.0EDL Subtotal 328.5 56.1
256.8585.3
22.5Backshell 28.6Parachute subsystem 5.320m Bridle 2.1Radar altimeter 0.4RAD subsystem 9.3Airbag subsystem 31.7
Landed payload mass 43.9Entry Mass
RDB Aug 2005306
MPL EDL Design Features• One lander, launched on a Delta II 7425
– Cost (roughly half cost of MPF)– High surface elevation landing site– 2.4 m maximum diameter– Entry allocation mass of 505 kg
• Maintain Viking and MPF heritage where possible (reliability, cost and schedule)– 70-degree sphere-cone geometry– SLA-561V thermal protection system– Disk-gap band parachute (MPF flight spare)– Viking-like terminal descent system
• Shallow entry flight path angle– Reduce loads and structural mass– Higher heat load with thinner TPS required additional qualification
• Terminal descent and landing modifications– Pulsed-control hydrazine system
RDB Aug 2005307
Mars Polar Lander
Landing Dates: Dec. 3, 1999Launch Mass: 618 kgLanded Mass: 290 kg
Target Coordinates: 76°S, 195°W (Polar Layered Terrain)MOLA Surface Elevation: +2.3 kmLanding Ellipse: 200 x 20 km (3σ)Miss Distance: UnknownLocal Time at Landing: Early Morning (5 AM)EDL Data Return: None
RDB Aug 2005308
MPL Entry System
Entry Mass: 494 kgAeroshell Diameter: 2.4 mBallistic Coefficient: 60 kg/m2
Forebody Shape: 70° Sphere-Cone
TPS Material: SLA-561V (Forebody)SLA-561S (Backshell)
CG Location (XCG/D): ??Inertial Entry Velocity: 6.9 km/sInertial Entry Angle: -13.25 deg at 125 kmTrim Angle of Attack: 0 degEntry System L/D: 0Peak Heating Rate: 80 W/cm2
Total Integrated Heating: 4322 J/cm2
RDB Aug 2005309
MPL EDL Sequence of Events
Entry Altitude: 125 kmParachute Deployment Altitude: 8.8 kmParachute Deployment Velocity: 493 m/sParachute Diameter: 12.7 mHeat Shield Release: Chute Deploy + 10 sTime on Parachute: 85 sTerminal Velocity on Parachute: 80 m/sDescent Engine Ignition Altitude: 1300 mDescent Engine Burn Duration: 40 sTerminal Descent Propellant: HydrazineRetrorocket Design: 12 Pulse-Modulated 266 N EnginesImpact Velocity: 2.4 m/sLanding System: 3 Landing Legs
RDB Aug 2005310
MER EDL Design Features
RDB Aug 2005311
• Two landers, launched on separate Delta II 7925 and Delta 7925H– System redundancy, mitigates large risk of Mars landing– 2.65 m maximum diameter– Entry allocation mass increased to 850 kg
• Initially formulated as a MPF build-to-print mission (schedule, reliability, and cost). Much of the planned BTP design lost as a result of landed mass growth during development.– 70-degree sphere-cone geometry– 2 rpm roll rate sufficient to provide inertial stability during coast phase– SLA-561V thermal protection system (thickness reduced)– Disk-gap band parachute (small changes to MPF chute, increased deploy Q)– RAD rockets and airbags (increased capability)– Descent rate limiter (reduced mass design, new materials)– Shallow entry flight path angle and baseline of ∆DOR navigation data type
resulted in greatly reduced loads, allowing structural mass savings• Additional landing risk augmentation
– TIRS system (small solid rockets, backshell IMU, descent imager and software) added to reduce horizontal velocity at surface impact. System added at CDR due to concerns over unknown near-surface winds and potential susceptibility of airbags to this failure mode
• Modifications required significant EDL qualification test program
MER Terminal Descent System DescriptionParachute: Disc-gap-band parachute used to slow theSystem from supersonic speeds to a terminal descent
velocity between 60 to 80 m/sec
Backshell: Structure that contains the lander during entry. Also supports the RAD Motors, and TIRS Rockets and BIMUBackshell IMU (BIMU):
Gyro and accelerometer packageLitton: LN-200SN
(Mounted inside backshell)
RDB Aug 2005312
Radar Altimeter (RAS): Measures Altitude, used to determine vertical velocityHoneywell <insert part number>Max altitude 2400 m (8000 ft)
Descent Imager: 45 degree FOV frame transfer imagerUsed to acquire images used to determine horizontal velocity
Rover IMU (RIMU):Gyro and accelerometer package
Litton: LN-200SN(Located inside rover)
Lander: Structure that contains the roverand all of the landing support equipment
(airbag system, EDL batteries, righting equipment)
Bridle: Structurally and electrically connects the Lander to the Backshell during terminal descent.
Retro-Rockets (RAD Motors):Three (3) Retro-rockets used to slow the
descent of the system just prior to landingSingle rocket thrust:
Single rocket Impulse:
Transverse Impulse Rockets:Three rockets pointed through the CG of backshell
Used to change orientation of backshell during RAD Motor firing
Single rocket thrust:Single rocket Impulse:
DRL: Descent Rate Limiter inside lander petalcontrols separation velocity after lander separation
Mass Growth Throughout MER Lifecycle
RDB Aug 2005313Months from LaunchMonths from Launch
Entry Mass(kg)
Entry Mass(kg)
-40 -36 -32 -28 -24 -20 -16 -12 -8 -4 0600
625
650
675
700
725
750
775
800
825
850
Preliminary Design Review
Preliminary Design Review
Critical Design Review
Critical Design Review
Mars Exploration Rovers
Landing Dates: Jan. 3, 2004 (Spirit)Jan. 24, 2004 (Opportunity)
Launch Mass: 1062 kgLanded Mass: 550 kg
RDB Aug 2005314
Target Coordinates: 15°S, 175°W (Gusev Crater, Spirit)2°S, 354°W (Meridiani, Opportunity)
Landing Ellipse: 80 x 20 km (3σ)Miss Distance: 9 km (Spirit)
15 km (Opportunity)MOLA Surface Elevation: -1.91 km (Spirit)
-1.44 km (Opportunity)Local Time at Landing: Afternoon (1-2 PM)EDL Data Return: DTE X-Band Semaphores and UHF relay
MER Entry SystemEntry Mass: 832 kgAeroshell Diameter: 2.65 mBallistic Coefficient: 89 kg/m2
Aeroshell Forebody Shape: 70° Sphere-ConeTPS Material: SLA-561V (Forebody)
SLA-561S (Backshell)TPS Stagnation Point Thickness: 1.57 cmCG Location (XCG/D): 0.27
Inertial Entry Velocity: 5.7 km/sInertial Entry Angle: -11.5 deg at 125 kmTrim Angle of Attack: 0 degEntry System L/D: 0Peak Heating Rate: 41 W/cm2
Total Integrated Heating: 3687 J/cm2
RDB Aug 2005315
MER EDL Sequence of Events
RDB Aug 2005316
Entry Altitude: 125 kmParachute Deployment Altitude: 9.5 kmParachute Deployment Velocity: 430 m/sParachute Diameter: 14.1 mHeat Shield Release: Chute Deploy + 20 sTime on Parachute: 122 sTerminal Velocity on Parachute: 68 m/sRAD Ignition Altitude: 150 mRAD Burn Duration: 2.8 sRAD Propellant: HTPBRetrorocket Design: 3 Solid Rockets on BackshellImpact Velocity: 8 m/s vertical, 11.5 m/s horizontal (Spirit)
7 m/s vertical, 9 m/s horizontal (Opportunity)Landing System: 4 Airbags, 6 Lobes per Bag,
Dual Bladder (Vectran)
MER-B EDL System MELSubsystem Flight Mass (kg) Flight Mass (%)Forebody heatshield 89.6
106.124.05.62.657.0
TIRS 5.5 1.4DIMES 1.5 0.4
119.6EDL Subtotal Mass 411.5 49.4
420.75832.25
21.8Backshell 25.8Parachute subsystem 5.820m Bridle 1.4Radar altimeter 0.6RAD subsystem 13.9
Airbag subsystem 29.1
Landed payload mass 50.6Entry Mass
RDB Aug 2005317
Viking, Pathfinder and MER Comparison
RDB Aug 2005318
Viking I, II Mars Pathfinder MER
Forebody geometry, deg 70 70 70Aftbody geometry, deg 39/62 (biconic) 49 49Relative Entry Velocity, km/s 4.5, 4.42 7.6 5.7Relative Entry FPA, deg -17.6* -13.8 -11.5Mass, kg 980 585 840m/(CDA), kg/m2 63.7 62.3 89.8 XCG/D: reference 0.22 0.27 0.27Nominal α, deg -11.1 0 0L/D 0.18 0 0G&C 3-axis (active) spin stabilized spin stabilized
MER, MPF and Viking Entry Comparison
RDB Aug 2005319
Comparison of MER and MPF Entry Conditions
Parameter MER-A MER-B MPFArrival Date 4 Jan 2004 25 Jan 2004 4 Jul 1997Arrival Season Mid Winter Late Winter Summer Inertial Velocity (at 125 km) 5.65 km/s 5.72 km/s 7.26 km/sEntry Direction Posigrade Posigrade RetrogradeLocal Landing Time Afternoon Noon Pre-DawnEntry mass 827 kg 832 kg 585 kgLanded mass 540 kg 540 kg 410 kgLanding site altitude –1.91 km –1.44 km –3.68 km
• MER entry mass, local time, and landing site altitude significantly increase the EDL challenge relative to MPF
RDB Aug 2005320
Comparison of MPF and MER Terminal Descent
MER MPF Terminal Velocity Change
Descent Mass (kg) 740 530 +20%
Atmospheric Density Mid-afternoon
Pre-dawn
+21%
Landing Site Altitude (km) -1.3 -2.6* +3*%Chute Drag Area (m2) 67 52.5 -13%
Upper-bound terminal descent velocity (m/s)
85 65 +32%
RDB Aug 2005321
*Reference: Stelzner, Desai, Lee and Bruno, “The MER EDL and the Use of Aerodynamic Decelerators,” AIAA 03-2125, May 2003.
Parachute System Heritage
VikingD0 = 53.0 ft
Mars PathfinderD0 = 41.8 ft
MERD0 = 46.3 ft
CD = 0.63 CD = 0.41 CD = 0.41
RDB Aug 2005322
Reference: Cruz, J.R., “MER Parachute Decelerator Subsystem, May 15, 2003.
Phoenix EDL Design Features• One lander, launched on a Delta II 7925H
– Inherited MSP’01 lander (MPL BTP) hardware• 2.4 m maximum diameter• Entry allocation mass of 550 kg
– Large launch mass margin– New suite of science instruments– Funding profile requires focused Phase B effort
• Minimize lander modification (reliability, cost and schedule)– 70-degree sphere-cone geometry– SLA-561V thermal protection system– Disk-gap band parachute (Viking derivative to enable higher surface elevation
sites, MSP’01 requirement)– MPL-like terminal descent system (additional testing underway)
• Precision landing and hazard avoidance technology demonstration proposed in Phase A but descoped prior to project confirmation.– Planned but not required by science (descopeable)– MSP’01 Lander has cg offset designed to produce entry L/D of 0.06 and
active RCS protruding through backshell – lift-up controlled entry• Terminal descent and landing system
– Pulsed-control hydrazine system
RDB Aug 2005323
Phoenix
Landing Date: May 25, 2008Launch Mass: 705 kgLanded Mass: 364 kgTarget Coordinates: 70°N, 130°W Landing Ellipse: 200 x 40 km (3σ) (unguided)
20 km radius (guided)Local Time at Landing: Late Afternoon (6 PM)EDL Data Return: DTE X-Band, UHF to Orbiter
RDB Aug 2005324
Phoenix Entry System
Entry Mass: 538 kgAeroshell Diameter: 2.65 mBallistic Coefficient: 64 kg/m2
Forebody Shape: 70° Sphere-Cone
TPS Material: SLA-561V (Forebody)SLA-561S (Backshell)
CG Location (XCG/D): ??Inertial Entry Velocity: 5.79 km/sInertial Entry Angle: -12.5 deg at 125 kmTrim Angle of Attack: 3.5 degEntry System L/D: 0.06Peak Heating Rate: 47 W/cm2
Total Integrated Heating: 2827 J/cm2
RDB Aug 2005325
Heatshield1.25” - 5056/F40-.014-2.1 Slotted Flexcore
.014” T300/BTCY-1 FS3” - 5056/F40-.014-2.1 Slotted Flexcore Outer Ring
EX1541 Core Fill.020” M55J/EX1515 Outer Ring Doubler
0.55” - SLA561V TPS
Parachute/Mortar Support Ring
2219-T851 Aluminum
Heatshield/Backshell
Sep Fittings (6x)2219-T851 Aluminum
Parachute Thrust Cone3 Section 6061-T6 Aluminum
3 Longitudinal Riveted Doublers0.30” - SLA561S TPS
BS/LanderBipods (3x)
1” ID - 0.1” - M55J/EX1515
BS/CruiseSep Fitting (6x)2219-T851 Aluminum
T0 Connector
Backshell0.5” - 5056/F40-.014-2.1 Slotted Flexcore
.028” T300/BTCY-1 FSEX1541 Core Fill
.060” M55J Outer Ring Doublers0.20” - SLA561s TPS
Parachute Canister2219-T851 Aluminum
Ballast Location (2x)
Separation Connector (2x)
Backshell/CS I/F Ring2219-T851 Aluminum
Phoenix Aeroshell Structural Design
RDB Aug 2005326
Phoenix EDL Sequence of Events
Entry Altitude: 125 kmParachute Deployment Altitude: 10.2 kmParachute Deployment Velocity: 366 m/sParachute Diameter: 12.4 mHeat Shield Release: Chute Deploy + 10 sTime on Parachute: 187 sTerminal Velocity on Parachute: 41 m/sDescent Engine Ignition Altitude: 220 mDescent Engine Burn Duration: 12.6 sTerminal Descent Propellant: HydrazineRetrorocket Design: 12 Pulse-Modulated 266 N EnginesImpact Velocity: 2.4 m/sLanding System: 3 Landing Legs
RDB Aug 2005327
Historical Experience – Case Studies• Entry Probes
– Pioneer Venus– Galileo– MSR EEV– DS-2– Huygens– Stardust– Genesis
RDB Aug 2005328
Pioneer-Venus Probe Design Features• Designed and operated by the NASA Ames Research Center• The Pioneer Venus mission consisted of two components launched
separately: an Orbiter and a Multi-probe spacecraft– Orbiter (517 kg), carrying 17 instruments, was launched on May 20, 1978
(Atlas-Centaur)– Multi-probe (875 kg), carrying one large (315 kg) and 3 small (91 kg) probes,
was launched on May 20, 1978 (Atlas-Centaur)• Large probe release: Nov 16, 1978 (E-25 days)
– 7 science instruments within a sealed spherical pressure vessel (73 cm dia.)– Parachute system and forebody TPS separation system
• Small probe release: Nov 20, 1978 (E-21 days)– 5 science instruments within a sealed spherical pressure vessel– No deployables– One probe (day probe) transmitted for over an hour on the surface
• All probes were 45-deg sphere cones with entries on Dec 9, 1978• The multi-probe cruise-stage (290 kg) carried 2 instruments into the upper portion
of the atmosphere before being destroyed• The probes sounded the clouds and lower atmosphere at 4 separate locations
returning chemical, physical, and meteorological data on the Venus atmosphere.
RDB Aug 2005329
Pioneer-Venus Probe Mission Summary
Large Probe
North Probe
Day Probe
Night Probe
Cruise Stage
Entry Time (200 km) 18:45:32 18:49:40 18:52:18 18:56:13 20:21:52
Impact Time 19:39:53 19:42:40 19:47:59 19:52:05 -
Loss of Signal 19:39:53 19:42:40 20:55:34 19:52:07 20:22:55
Impact Latitude 4.4 N 59.3 N 31.3 S 28.7 S (37.9 S)
Impact Longitude 304.0 4.9 317.0 56.7 (290.9)
Solar Zenith Angle 65.7 108.0 79.9 150.7 60.7
Local Venus Time 7:38 3:35 6:46 0:07 8:30
• All times in UT (= EST + 5 hours) on December 9, 1978• Cruise-stage signal lost at 110 km altitude
RDB Aug 2005330
Pioneer Venus, Small North Probe
Entry Date: Dec. 9, 1978Entry Mass: 91 kgEntry Latitude: 60°NEntry Angle: -68.7° at 200 km AltitudeInertial Entry Velocity: 11.54 km/sAeroshell Diameter: 0.76 mBallistic Coefficient: 190 kg/m2
Geometry: 45° sphere cone with hemispherical afterbody
TPS Material: Carbon PhenolicTPS Stagnation Point Thickness: 1.2 cmCG Location (XCG/D): 0.4Peak Heating Rate: 7,200 W/cm2
Total Integrated Heating: 11,700 J/cm2
RDB Aug 2005331
Pioneer Venus, Small Night Probe
Entry Date: Dec. 9, 1978Entry Mass: 91 kgEntry Latitude: 30°NEntry Angle: -41.5° at 200 km AltitudeInertial Entry Velocity: 11.54 km/sAeroshell Diameter: 0.76 mBallistic Coefficient: 190 kg/m2
Geometry: 45° sphere cone with hemispherical afterbody
TPS Material: Carbon PhenolicTPS Stagnation Point Thickness: 1.2 cmCG Location (XCG/D): 0.4Peak Heating Rate: 5,500 W/cm2
Total Integrated Heating: 12,500 J/cm2
RDB Aug 2005332
Pioneer Venus, Small Day Probe
Entry Date: Dec. 9, 1978Entry Mass: 91 kgEntry Latitude: 34° SEntry Angle: -25.4° at 200 km AltitudeInertial Entry Velocity: 11.54 km/sAeroshell Diameter: 0.76 mBallistic Coefficient: 190 kg/m2
Geometry: 45° sphere cone with hemispherical afterbody
TPS Material: Carbon PhenolicTPS Stagnation Point Thickness: 1.2 cmCG Location (XCG/D): 0.4Peak Heating Rate: 3,900 W/cm2
Total Integrated Heating: 14,000 J/cm2
RDB Aug 2005333
Pioneer Venus, Large Probe
Entry Date: Dec. 9, 1978Entry Mass: 315 kgEntry Latitude: 60°NEntry Angle: -32.4° at 200 km AltitudeInertial Entry Velocity: 11.54 km/sAeroshell Diameter: 1.42 mBallistic Coefficient: 188 kg/m2
Geometry: 45° sphere cone with biconic afterbody
TPS Material: Carbon PhenolicTPS Stagnation Point Thickness: 1.6 cmCG Location (XCG/D): 0.4Peak Heating Rate: 4,500 W/cm2
Total Integrated Heating: 12,400 J/cm2
RDB Aug 2005334
Galileo Probe Design Features
• Highest speed entry of all time• 60 km/s• Approximately 50 times the peak heat rate of an Apollo capsule• 50% TPS mass fraction
• Probe designed and operated by the NASA Ames• 45 deg sphere-cone for stability
• Carbon-phenolic forebody TPS• Dual-chute system
• Science instruments observe atmospheric structure, winds, and atmospheric composition
RDB Aug 2005335
Galileo Probe
RDB Aug 2005336
Launch Date: October 18, 1989Release Date: July 13, 1995 (E-5 months)Entry Date: December 7, 1995Data Collection: 59 minutes (3.5 Mb)Entry Mass: 335 kgEntry Latitude: ??Entry Angle: -6.64° at 450 km AltitudeInertial Entry Velocity: 59.92 km/sAeroshell Diameter: 1.26 mBallistic Coefficient: 256 kg/m2
Forebody Shape: Blunt-Nosed 45° ConeTPS Material: Carbon PhenolicTPS Stagnation Point Thickness: 14.6 cmCG Location (XCG/D): 0.344Peak Heating Rate: 17,000 W/cm2
Total Integrated Heating: 200,000 J/cm2
MSR EEV Design Features• Containment assurance: risk-based design
– High reliability “fail-safe” entry system– Functional redundancy – Large margins– Simple system that minimizes the number of failure modes– Probabilistic risk assessment incorporated in design cycle
• Simple, passive ballistic entry system– Low ballistic coefficient– 60 deg sphere-cone to balance stability and drag– Hemispherical aftbody to mitigate off-nominal entry attitude– No deployments– Energy absorption system to ensure sample containment at landing– Carbon-phenolic forebody TPS with carbon-carbon structure– Forward cg with 2 rpm roll rate; no aerodynamic instabilities
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EEV Risk-Based Design, (CP5.7, Feb 2000)
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Preliminary: Work in ProgressMass: 45 kgDiameter: 0.9 m
Integrated Design-Test Development Program• Testing
– Subsonic aerodynamics testing in 20-Foot Vertical Spin Tunnel
– Over 200 static and dynamic crush tests of carbon foam and other candidate energy absorbing materials, machining approach to strength tailoring
– Acoustic testing of carbon foam
– PICA-15 crush tests
– Crush test of aluminum hemispheres
– Crush test of graphite epoxy sandwich structures
– Ground impact testing, UTTR, commercial clay, sod, sand, dirt
• Analysis– Computational aerodynamics
– Aerothermodynamics
– TPS sizing
– 3-DOF trajectory simulations, footprint determination
– I-DEAS solid-modeling (packaging, mass)
– Pre- and post-impact thermal
– NASTRAN structural analysis
– DYTRAN finite element modeling
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EEV Impact Dynamics Testing
Crush energy management drop test at IDRF, 3/99
Impact sphere crush tests at IDRF, 8/98-2/99
Graphite/Kevlar/Rhoacellarticle tested at IDRF 6/99
Ground characterizationtests at UTTR, 12/98
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Probabilistic Risk Assessment Scoping Study
Date: Nov-99 to May-00Objective: Evaluate viability of EEV design and development plan towards meeting draft 10-6 containment assurance requirement.Approach: Quantitatively evaluate risk and uncertainty, and prioritize the factors that contribute to risk.
PRA Task-1 completed on April 7, 2000Development path for demonstrating adherence to draft containment assurance requirement appears viable.Final review meeting at LaRC on May-11
EEV PRA realizationsCarbon-phenolic heatshield is required.Adequate EEV structure analysis and ground based test program in development plan.System validation flight test(s) are likely required.Physics-based verification of containment assurance seal(s) is vital.
PRA methodology/approach adopted by ProjectPRA effort across Project elements in workFollow-on EEV PRA analysis in workPhase II SAIC contract start expected in mid-June
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Preliminary: Work in ProgressPortion of EEV Fault Tree Analysis
EEV System Validation Flight Test
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• Perform system-level flight test at nominal conditions to verify EEV design performance and critical events that can not be fully duplicated with ground testing and analysis.
– Demonstrate TPS reliability and performance
– Demonstrate spin-eject orientation and aerodynamic stability
– Demonstrate structural integrity and seal performance of containment seals through entry and impact loads.
– Demonstrate tracking and recovery of EEV.
– Verify design performance is nominal and all critical risk parameters are within their design limits.
– Assess and reduce uncertainty of PRA quantification.• Successful SVFT will validate EEV performance
predictions and demonstrate no unknown system-level issues.
MSR EEV System MEL
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Subsystem Growth Mass (kg)
Growth Mass (%)
Mars sample 0.53.12.03.6
11.00.9
17.12.23.6
44.0
1.1OS, empty 7.0Containment vessel 4.5Impact sphere 8.2Body structure 25Lid structure 2.0Forebody TPS 39Aft & lid TPS 5.0Latches, bolts, vents, sensors, misc.
8.2
Total
DS-2 Mars Microprobe Design Features• Two surface penetrators housed in separate entry probes
– Piggback flight to Mars on MPL cruise stage– Extremely low cost (order $25M)– Mass and volume constraints imposed by MPL
• 0.35 m maximum diameter• Entry allocation mass of 3.75 kg
• Simple, passive entry system– 45-degree sphere-cone geometry (stability and impact speed)– Hemispherical afterbody (hypersonic reorientation capability)– No atmospheric deployments or separations– Mechanical spring s/c separation system
• New entry system developments– Entry system geometry– Low mass system with forward cg– SIRCA SPLIT thermal protection system– Design for impact
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Deep Space 2 (Mars Microprobes)
Entry Date: Dec. 3, 1999Entry Mass: 3.67 kgEntry Latitude: 61°SEntry Angle: -13.25° at 125 km AltitudeInertial Entry Velocity: 6.9 km/sAeroshell Diameter: 0.35 mBallistic Coefficient: 36.2 kg/m2
Forebody Shape: 45° sphere-cone with hemispherical aftbody
TPS Material: Sirca SPLITTPS Stagnation Point Thickness: 1 cmCG Location (XCG/D): 0.24Peak Heating Rate: 194 W/cm2
Total Integrated Heating: 8,712 J/cm2
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Deep Space 2 (Mars Microprobes)
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Huygens Probe Design Features
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• Designed and operated by the European Space Agency• 60 deg sphere-cone balances drag and stability requirements
• Tile-like forebody TPS• Multi-chute system
• Six instruments observe atmospheric structure, winds, atmospheric composition, surface imagery and spectroscopy, surface sounding during descent
Huygens Probe (Titan)
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Interplanetary Journey: 6.7 yearsSeparation Date: December 25, 2004 (E-22 days)Entry Date: January 14, 2005Entry Mass: 318 kgEntry Latitude: High latitude siteEntry Angle: -64° at 1250 km AltitudeInertial Entry Velocity: 6.2 km/sAeroshell Diameter: 2.7 mBallistic Coefficient: 35 kg/m2
Forebody Shape: 60° Sphere-ConeTPS Material: AQ60 Silica Fibers Reinforced
with Phenolic Resin (Aerospatiale)(approx. 79 kg forebody heatshield)
TPS Stagnation Point Thickness: ?? cmCG Location (XCG/D): ??Peak Heating Rate: 50 W/cm2
Total Integrated Heating: ?? J/cm2
Stardust Sample Return Capsule Design Features
• Discovery program mission – cost capped, competitive selection, PI-led• Simple entry system
– Low cost, passive, ballistic entry system– 60-deg forebody to balance drag and stability requirements– 0.8 m maximum diameter– No heatshield separation– No parachute deployment software (g-switch & timers)– 16 rpm roll rate during 4-hour coast
• Highest velocity Earth entry planned to date, 12.9 km/s– Combined with small nose radius, highest peak heat rate at Earth, 1200 W/cm2
– First flight of PICA forebody heatshield (single-piece low mass system)– SLA-561V aftbody TPS
• Development items– PICA – NASA Ames Research Center– UTTR landing site– Aft cg coupled with rarefied flow regime static instability, transonic dynamic
instability– Use of supersonic DGB drogue chute
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Stardust Sample Return Capsule
Entry Date: Jan. 15, 2006Entry Mass: 45.8 kgEntry Latitude: 40°NEntry Angle: -8.2° at 125 km AltitudeInertial Entry Velocity: 12.8 km/sAeroshell Diameter: 0.827 mBallistic Coefficient: 60 kg/m2
Forebody Shape: Blunt-Nosed 60° ConeTPS Material: PICA-15TPS Stagnation Point Thickness: 5.82 cmCG Location (XCG/D): 0.35Peak Heating Rate: 1,200 W/cm2
Total Integrated Heating: 36,000 J/cm2
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Stardust SRC
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Genesis Sample Return Capsule Design Features
• Discovery program mission – cost capped, competitive selection, PI-led• Simple entry system
– Low cost, passive, ballistic entry system– 11 km/sec entry velocity, 1.5 m maximum diameter– No heatshield separation– No parachute deployment software (improper installation of mechanical g-trigger
led to entry and descent without initiation of parachute deployment event)• Maximize Stardust heritage
– Forebody shape– UTTR landing site– Parachute deploy sequence– SLA-561V aftbody TPS– 16 rpm roll rate during 4-hour coast
• Development items– Carbon-carbon forebody with penetrations– Air-snatch SRC retrieval (planned)– Aft cg coupled with rarefied flow regime static instability, transonic dynamic
instability
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Genesis Sample Return Capsule
Entry Date: Sept. 8, 2004Entry Mass: 210 kgEntry Latitude: 40°NEntry Angle: -8° at 125 km AltitudeInertial Entry Velocity: 12.8 km/sAeroshell Diameter: 1.51 mBallistic Coefficient: 80 kg/m2
Forebody Shape: Blunt-Nosed 60° ConeTPS Material: Carbon-carbon (Forebody)
SLA-561V (Aftbody)TPS Stagnation Point Thickness: 6 cmCG Location (XCG/D): 0.33Peak Heating Rate: 700 W/cm2
Total Integrated Heating: 16,600 J/cm2
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Genesis SRC Overview
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DACSDrogue ChuteMain Parachute
Backshell TPS
Carbon-carbonHeatshield
DACS R&RMechanism
CanisterSupport Strut
SRC Hinge
ParachuteDeck
HeatshieldStructure
BackshellStructure
Canister
AvionicsDeck
Genesis SRC
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Common Entry Probe FeaturesMission Forebody
Cone Angle (deg)
AftbodyGeometry
ForebodyHeatshield
Material
Terminal Descent System
Pioneer-Venus
45 Hemispheric section
Carbon-phenolicCarbon-phenolic
SIRCAMSR EEV 60 Hemispheric
sectionCarbon-phenolic
None
Huygens 60 Hemispheric section
Silica Parachute
Stardust 60 Cone PICA Supersonic DGB & subsonic chute
Carbon-carbon
Parachute
Galileo 45 Hemispheric section
Parachute
DS-2 45 Hemisphere None
Genesis 60 Biconic Supersonic DGB & subsonic parafoil
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Common Features of Pioneer Venus Probes, Galileo Entry Probes and DS-2 Microprobes
• Each of these probes chose a 45-deg sphere-cone geometry for increased aerodynamic stability– Venus, Jupiter and Earth atmospheric densities provide sufficient
deceleration at moderate-high altitudes that the increased drag coefficient provided by larger cone angle is not required
– DS-2 Microprobes required increased stability and a lower drag coefficient to achieve proper Mars surface impact conditions
• High peak heat rate of Galileo and Pioneer-Venus probes necessitated used of carbon-phenolic heatshield material– Carbon-phenolic is a relatively high-mass material developed by the
DoD for use on ballistic missiles
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Common Features of Huygens Probe, MSR EEV, Stardust and Genesis Sample Return Capsules
• Each of these probes chose a 60-deg sphere-cone geometry to balance the conflicting requirements of aerodynamic drag and stability
• Three of these entry systems rely on a parachute during terminal descent• Genesis chose not to use PICA (in use for the higher entry heating of
Stardust) due to manufacturing issues and the Genesis structuralrequirements for forebody penetrations
• MSR EEV baselined carbon-phenolic due to reliability and performance margin of this material relative to lower mass systems
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Historical Experience – Case Studies• Aerobraking Spacecraft
– Magellan– MGS– Odyssey– MRO
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Magellan Aerobraking Design Features
• The Magellan spacecraft was not designed with aerobraking in mind. Following the primary surface mapping mission, aerobrakingwas used to reduce the ellipticity of the orbit, to obtain a high resolution global gravity map of Venus.
• Two solar panels and the High Gain Antenna provided the primary drag surfaces.
– Magellan used a “tail first” attitude, with the HGA trailing thespacecraft, for aerodynamic stability.
• Aerobraking corridor design allowed a 3σ atmospheric variability of 40%.
• Magellan used 2.8 kg of propellant during aerobraking for corridor control (controlling periapsis altitude), and 24.5 kg for attitude control.
– The large propellant usage was driven by thruster control of thes/c attitude following drag passes. Post-drag pass body rates were damped propulsively, costing propellant.
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Magellan (Venus)
Start of Aerobraking: May 25, 1993End of Aerobraking: August 3, 1993Aerobraking Duration: 70 daysAerobraking Revs: 730 revsInitial Orbit Period: 3.26 hrsFinal Orbit Period: 1.56 hrs
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Magellan Aerobraking ParametersSpacecraft Mass: 1,100 kgDrag Area: 23 m2
Ballistic Coefficient: 22 kg/m2
Initial Apoapsis Altitude: 8,470 kmFinal Apoapsis Altitude: 541 kmPeriapsis ∆V from Aerobraking: 1,208 m/sPeriapsis Altitude During Aerobraking: 135 - 141 kmAverage Periapsis Density: 8.3 kg/m3
Average Main Phase Periapsis Heating Rate: 0.3 W/cm2
Maximum Heating Rate: 0.4 W/cm2
Periapsis Density Variability (1σ): 5%Drag Pass Duration: 4 - 12 min
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Mars Global SurveyorAerobraking Design Features
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• Originally designed to complete aerobraking in 130 days and 405 orbits, MGS ultimately took 299 days and 891 orbits to complete aerobraking, due to a damper mechanism that failed upon solar panel deployment after launch.
• After 11 aerobraking orbits, it became clear that the damage to the solar panel was more serious than previously thought. During drag passes, the solar panel was bending beyond the fully closed (latched) position.
– Analysis and concurrent ground testing indicated that one of thesolar panel facesheets had been cracked when the undampedpanel was deployed.
• Aerobraking was halted for a month to replan the aerobraking phase.– Aerobraking resumed with a dynamic pressure profile that was
less than half the originally planned value.– Solar panels were swept back 30° for aerodynamic stability.– The unlatched panel was rotated 180°, putting the cell side of the
array into the flow so that aerodynamic forces would push the unlatched hinge toward the closed position.
Mars Global Surveyor
C.G.
+Y
+X
+Z
Aero Flow(Nominal)
C.P.
33.8° [30°] Sweep
30.5° [30°] Sweep
Solar Cells
Jammed Hinge
Solar Cells
Start of Aerobraking: September 16, 1997Science Phasing Orbit: March 27, 1998 - Sept. 9, 1998End of Aerobraking: February 4, 1998Aerobraking Duration: 299 daysAerobraking Revs: 891 revsInitial Orbit Period: 45 hrsFinal Orbit Period: 1.89 hrs
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MGS Aerobraking ParametersSpacecraft Mass: 760 kgDrag Area: 17 m2
Ballistic Coefficient: 22 kg/m2
Initial Apoapsis Altitude: 54,028 kmFinal Apoapsis Altitude: 453 kmPeriapsis ∆V from Aerobraking: 1,217 m/sPeriapsis Altitude During Aerobraking: 100 - 134 kmAverage Periapsis Density: 19.4 kg/m3
Average Main Phase Periapsis Heating Rate: 0.08 W/cm2
Maximum Heating Rate: 0.43 W/cm2
Periapsis Density Variability (1σ): 31%Drag Pass Duration: 6 - 37 minNumber of Aerobraking Trim Maneuvers: 92
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Mars Odyssey AerobrakingDesign Features & Cost
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• Single solar array was stowed for drag passes, with the spacecraft bus oriented in the direction of the flow.
• Many of the aerobraking processes, design tools and personnel from MGS were used in Odyssey aerobraking, providing a major benefit to the project.
– Odyssey performed aerobraking with no major anomalies, completing the phase several days earlier than planned.
• Odyssey demonstrated an on-board periapsis timing estimator, which computed the approximate periapsis time based on peak acceleration from one drag pass and applied a correction to the sequenced periapsis time for the next drag pass.
• The cost of developing and executing aerobraking for Mars Odyssey were estimated as $9.3M.
– Aerobraking development: $1.5M– Aerobraking operations: $4.8M– Science operations during aerobraking: $3.0M– DSN costs were not charged to the project and are not included
Mars Odyssey
VelocityNadir
Start of Aerobraking: October 27, 2001End of Aerobraking: January 11, 2002Aerobraking Duration: 76 daysAerobraking Revs: 332 revsInitial Orbit Period: 18.6 hrsFinal Orbit Period: 1.9 hrs
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Odyssey Aerobraking ParametersSpacecraft Mass: 461 kgDrag Area: 11 m2
Ballistic Coefficient: 19 kg/m2
Initial Apoapsis Altitude: 26,700 kmFinal Apoapsis Altitude: 503 kmPeriapsis ∆V from Aerobraking: 1,080 m/sPeriapsis Altitude During Aerobraking: 95 - 158 kmAverage Periapsis Density: 47.5 kg/m3
Average Main Phase Periapsis Heating Rate: 0.27 W/cm2
Maximum Heating Rate: 0.57 W/cm2
Periapsis Density Variability (1σ): 30%Drag Pass Duration: 5 - 21 minNumber of Aerobraking Trim Maneuvers: 33
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Mars Reconnaissance OrbiterAerobraking Design Features
• Two solar panels and HGA provide much larger drag area than previous aerobraking spacecraft.
• Planned long aerobraking duration (167 days) allows target heating rates to be less than half that flown on Odyssey.
– Lower susceptibility to atmospheric density variations (150% margin).
• HGA and solar arrays remain fixed (are not articulated) during aerobraking operations.
• Minimum orbit lifetime of 48 hours will be maintained during orbit walkout, to guard against communication outages or safe mode entries.
• Periapsis Timing Estimator to be improved for MRO, using acceldata to calculate ∆V accumulated during each drag pass, and computing estimated time of next periapsis.
• High downlink data rates available (500 kbps vs. 28.8 kbps for Odyssey).
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Mars Reconnaissance Orbiter
Direction of Flight
Nadir Nadir
View Along the Velocity Vector
Start of Aerobraking (Planned): March 16, 2006End of Aerobraking (Planned): August 30, 2006Aerobraking Duration (Planned): 167 daysAerobraking Revs (Planned): 495 revsInitial Orbit Period (Planned): 35 hrsFinal Orbit Period (Planned): 1.9 hrs
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MRO Aerobraking ParametersSpacecraft Mass: 1400 kgDrag Area: 37.7 m2
Ballistic Coefficient: 18 kg/m2
Initial Apoapsis Altitude: 44,000 kmFinal Apoapsis Altitude: 450 kmPeriapsis ∆V from Aerobraking: 1,193 m/sPeriapsis Altitude During Aerobraking: 107 - 150 kmAverage Main Phase Periapsis Heating Rate: 0.11 W/cm2
Maximum Expected Heating Rate: 0.16 W/cm2
Expected Periapsis Density Variability (1σ): 30%Number of Aerobraking Trim Maneuvers (Planned): 30
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Future Expectations for Robotic Exploration• Larger robotic entry systems
– 4-5 m diameter aeroshells• TPS material and aerothermodynamic complexity accommodated through
increased margins– Much larger attached and/or trailing inflatables (fabric & thin-film
concepts)• Nanoprobes
– Dozens of ballistic entry systems of order 1 kg enabling a network science mission. Mars, Venus and outer planet applications.
• Hypersonic aeromaneuvering– Lifting entry– Precision and pinpoint landing– Guidance initiation of parachute deployment– Aerocapture
• Terminal descent– Large diameter supersonic parachutes, possibly deploying at Mach 2.7– Ring-sail and other chute types for improved subsonic performance
• Propulsive terminal descent and soft landingRDB Aug 2005
374
Future Requirements for Human Exploration
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• Much larger entry systems– TPS material and aerothermodynamic complexity accommodated
through human-rated TPS qualification program– On-orbit assembly & certification of heatshield integrity and/or
mammoth launch vehicle– Dual-use heatshields for aerocapture followed by entry from orbit
• Higher L/D configurations– Required to reduce deceleration to acceptable levels– Need a configuration that gains lift without reducing drag
• Hypersonic aeromaneuvering– Pinpoint landing– Guidance initiation of parachute deployment– Aerocapture - advantages increase as trip time is reduced
• Terminal descent– Supersonic propulsive initiation– High Mach parachutes or other aerodynamic decelerator– Large or multiple subsonic parachutes
• Propulsive terminal descent and soft landing
A Few Aeroassist Contacts
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NASA LaRC NASA MSFC– Dick Powell - Bonnie James– Prasun Desai - Michelle Munk– Neil Cheatwood LMA– Juan Cruz - Bill Willcockson– Peter Gnoffo - Doug Gulick
NASA Ames Ball Aerospace– Raj Venkatapathy - Kevin Miller – Jim Arnold - Jim Mascarelli– Paul Wercinski Draper Labs– Dean Kontinos - Gregg Barton
NASA JSC Universities– Claude Graves - Bobby Braun (GA Tech)– Lee Bryant - Evans Lyne (U of Tennesee)
JPL - Bob Tolson (NCSU)– Bob Mitcheltree - Bob Blanchard (NIA)– Rob Manning - Ken Mease (UCSB)– Tom RIvellini - Graham Candler (U of Minn)– Dara Sabahi - Wallace Fowler (U of Texas)– Jeff Hall– Adam Steltzner
Classic Planetary Entry ReferencesTextbooks• Vinh, Busemann and Culp, Hypersonic and Planetary Entry Flight
Mechanics, 2nd edition, University of Michigan Press, 1980.• Anderson, John D., Hypersonic and High Temperature Gas Dynamics,
McGraw-Hill Book Company, 1989.• Martin, J.J., Atmospheric Entry, Prentice-Hall, 1966.• Loh, W.H.T., Re-entry and Planetary Entry Physics and Technology,
Volume I and II, Springer-Verlag, 1969.• Regan, Reentry Vehicle Aerodynamics, AIAA Education Series, 1984.Overview• Walberg, G.D., “A Survey of Aeroassisted Orbit Transfer,” Journal of
Spacecraft and Rockets, Vol. 22, No. 1, Jan-Feb 1985.Aerodynamics and Heating• Fay, J.A., and Riddell, F.R., “Theory of Stagnation-Point Heat Transfer in
Dissociated Air,” Journal of Aeronautical Science, Feb 1958.• Allen, H.J., Seiff, A., and Winovich, W., “Aerodynamic Heating of Conical
Entry Vehicles at Speeds in Excess of Earth Parabolic Speed,“ NASA TR R-185, Dec 1963.
• Marvin, J.G. and Deiwert, G.S., “Convective Heat Transfer in Planetary Gases,” NASA TR R-224, July 1965.
• Tauber, M.E., and Wakefield, R.M., “Heating Environment and Protectionduring Jupiter Entry,” AIAA Journal of Spacecraft & Rockets, Vol. 8, No. 6, June 1971.
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Classic Planetary Entry ReferencesFlight Mechanics• Allen, H.J. and Eggers, A.J., “A Study of the Motion and Aerodynamic Heating of
Missiles Entering the Earth’s Atmosphere at Supersonic Speeds,” NACA TN-4047, 1957.
• Chapman, D.R., “An Approximate Analytical Method for Studying Entry into Planetary Atmospheres,” NASA TR R-111, 1959.
• Chapman, D.R., “An Analysis if the Corridor and Guidance Requirements for Supercircular Entry into Planetary Atmsopheres,” NASA TR R-55, 1960.
• Citron, S.J., and Meir, T.C., “An Analytic Solution for Entry into Planetary Atmospheres,” AIAA Journal, March 1965, pp. 470-475.
• Loh, W.H.T., “Extension of 2nd Order Theory of Entry Mechanics to Oscillatory Entry Solutions, AIAA Journal, Sept. 1965, pp. 1688-1697.
Apollo/Shuttle• Curry, D.M., and Stephens, E.W., “Apollo Ablator Thermal Performance at
Superorbital Entry Velocities,” NASA TN D-5969, Sept. 1970.• Lee, D.B. and Goodrich W.D., “The Aerothermodynamic Environment of the Apollo
Command Module during Superorbital Entry,” NASA TN D-6792, Apr 1972.• Graves, C.A., and Harpold, J.C.; “Apollo Experience Report: Mission Planning for
Apollo Reentry,” NASA TN-D-6725, March 1972.• Harpold, J.C., and Graves, C.A.; “Shuttle Entry Guidance,” NASA TM-79949, 1979.
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Past Decade Mission/Flight System References• Mars Pathfinder atmospheric entry navigation operations; Braun, R D; Spencer, D A;
Kallemeyn, P H; Vaughan, R M.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 348-356, 1999.
• Mars Pathfinder entry, descent, and landing reconstruction; Spencer, D A; Blanchard, R C; Braun, R D; Kallemeyn, P H; Thurman, S W.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 357-366. May-June 1999.
• Mars Pathfinder Status at Launch; Spear, A.J.; Freeman, D.C., Jr.; Braun, R.D.; IAF Paper IAF-96-Q.3.02, 1996.Magellan aerobraking control corridor - design and implementation; Willcockson, W H. Advances in the Astronautical Sciences. 1994
• Aerobraking Magellan: plan versus reality; Lyons, D T.; Advances in the AstronauticalSciences, 1994.
• Determining Venusian upper atmosphere characteristics using Magellan attitude control data; Espiritu, R C; Tolson, R H,; Proceedings of the 5th AAS/AIAA Spaceflight Mechanics Conference, Albuquerque, NM; 13-16 Feb. 1995. pp. 377-393.
• Aerobraking at Mars: The MGS Mission; J. Beerer; R. Brooks; P. Esposito; D. Lyons; W. Sidney; H.L. Curtis; W. Willcockson; Journal of Spacecraft and Rockets, Jan. 1996.
• Mars Global Surveyor - Aerobraking with a broken wing; Lyons, D T.; Proceedings of the AAS/AIAA Astrodynamics Conference, Sun Valley, ID; Aug. 1997. pp. 275-294.
• The development and evaluation of an operational aerobraking strategy for the Mars2001 Odyssey Orbiter; Tartabini, P M; Munk, M M; Powell, R W.; AIAA/AAS Astrodynamics Specialist Conference and Exhibit, Monterey, CA, Aug. 2002.
• Approaches to autonomous aerobraking at Mars; Hanna, J L; Tolson, R H.; Advances in the Astronautical Sciences, 2002.
• Application of Accelerometer Data to Mars Odyssey Aerobraking and Atmospheric Modeling; Tolson, R H; Keating, G M; George, B E; Escalera, P E; Werner, M R; Dwyer, A M; Hanna, J L; NASA TM-20030002226.
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Past Decade Mission/Flight System References• Modeling reaction control system effects on Mars Odyssey; Hanna, J L; Chavis, Z Q;
Wilmoth, R G.; AIAA/AAS Astrodynamics Specialist Conference, Aug. 2002.• Mars Reconnaissance Orbiter - Aerobraking reference trajectory; Lyons, D T.;
AIAA/AAS Astrodynamics Specialist Conference and Exhibit, Monterey, CA, Aug. 2002.• Mars Microprobe entry-to-impact analysis; Braun, R D; Mitcheltree, R A; Cheatwood,
F M.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 412-420. May-June 1999.• Stardust Sample Return Capsule design experience; Willcockson, W H.; Journal of
Spacecraft and Rockets. Vol. 36, no. 3, pp. 470-474. May-June 1999.• Mars Polar Lander Entry, Descent and Landing design; Willcockson, W H.;
Proceedings of the AAS/AIAA Astrodynamics Conference, Aug. 1999. pp. 33-52. • Aerobraking at Venus and Mars - A comparison of the Magellan and Mars Global
Surveyor aerobraking phases; Lyons, D T; Proceedings of the AAS/AIAA Astrodynamics Conference, Aug. 1999. pp. 859-877.
• Entry, Descent and Landing Scenario for the Mars Exploration Rover Mission; Desai, P N; Lee, W J.; International Workshop on Planetary Probe Atmospheric Entry and Descent Trajectory Analyses and Sciences, October 2003, Lisbon, Portugal.
• Entry Trajectory and Atmospheric Reconstruction Methodologies for the Mars Exploration Rover Mission; Desai, P N; Blanchard, R.C.; Powell, R.W.; International Workshop on Planetary Probe Atmospheric Entry and Descent Trajectory Analyses and Sciences, October 2003, Lisbon, Portugal.
• Mission Design Overview for the Mars Exploration Rover Mission; Roncoli, R.B and Ludwinski, J.M.; AIAA Paper 2002-4823, August 2002.
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Past Decade Advanced Study References
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• Cost-Benefit Analysis of the Aerocapture Mission Set; Hall, J.L.; Noca, M.A.; Bailey, R.W.; Journal of Spacecraft & Rockets, Vol. 42, No. 2, 2005, pp. 309-320.
• Aeroassist technology planning for exploration; Munk, M M; Powell, R.W.; Proceedings of the AAS/AIAA Space Flight Mechanics Mtg, Jan. 2000. pp. 1073-1083.
• Developments in nanotechnology and implications for future atmospheric entry probes; Arnold, J.O.; Venkatapathy, E.; European Space Agency Special Publication, ESA SP, n 544, February, 2004, p 253-265.
• A passive Earth-entry capsule for Mars Sample Return; Mitcheltree, R A; Kellas, S; Dorsey, J T; Desai, P N; Martin, C J.; AIAA/ASME Joint Thermophysics and Heat Transfer Conference, 7th, Albuquerque, NM, June 15-18, 1998
• Sample returns missions in the coming decade; Desai, P N; Mitcheltree, R A; McNeil Cheatwood, International Astronautical Congress, 51st, Rio de Janeiro, Brazil, Oct. 2000.
• Flyby Delivers Multiple Deep Jupiter Probes; Spilker, T R; Hubbard, W B; Ingersoll, A P.; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020.
• Saturn Deep Atmospheric Entry Probes Delivered by INSIDE Jupiter Derivative Spacecraft; Spilker, T R.; Forum on Innovative Approaches to Outer Planetary Exploration 2001-2020.
• Earth Entry Vehicle for Mars Sample Return; Mitcheltree, R A; Braun, R D; Hughes S J; Simonsen, L C.; 51st International Astronautics Federation Congress, Rio de Janeiro, Brazil, Oct. 2000.
• Small Neptune orbiter using aerocapture; Lemmerman, L A; Wercinski, P F.; Space Technology & Applications International Forum - Conference on Future Science & Earth Science Missions, Jan. 1997. pp. 101-110.
• Low cost atmospheric probe missions to the outer planets; Wallace, R A; Rowley, R W; Wercinski, P F.; Space Technology & Applications International Forum - Conference on Future Science & Earth Science Missions, Jan. 1997. pp. 95-100.
• Uranus and Neptune atmospheric-entry probe study; Tauber, M.; Wercinski, P.; Henline, W.; Paterson, J.; Journal of Spacecraft and Rockets, v 31, n 5, Sept-Oct, 1994, p 789-805.
Past Decade Advanced Study References
RDB Aug 2005382
• Aerobraking technology for manned space transportation systems; ARNOLD, J O; TAUBER, M E; GOLDSTEIN, H E.; 43rd International Astronautical Congress, 28 Aug.-5 Sept. 1992.
• Manned Mars aerobrake vehicle design issues; Freeman, D C J; Powell, R W; Braun, R D.; Space Technology. Vol. 12, no. 3, pp. 313-334. 1992.
• Departure Energies, Trip Times and Entry Speeds for Human Mars Missions; Munk, M M; AAS Paper 99-103, 1999.
• Configurational analysis of the SHARP-L1 re-entry vehicle; Starkey, R P; Reuster, J G; Lewis, M J; Kolodziej, P.; 41st AIAA Aerospace Sciences Meeting & Exhibit, Jan. 2003.
• Design of aerogravity-assist trajectories; Johnson, W R; Longuski, J M.; Journal of Spacecraft and Rockets. Vol. 39, no. 1, pp. 23-30. Jan. 2001.
• Aerocapture simulation and performance for the Titan Explorer mission; Way, D W; Powell, R W; Edquist, K T; Masciarelli, J P; Starr, B R.; 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, Huntsville, AL, July 2003.
• Aerocapture Technology Development Needs for Outer Planet Exploration; Wercinski, P; Munk, M; Powell, R; Hall, J; Graves, C.; RECON no. 20020077966.
• Earth return aerocapture of the TransHab vehicle for a manned Mars mission; Muth, D; Lyne, J E.; Proceedings of the AAS/AIAA Space Flight Mechanics Meeting, Jan. 2000. pp. 1531-1538.
• High L/D Mars aerocapture for 2001, 2003 and 2005 mission opportunities; Jits, R Y; Walberg, G D.; AIAA 36th Aerospace Sciences Meeting & Exhibit, 36th, Jan. 1998.
• An overview of the aerocapture flight test experiment (AFTE); Hall, J L.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Monterey, CA, Aug. 2002.
• Second generation Mars landed missions; Graf, J; Thurman, S; Eisen, H; Rivellini, T; Sabahi, D.; 2001 IEEE Aerospace Conference, Big Sky, MT; Mar. 2001. pp. 243-254.
• Entry configurations and performance comparisons for the Mars Smart Lander; Lockwood, M K; Sutton, K; Prabhu, R; Powell, R; Graves, C; Epp, C; Carman, G.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Monterey, CA, Aug. 2002.
Past Decade Advanced Study References
RDB Aug 2005383
• Physiological constraints on deceleration during the aerocapture of manned vehicles; Lyne, J E; Journal of Spacecraft and Rockets. Vol. 31, no. 3, pp. 443-446. May-June 1994.
• Earth aerobraking strategies for manned return from Mars; Braun, R.D.; Powell, R.W.; Lyne, J. E.; Journal of Spacecraft and Rockets, Vol. 29, Jun 1992.
• Parametric study of manned aerocapture. Part I: Earth return from Mars; Lyne, J.Evans; Tauber, M.E.; Braun, R.D.; Journal of Spacecraft and Rockets, v 29, n 6, Nov-Dec, 1992, p 808-813
• Parametric study of manned aerocapture. II - Mars entry; LYNE, J E; ANAGNOST, A; TAUBER, M E.; Journal of Spacecraft and Rockets , vol. 29, no. 6, Nov.-Dec. 1992.
• Earth atmospheric entry studies for manned mars missions; Tauber, M.E.; Palmer, G.E.; Journal of Thermophysics and Heat Transfer, v 6, n 2, Apr-Jun, 1992, p 193-199
• Minimum Mars Mission Approach; Bryant, L.; Cockrell, B.; Condon, G.; Kennedy, K.; Lewis, S.; Masciarelli, J.; Munk, M.; Tigges, M.; Wilson, S.; Proceedings of the 4th International Conference on Engineering, Construction and Operations in Space, 1994, p 1309-1322.
• Aerobrake design studies for manned Mars missions; Tauber, M; Chargin, M; Henline, W; Hamm, K R J; Miura, H; Chiu, A; Yang, L.; Journal of Spacecraft and Rockets. Vol. 30, no. 6, pp. 656-664. Nov.-Dec. 1993
• Unmanned and manned Mars missions - Aeroassist technology needs and issues; WILLCOCKSON, W H.; 2nd International Conference on Solar System Exploration, Pasadena, CA; Aug. 1989.
• Low-cost entry systems for future planetary exploration missions; Rasky, D J; Tran, H K.; Acta Astronautica. Vol. 45, no. 4, pp. 347-355. 1999.
• Ultra-light entry systems for planetary missions; Murbach, M S; Kourtides, D; Chen, Y K.; AIAA 34th Aerospace Sciences Meeting and Exhibit, Jan. 1996.
• The Aeroassist Flight Experiment; WALBERG, G D; SIEMERS, PMIII; CALLOWAY, R L; JONES, J J.; 38th International Astronautical Congress, Brighton, England, Oct. 1987.
Past Decade Aerothermodynamic References• Computational aerothermodynamic design issues for hypersonic vehicles; Gnoffo,
P A; Weilmuenster, K J; Hamilton, H H; Olynick, D R; Venkatapathy, E.; Journal of Spacecraft and Rockets. Vol. 36, no. 1, pp. 21-43. Jan.-Feb. 1999.
• Computational aerothermodynamics in aeroassist applications; Gnoffo, P A; AIAA Computational Fluid Dynamics Conference, 15th, Anaheim, CA, June 2001.
• Planetary-entry gas dynamics; Gnoffo, P A.; Annual review of fluid mechanics. Vol. 31 (A99-35451 09-34), Palo Alto, CA, Annual Reviews, 1999, p. 459-494.
• Aerothermal environment for hypersonic vehicle design current practices andfuture requirements; Venkatapathy, E.; European Space Agency Special Publication ESA SP-426, Jan, 2000, p 239.
• Stagnation-point radiative heating relations for earth and Mars entries; TAUBER, M E; SUTTON, K.; Journal of Spacecraft and Rockets. Vol. 28, pp. 40-42. Jan.-Feb. 1991.
• Mars Pathfinder rarefied aerodynamics: Computations and measurements; Moss, J N; Blanchard, R C; Wilmoth, R G; Braun, R D.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 330-339. 1999.
• Prediction and validation of Mars Pathfinder hypersonic aerodynamic database; Gnoffo, P A; Braun, R D; Weilmuenster, K J; Mitcheltree, R A; Engelund, W C; Powell, R W.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 367-373. 1999
• Influence of sonic-line location on Mars Pathfinder probe aerothermodynamics; Gnoffo, P A; James, K; Weilmuenster, J; Braun, R D; Cruz, C I.; Journal of Spacecraft and Rockets. Vol. 33, no. 2, pp. 169-177. 1996
• Wake flow about the Mars Pathfinder entry vehicle; Mitcheltree, R A; Gnoffo, P A.; Journal of Spacecraft and Rockets. Vol. 32, no. 6, pp. 771-776. Sept.-Oct. 1995.
• Mars Pathfinder trajectory based heating and ablation calculations; Chen, Y K; Henline, W D; Tauber, M E.; Journal of Spacecraft and Rockets. Vol. 32, no. 2, pp. 225-230. Mar.-Apr. 1995.
RDB Aug 2005384
Past Decade Aerothermodynamic References• Aerodynamics of the Mars Microprobe entry vehicles; Mitcheltree, R A; Moss, J N;
Cheatwood, F M; Greene, F A; Braun, R D.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 392-398. 1999
• Aerothermal heating predictions for Mars Microprobe; Mitcheltree, R A; DiFulvio, M; Horvath, T J; Braun, R D.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 405-411. May-June 1999.
• Direct simulation Monte Carlo calculations of aerothermodynamics for Mars Microprobe capsules; Moss, J N; Wilmoth, R G; Price, J M.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 399-404. 1999.
• Subsonic static and dynamic aerodynamics of blunt entry vehicles; Mitcheltree, R A; Fremaux, C M; Yates, L A.; AIAA 37th Aerospace Sciences Meeting, Jan. 1999.
• Aerodynamics of Stardust Sample Return Capsule; Mitcheltree, R A; Wilmoth, R G; Cheatwood, F M; Brauckmann, G J; Greene, A.F.; AIAA 15th Applied Aerodynamics Conference, June 1997. pp. 697-707.
• Low-density aerodynamics of the Stardust Sample Return Capsule; Wilmoth, R G; Mitcheltree, R A; Moss, J N.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 436-441. May-June 1999.
• Transonic and low supersonic static and dynamic aerodynamic characteristics of the Stardust sample return capsule; Chapman, G T; Mitcheltree, R A; Hathaway, W H.; AIAA, Aerospace Sciences Meeting and Exhibit, 37th, Reno, NV, Jan. 1999.
• Aerothermodynamic analysis of Stardust Sample Return Capsule with coupled radiation and ablation; Gupta, R N.; Journal of Spacecraft and Rockets. Vol. 37, no. 4, pp. 507-514. July-Aug. 2000.
• Aerothermodynamics of the Stardust Sample Return Capsule; Olynick, D; Chen, Y K; Tauber, M E.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 442-461. May-June 1999.
• CFD code comparisons for Mars entry simulations; Papadopoulos, P; Prabhu, D; Olynick, D; Chen, Y K; Cheatwood, F M.; 36th Aerospace Sciences Meeting,Jan. 1998.
RDB Aug 2005385
Past Decade Aerothermodynamic References• DSMC simulations of blunt body flows for Mars entries - Mars Pathfinder and Mars
Microprobe capsules; Moss, J N; Wilmoth, R G; Price, J M.; 32nd AIAA ThermophysicsConference, June 1997.
• Magellan Aerodynamic Characteristics During the Termination Experiment Including Thruster Plume-Free Stream Interaction; Cestero, F J; Tolson, R H.; NASA TM-19980029679, 1998.
• Reentry-F Flowfield Solutions at 80,000 ft.; William A. Wood; Christopher J. Riley; McNeil Cheatwood; NASA TM-112856, May 1997.
• Galileo probe aerodynamics; Seiff, A; Venkatapathy, E; Haas, B; Intrieri, P.; AIAA 14th
Applied Aerodynamics Conference, 14th, New Orleans, LA, June 18-20, 1996.• Ballistic range and aerothermodynamic testing; Strawa, A.W.; Chapman, G.T.;
Arnold, J.O.; Canning, T.N.; Journal of Aircraft, 28, 443-449, July 1991.• Aerothermodynamic study of slender conical vehicles; Thompson, R.A.; Zoby, E.V.;
Wurster, K.E.; Gnoffo, P.A.; Journal of Thermophysics and Heat Transfer, v 3, n 4, p 361-367, Oct. 1989.
• Wake closure characteristics and afterbody heating on a Mars sample return orbiter; Horvath, T J; Cheatwood, F M; Wilmoth, R G; Alter, S J.; Space Technology and Applications International Forum - STAIF 2002; Albuquerque, NM; 3-6 Feb. 2002. pp. 318-336.
• Aerothermodynamic Environment Definition for the Genesis Sample ReturnCapsule; Cheatwood, F M N; Merski, N R J; Riley, C J; Mitcheltree, R A.; 35th AIAA Thermophysics Conference, Jun. 2001.
• Aerothermal Effects of Cavities and Protuberances for High-Speed Sample Return Capsules; Olynick, D; Kontinos, D.; 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, United States, Jan. 1999.
• Reassessment of effect of dust erosion on heatshield of Mars entry vehicle; Palmer, G; Chen, Y K; Papadopoulos, P; Tauber, M.; Journal of Spacecraft and Rockets. Vol. 37, no. 6, pp. 747-752. Nov.-Dec. 2000.
RDB Aug 2005386
Past Decade Aerothermodynamic References
RDB Aug 2005387
• Measuring lift coefficient in free molecular flow while aerobraking Magellan; Lyons, D T; Hulburt, F C.; Rarefied gas dynamics: Space science and engineering; Proceedings of the 18th International Symposium, Univ. of British Columbia, Vancouver, Canada; July 1992. pp. 53-63.
• Rarefied aerothermodynamic predictions for Mars Global Surveyor; Wilmoth, R G; Rault, D F; Cheatwood, F M; Engelund, W C; Shane, R W.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 314-322. 1999
• Aerothermodynamics of the Mars Global Surveyor Spacecraft; Shane, R W; Tolson, R. H.; NASA TM-19980041304, 1998.
• Mars Global Surveyor aerodynamics for maneuvers in Martian atmosphere; Shane, R W; Tolson, R H; Rault, D F.; AIAA 32nd Thermophysics Conference, June 1997.
• Aerodynamics of Mars Odyssey; Takashima, N; Wilmoth, R G.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Monterey, CA, Aug. 2002
• Plume modeling and application to Mars 2001 Odyssey aerobraking; Chavis, Z; Wilmoth, R.; 8th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, Saint Louis, MO, June 2002.
• Sharp-L1 technology demonstrator development - An aerothermodynamicperspective; Kolodziej, P; Bowles, J V; Brown, J L; Cornelison, C J; Lawrence, S L; Loomis, M P; Merriam, M L; Rasky, D J; Tam, T C; Wercinski, P F.; AIAA 34th ThermophysicsConference, June 2000.
• A CFD analysis of the orbital reentry experiment vehicle; Palmer, G; Prabhu, D; Venkatapathy, E.; First Europe-U.S. High Speed Flow Field Database Workshop, Naples, Italy, pp. 401-413, Nov. 1997.
• Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap; Horvath, Thomas J.; O’Connell, Todd F.; Cheatwood, F McNeil; Prabhu, Ramadas K.; Alter, Stephen J.; AIAA Paper 2002-4408; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Monterey, CA, Aug. 2002
• An aerothermal analysis and TPS sizing of the Mars 2001 Lander vehicle; Palmer, G; Chen, Y K; Papadopoulos, P; Tauber, M.; AIAA 37th Aerospace Sciences Mtg, Jan 1999.
Past Decade Aerothermodynamic References
RDB Aug 2005388
• Heat shield aeroheating predictions for a Mars Orbiter angle of attack aerocapture; Gulick, D S; Edquist, C T.; AIAA/ASME 7th Joint Thermophysics Conference, June 1998.
• Aerothermal heating simulations with surface catalysis for the Mars 2001 aerocapturemission; Papadopoulos, P; Venkatapathy, E; Henline, W; Wercinski, P F.; AIAA 35th
Aerospace Sciences Meeting & Exhibit, Jan. 1997.• Trajectory, aerothermal conditions, and thermal protection system mass for the
MARS 2001 aerocapture mission; Wercinski, P F; Henline, W; Tran, H; Milos, F; Papadopoulos, P; Chen, Y K; Venkatapathy, E; Tauber, M.; AIAA 35th Aerospace Sciences Meeting & Exhibit, Jan. 1997.
• A 3-D coupled CFD-DSMC solution method with application to the Mars Sample Return Orbiter; Glass, C E; Gnoffo, P A; Rarefied gas dynamics; Proceedings of the 22nd International Symposium, Sydney, Australia; July 2000. pp. 723-729.
• Convective and Radiative Heating for Vehicle Return from the Moon and Mars; R.B. Greendyke; P.A. Gnoffo; NASA TM-110185, July 1995.
• Simulated rarefied entry of the Galileo Probe into the Jovian atmosphere; Haas, B L; Milos, F S.; Journal of Spacecraft and Rockets. Vol. 32, no. 3, pp. 398-403. May-June 1995.
• Thermal protection system design studies for lunar crew module; Williams, S D; Curry, D M; Bouslong, S A; Rochelle, W C.; Journal of Spacecraft and Rockets. Vol. 32, no. 3, pp. 456-462. May-June 1995.
• High energy entry heating study for lunar/Mars aerocapturing vehicles; Rochelle, W C; An, M Y; Tam, L T; Williams, S D; Curry, D M; AIAA and ASME 6th Joint ThermophysicsConference, June 1994.
• Aerodynamic heating to spherically blunted cones at angle of attack; Shimshi, J P; Walberg, G D.; Journal of Spacecraft and Rockets. Vol. 32, no. 3, pp. 559, May-June 1995.
• Mars entry vehicle aerodynamic flight measurements; Blanchard, R C; Wilmoth, R G; Moss, J N.; ICAS, Congress, 21st, Melbourne, Australia, Sept. 1998.
• Rarefied Transitional Bridging of Blunt Body Aerodynamics; R.G. Wilmoth; R.C. Blanchard; J.N. Moss; 21st International Symposium on Rarefied Gas Dynamics, Marseille, France, July 1998.
Past Decade TPS References• Heatshielding problems of planetary entry - A review; Park, C; Tauber, M E.; AIAA
Fluid Dynamics Conference, 30th, Norfolk, VA, June 28-July 1, 1999.• Thermal protection system technology and facility needs for demanding future
planetary missions; Laub, B.; Venkatapathy, E.; European Space Agency Special Publication ESA SP, n 544, February, 2004, p 239-247.
• Thermal protection concepts and issues for aerocapture at Titan; Laub, B.; 39th AIAA/ASME/SAE Joint Propulsion Conference and Exhibit, Huntsville, AL, July 2003.
• Mars Exploration Rover TIRS Cover Thermal Protection System design verification; Szalai, Christine; Chen, Y-K; Loomis, Mark; Scrivens, Larry; Thoma, Benjamin; Buck, Stephanie; Hui, Frank; 36th AIAA Thermophysics Conference, June 2003.
• Arc jet screening of candidate ablative thermal protection materials for Mars Smart Lander; Laub, B; White, S.; AIAA Atmospheric Flight Mechanics Conference, Aug. 2002.
• New TPS materials for aerocapture; Laub, B.; Space Technology and Applications International Forum - STAIF 2002; Albuquerque, NM; 3-6 Feb. 2002. pp. 337-344.
• Probabilistic design of a Mars Sample Return Earth entry vehicle thermal protection system; Dec, J A; Mitcheltree, R A.; AIAA 40th Aerospace Sciences Meeting & Exhibit, Reno, NV, Jan. 2002.
• SHARP – NASA’s research and development activities in ultra high temperature ceramic nose caps and leading edges for future space transportation vehicles; Arnold, J.; Johnson, S.; Wercinski, P.; IAF Paper 01-V502, Oct. 2001.
• Two-Dimensional Implicit Thermal Response and Ablation Program for charring materials; Chen, Y-K; Milos, F.S.; Journal of Spacecraft and Rockets, Volume 38, 473-481, Aug 2001.
• Evaluation of high-temperature multilayer insulation for inflatable ballute; Kustas, F.M.; Rawal, S.P.; Wilcockson, W.H.; Edquist, C.T.; Thornton, J.M.; Sandy, C.; Journal of Spacecraft and Rockets, v 38, n 4, July-Aug 2001, p 630-631.
RDB Aug 2005389
Past Decade TPS References
RDB Aug 2005390
• Thermal performance of advanced charring ablator systems for future robotic and manned missions to Mars; Congdon, W M; Curry, D M.; AIAA ThermophysicsConference, 35th, Anaheim, CA, June 2001.
• Preliminary thermal analysis of a Mars Sample Return Earth Entry Vehicle; Amundsen, R M; Dec, J A; Mitcheltree, R A; Lindell, M C; Dillman, R A.; AIAA 34th
Thermophysics Conference, June 2000.• Mars Pathfinder heatshield design and flight experience; Willcockson, W H.; Journal
of Spacecraft and Rockets. Vol. 36, no. 3, pp. 374-379. 1999.• Analysis of Galileo probe heatshield ablation and temperature data; Milos, F S;
Chen, Y K; Squire, T H; Brewer, R A.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 298-306. May-June 1999
• Galileo Probe heat shield ablation experiment; Milos, F S; Journal of Spacecraft and Rockets. Vol. 34, no. 6, pp. 705-713. Nov.-Dec. 1997.
• Mars Pathfinder entry temperature data, aerothermal heating, and heatshieldmaterial response; Milos, F S; Chen, Y K; Congdon, W M; Thornton, J M.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 380-391. May-June 1999.
• Ablation and thermal response program for spacecraft heatshield analysis; Chen, Y-K; Milos, F.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 475-483. May-June 1999.
• Aerothermal Effects of Cavities and Protuberances for High-Speed Sample Return Capsules; Olynick, D; Kontinos, D.; 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, NV, United States, Jan. 1999.
• New TPS design strategies for planetary entry vehicle design; Olynick, D; Loomis, M; Chen, Y K; Venkatapathy, E; Allen, G.; AIAA 37th Aerospace Sciences Meeting and Exhibit, 37th, Reno, NV, Jan. 11-14, 1999.
• An aerothermal analysis and TPS sizing of the Mars 2001 Lander vehicle; Palmer, G; Chen, Y K; Papadopoulos, P; Tauber, M.; AIAA 37th Aerospace Sciences Meeting and Exhibit, 37th, Reno, NV, Jan. 11-14, 1999.
Past Decade TPS References• Heatshield Erosion in a Dusty Martian Atmosphere; Papadopoulos, P; Chang, I D;
Tauber, M E; Journal of Spacecraft and Rockets. Vol. 30, no. 2, pp. 140-151. Mar.-Apr. 1993
• Particle impact risk assessment for ablative thermal protection systems; Naughton, J W; Venkatapathy, E; Loomis, M P.; AIAA 36th Aerospace Sciences Meeting, Jan. 1998.
• Phenolic Impregnated Carbon Ablators (PICA) as Thermal Protection Systems for Discovery Missions; TRAN, HUYK; JOHNSON, CHRISTINEE; RASKY, DANIELJ; HUI, FRANKCL; HSU, MING-TA; CHEN, TIMOTHY; CHEN, Y K; PARAGAS, DANIEL; KOBAYASHI, LOREEN; NASA-TM-110440;1997.
• Qualification of the forebody heatshield of the Stardust's Sample Return Capsule; Tran, H K; Johnson, C E; Hsu, M T; Chem, H C; Dill, H; Chen-Johnson, A.; AIAA 32nd
Thermophysics Conference, 32nd, June 1997.• Trajectory based, 3-dimensional heating and ablation calculations for the Apollo
Lunar/Earth return capsule; Henline, W.D.; Chen, Y-K; Palmer, G.E.; Stewart, D.A.; AIAA Paper 93-2788; AIAA, Thermophysics Conference, July 1993.
• TPS design for aerobraking at Earth and Mars; WILLIAMS, S D; GIETZEL, M M; ROCHELLE, W C; CURRY, D M.; NASA-TM-104739; 1991.
• Thermal protection systems manned spacecraft flight experience; Curry, D M; In NASA Langley Research Center, Current Technology for Thermal Protection Systems p 19-41 (SEE N93-12447 02-18), 1992.
• Thermal Protection Materials: Thermophysical Property Data; Williams, S.D.; Curry, Donald M.; NASA-RP-1289, 1992.
• Heat Protection for Atmospheric Entry into Saturn, Uranus and Neptune; Tauber, M.E.; AAS 71-145, June 1971.
RDB Aug 2005391
Past Decade Flight Dynamics References
RDB Aug 2005392
• Mars Pathfinder Six-degree-of-freedom entry analysis; Braun, R D; Powell, R W; Engelund, W C; Gnoffo, P A; Weilmuenster, K J; Mitcheltree, R A.; Journal of Spacecraft and Rockets. Vol. 32, no. 6, pp. 993-1000. Nov.-Dec. 1995
• Mars Pathfinder atmospheric entry - Trajectory design and dispersion analysis; Spencer, D A; Braun, R D.; Journal of Spacecraft and Rockets. Vol. 33, no. 5, pp. 670-676. Sept.-Oct. 1996.
• Six-degree-of-freedom entry dispersion analysis for the METEOR recovery module; Desai, P N; Braun, R D; Powell, R W; Engelund, W C; Tartabini, P V.; Journal of Spacecraft and Rockets. Vol. 34, no. 3, pp. 334-340. 1997
• Mars Polar Lander aerothermodynamic and entry dispersion analysis; Queen, E M; Cheatwood, F M N; Powell, R W; Braun, R D; Edquist, C T.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 421-428. May-June 1999.
• Entry dispersion analysis for the Genesis sample return capsule; Desai, P N; Cheatwood, F M.; Journal of Spacecraft and Rockets. Vol. 38, no. 3, pp. 345-350. 2001
• Entry dispersion analysis for the Stardust comet sample return capsule; Desai, P N; Mitcheltree, R A; Cheatwood, F M.; Journal of Spacecraft and Rockets. Vol. 36, no. 3, pp. 463-469. 1999.
• Mars Exploration Rovers Six Degree of Freedom Entry Analysis; Desai, P.N.; Schoenenberger, M. and Cheatwood, F.M., AAS Paper 03-642, August 2003.
• Aeromaneuvering in the martian atmosphere: simulation-based analyses; Smith, R S; Mease, K D; Bayard, D S; Farless, D L.; Journal of Spacecraft and Rockets. Vol. 37, no. 1, pp. 139-142. Feb. 2000.
• Atmospheric maneuvering during Martian entry; Tauber, M E; Bowles, J V; Yang, L.; AIAA Atmospheric Flight Mechanics Conference, Aug. 1988. pp. 124-133.
• An Atmospheric Guidance Algorithm Testbed for the Mars Surveyor Program 2001 Orbiter and Lander; Striepe, S A; Queen, E M; Powell, R W; Braun, R D; Cheatwood, F M N; Aguirre, J T; Sachi, L A; Lyons, D T.; NASA TM-19980219469.
• Mars Smart Lander simulations for entry, descent, and landing; Striepe, S A; Way, D W; Dwyer, A M; Balaram, B.; AIAA Atmospheric Flight Mechanics Conference, Aug. 2002.
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Past Decade GN&C References• Atmospheric Guidance Concepts for an Aeroassist Flight Experiment; Gamble, J.D.;
Cerimele, C.J.; Moore, T.E.; Higgins, J.; Journal of the Astronautical Sciences, v 36, n 1 pt 2, Jan-Jun, 1988, p 45-71.
• Predictor-corrector guidance algorithm for use in high-energy aerobraking system studies; Braun, R.D.; Powell, R.W.; Journal of Guidance, Control, and Dynamics, Vol. 15, Jun 1992.
• Six-degree-of-freedom guidance and control analysis of Mars aerocapture; Powell, R W; Braun, R D.; Journal of Guidance, Control, and Dynamics. Vol. 16, no. 6, Nov-Dec 1993.
• Analytic drag control for precision landing and aerocapture; Bryant, L E; Tigges, M A; Ives, D G.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Aug. 1998.
• Blended control, predictor-corrector guidance algorithm - An enabling technology for Mars aerocapture; Jits, R Y; Walberg, G D.; International Astronautical Congress, 52nd, Toulouse, France, Oct. 2001.
• Mars aerocapture - Extension and refinement; Wercinski, P F; Lyne, J E.; Journal of Spacecraft and Rockets. Vol. 31, no. 4, pp. 703-705. July-Aug. 1994.
• Nondimensional analysis of reaction-wheel control for aerobraking; Johnson, W R; Longuski, J M; Lyons, D T.; Journal of Guidance, Control, and Dynamics. Vol. 26, no. 6, pp. 861-868. Nov. 2003.
• An analytical assessment of aerocapture guidance and navigation flight demonstration for applicability to other planets; Graves, C A; Masciarelli, J P.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Aug. 2002.
• Aerocapture guidance algorithm comparison campaign; Rousseau, S; Perot, E; Graves, C; Masciarelli, J; Queen, E.; AIAA/AAS Astrodynamics Specialist Conference and Exhibit, Monterey, CA, Aug. 2002.
• CNES-NASA studies of the Mars Sample Return Orbiter Aerocapture Phase; Fraysse, H.; Rousseau, S.; Powell, R.; Striepe, S.; IAF Paper 00-A605, Oct 06, 2000.
• An analytic aerocapture guidance algorithm for the Mars Sample Return Orbiter; Masciarelli, J P; Rousseau, S; Fraysse, H; Perot, E.; AIAA Atmospheric Flight Mechanics Conference, Denver, CO; 14-17 Aug. 2000. pp. 525-532.
Past Decade GN&C References• The Mars Surveyor 2001 Lander: A First Step Toward Precision Landing; Braun, R.D.;
Powell, R.W.; Cheatwood, F.M.; Spencer, D.A.; and Mase, R.A.; IAF-98-Q.3.03, 1998.• Navigation strategy for the Mars 2001 lander mission; Mase, R A; Spencer, D A; Smith, J
C; Braun, R D.; Proceedings of the AAS/AIAA Astrodynamics Conference, Aug. 1999. pp. 2193-2208.
• Numerical Roll Reversal Predictor-Corrector Aerocapture and Precision Landing Guidance Algorithms for the Mars Surveyor Program 2001 Missions; Powell, R W.; NASA TM-19980237136.
• Navigation and guidance for the Mars Surveyor '98 mission; Kallemeyn, P H J; Knocke, P C; Burkhart, P D; Thurman, S W.; AIAA/AAS Astrodynamics Specialist Conference and Exhibit, Boston, MA; 10-12 Aug. 1998. pp. 471-481. 1998
• Mars Surveyor Program landing radar - Overview of flight tests and GN&C interfaces; Cuseo, J A; Haack, B R; Proceedings of the 21st Annual AAS Rocky Mountain Guidance and Control Conference, Breckenridge, CO; Feb. 1998. pp. 539-558.
• Overview - Precision landing/hazard avoidance concepts and MEMS technologyinsertion for human Mars lander missions; Benjamin, A L; Bolen, S M; Smit, G N; Cuseo, J A; Lindell, S D.; AIAA/IEEE 16th Digital Avionics Systems Conference (DASC), Irvine, CA; 26-30 Oct. 1997. pp. 8.5-18 to 8.5-25.
• Autonomous guidance and control design for hazard avoidance and safe landing on Mars; Wong, E C; Singh, G; Masciarelli, J.; AIAA Atmospheric Flight Mechanics Conference and Exhibit, Monterey, CA, Aug. 2002.
• Fuel-optimal bank-angle control for lunar-return aerocapture; Meyer, J L; Silverberg, L; Walberg, G D.; Journal of Spacecraft and Rockets. Vol. 32, no. 1, pp. 149-155. Jan-Feb 1995.
• Mars aerocapture using continuous roll techniques; Willcockson, W H.; Proceedings of the AAS/AIAA Astrodynamics Conference, Pt. 3; Aug. 1991. pp. 1859-1881.
• Mars entry-to-landing trajectory optimization and closed loop guidance; Ilgen, Marc R.; Manning, Raymund A.; Cruz. Manuel I.; AAS Paper 91-501, Aug 1, 1991
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Past Decade Supersonic Parachute References
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• System Design Overview of the Mars Pathfinder Parachute Decelerator Subsystem; Fallon, E.J.; AIAA Paper 97-1511, 1997.
• Mars Exploration Rover Parachute Decelerator System program overview; Witkowski, A; Bruno, R.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
• Flight reconstruction of the Mars Pathfinder Disk-Gap-Band parachute drag coefficient; Desai, P N; Schofield, J T; Lisano, M E.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003, AIAA Paper 2003-2126.
• Wind tunnel testing of various Disk-Gap-Band parachutes; Cruz, J R; Mineck, R E; Keller, D F; Bobskill, M V.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
• Structural testing of parachutes in the National Full-Scale Aerodynamics Complex 80-by-120-Foot Wind Tunnel at NASA Ames Research Center; Zell, P T; Cruz, J R; Witkowski, A.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
• Opening loads analyses for various Disk-Gap-Band parachutes; Cruz, J R; Kandis, M; Witkowski, A.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
• The Mars Exploration Rover Entry, Descent and Landing and the Use of Aerodynamic Decelerators; Steltzner A., Desai, P., Lee, W., and Bruno, R.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003, AIAA Paper 2003-2125.
• Development of an improved performance parachute system for Mars missions; Masciarelli, J P; Cruz, J R; Hengel, J E.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
• Hypercone inflatable supersonic decelerator; Brown, G J; Epp, C; Graves, C; Lingard, S; Darley, M; Jordan, K.; 17th AIAA Aerodynamic Decelerator Systems Technology Conference and Seminar, Monterey, CA, May 19-22, 2003.
Past Decade Deployable Aeroshell References• A Survey of Ballute Technology for Aerocapture; Rohrschneider, R.R.; and Braun,
R.D.; Submitted for publication in the Journal of Spacecraft and Rockets, 2005.• Computational Analysis of Towed Ballute Interactions; Gnoffo, Peter A.; Anderson,
Brian P.; AIAA Paper 2002-2997; 8th AIAA/ASME Joint Thermophysics and Heat Transfer Conference, June 2002.
• Technology development for deployable aerodynamic decelerators at Mars; Masciarelli, J P.; Space Technology and Applications International Forum - STAIF 2002; Proceedings; Albuquerque, NM; 3-6 Feb. 2002. pp. 345-352.
• Aerocapture trajectories for spacecraft with large, towed ballutes; Hall, J L; Le, A K.; Proceeding of the 11th Annual AAS/AIAA Space Flight Mechanics Meeting, Feb. 11-15, 2001, vol. 2, p. 1857-1872.
• Experimental investigation of the flow over a toroidal aerocapture ballute; Rasheed, A; Fujii, K; Hornung, H G; Hall, J L.; AIAA 19th Applied Aerodynamics Conference, June 2001.
• Attached inflatable ballute for spacecraft deceleration; Kustas, F.M.; Rawal, S.P.; Willcockson, W.H.; Edquist, C.T.; Thornton, J.M.; Giellis, R.T.; IEEE Aerospace Conference Proceedings, v 7, 2000, p 421-427.
• A Review of Ballute Technology for Planetary Aerocapture; J. Hall; International Conference on Low-Cost Planetary Missions, May 2, 2000.
• A light-weight inflatable hypersonic drag device for Venus entry; McRonald, A; Proceedings of the AAS/AIAA Astrodynamics Conference, Aug. 1999. pp. 819-830.
• A light-weight hypersonic inflatable drag device for a Neptune orbiter; McRonald, A D; AAS/AIAA Space Flight Mechanics Meeting, Jan. 2000. pp. 1085-1099. 2000.
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