strategies for mars network missions via an alternative entry, descent, and landing architecture
DESCRIPTION
10 th International planetary probe workshop. Strategies for mars network missions via an alternative entry, descent, and landing architecture. 17-21 June, 2013; San Jose State University, CA, United States. Sarag J. Saikia, Blake Rogers, James M. Longuski - PowerPoint PPT PresentationTRANSCRIPT
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STRATEGIES FOR MARS NETWORK MISSIONS VIA AN ALTERNATIVE ENTRY, DESCENT, AND LANDING ARCHITECTURE
10TH INTERNATIONAL PLANETARY PROBE WORKSHOP17-21 June, 2013; San Jose State University, CA, United States
Sarag J. Saikia, Blake Rogers, James M. LonguskiSchool of Aeronautics and Astronautics, Purdue University
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MISSION CONCEPT AND ARCHITECTUREGOAL: Deliver four Mars Phoenix-class landers with a minimum separation of 3,000 km via a single launch from Earth
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LAUNCH AND CRUISE CONFIGURATION
Payload mass of 60 kg (x4) Flight system mass of 1380 kg + 2% reserve (x2)
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SINGLE-EVENT DRAG MODULATION
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BALLISTIC COEFFICIENT ANALYSIS
0
0.5
1
05
1015
200
0.5
1
1.5
ratio
fs/
s
p/s
& Dds ds
s D s
C Amm C A
21fs s
11p s
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INTERPLANETARY TRAJECTORYTRAJECTORY CONSTRAINTS
• Atlas V 541 Launch Vehicle• Maximum launch V∞ for a mass of 1382 kg ≈ 7 km/s
•Maximum entry speed of 6 km/s•Reduces the heating rates and heat loads of EDL•Corresponds to a maximum arrival V∞ of ≈ 3.5 km/s
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INTERPLANETARY TRAJECTORY7 DAY SEPARATION: LOW THRUST
Vehicle Number
Launch Date(d/m/y)
Arrival Date(d/m/y)
Launch V∞ (km/s)
Arrival V∞(km/s)
Onboard Propellant
Required (kg)1 7/29/2020 2/22/2021 3.82 2.54 02 7/29/2020 3/1/2021 3.82 1.64 58.5 1 10/12/2022 6/22/2023 6.00 2.37 02 10/12/2022 8/20/2023 6.00 1.39 63.1 1 4/12/2033 10/27/2033 2.99 3.37 02 4/12/2033 11/26/2033 2.99 2.58 56.0
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RANGE SEPARATION AND DIVERT CAPABILITY
Global Reach Capability
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MONTE CARLO RESULTS: RANGE SEPARATIONLANDING ERROR
Flight System 1
Flight System 2
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STAGNATION-POINT HEATING RATE
0 1 2 3 4 5 60
5
10
15
20
25
30
Velocity, km/s
Stag
natio
n-Po
int H
eat R
ate,
W/c
m2
SecondaryRelease
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DECELERATION, RELEASE
0 1 2 3 4 5 60
2
4
6
8
10
12
14
Velocity, km/s
Dec
eler
atio
n, E
arth
g
Secondary Release
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MONTE CARLO RESULTSSPACECRAFT RANGE AND RANGE SEPARATION DISTRIBUTION
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MONTE CARLO RESULTSPRIMARY SPACECRAFT RANGE DISTRIBUTION
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MONTE CARLO RESULTSINTEGRATED HEAT LOAD: PRIMARY SPACECRAFT
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MONTE CARLO RESULTSINTEGRATED HEAT LOAD: SECONDARY SPACECRAFT
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OTHER POTENTIAL APPLICATIONSRange Sep. [km]
FPA Release Time
[s]
Secondary Landing Error
3-σ [km]
Secondary Heat Load
[J/cm2]
Stagnation Heat Rate
[J/cm2]
Comment
1054 -10.2° 122 60 2900 28 ~ Phoenix
346 -10.8° 122 15 1310 28
140 -10.8° 144 12 425 11 TPS Needed?
706 -10.0° 144 45 2004 28
270 -10.0° 164 13 806 13 TPS Needed?
Primary heat load for all the cases is < 2200 J/cm2
Primary landing 3-σ error for all the cases is < ±10 km
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CONCLUSIONS Low-Thrust Propulsion represents an attractive ‘augmentation’ for any
future mission to Mars Benign aerothermal environments, reduced heat rates and loads Very low ballistic coefficient achievable: no supersonic decelerator
(parachute) required Increased risks of separation: flight systems, spacecraft from a flight
system Mass increase: due to extra spacecraft adapter; Decrease due to
reduction in cruise stages and supersonic parachutes Other potential applications of multiple spacecraft lander/orbiter
missions Single Atlas V 541 launch required Incorporation of the guidance on the second will reduce the landing
error
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ACKNOWLEDGMENT
Thanks to the IPPW10 student organizing committee for providing the ‘generous’ scholarship to attend the workshop
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QUESTIONS?
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BACKUP
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BACK UP
InstrumentsMass Breakdown
Payload mass of 60 kg (same as Phoenix mission)
Mass of flight system is 1380 kg + 2% reserve
Mission Mass BreakdownFlight System 1 or 2 (Identical)
Mass (kg)
Additional Margins
Lander 1 343
Lander 2 343Backshell and Parachute 1 110Backshell and Parachute 2 110
Ex.Heat Shield 1 124
Heat Shield 2 62
Cruise Stage 100
Propellant 1 65
Propellant 2 65 40% Secondary Spacecraft Adapter and Release Mechanism
60
Total Mass 1382 ~30 kg
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MONTE CARLO RESULTSINTEGRATED HEAT LOAD
% TPS mass is estimated using an empirical formula based on previous probe missions
Slightly high TPA mass for primary, and lower for secondary
Total range separation requirement is the determinant of % TPS mass of secondary
For low range requirements (<500km) secondary needs no TPS mass at all!
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ANALYSIS OF DRAG MODULATION
Parameter Value
70.0 kg/m2
37.6 kg/m2
20.5 kg/m2
s
pfs
Combine with the previous slide #4
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DRAG SKIRT OPTIONS
Heat Shield Extension
Rigid Deployable Decelerator (ADEPT)
Hypersonic Inflatable Aerodynamic Decelerator (HIAD)
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UNCERTAINTY ANALYSIS: MONTE CARLOSIMULATION UNCERTAINLY MODEL PARAMETERS AND INPUT
Parameter Details/ Models 3σ Comments
No. of Runs 1000 -
Density Model Mars-GRAM 2005 Include Details
Check Papers
FPA 0.003° Phoenix
Velocity 0.439 m/s Phoenix
Altitude 0 m
Range 0.002° Phoenix
Ballistic coefficient
1 kg/m2 Include dispersions in mass and
Aerodynamic Coefficients
Time of release Time Trigger 1-2 seconds
Correlated with Velocity Trigger