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

MISSION CONCEPT AND ARCHITECTUREGOAL: Deliver four Mars Phoenix-class landers with a minimum separation of 3,000 km via a single launch from Earth

LAUNCH AND CRUISE CONFIGURATION

Payload mass of 60 kg (x4) Flight system mass of 1380 kg + 2% reserve (x2)

SINGLE-EVENT DRAG MODULATION

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

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

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

RANGE SEPARATION AND DIVERT CAPABILITY

Global Reach Capability

MONTE CARLO RESULTS: RANGE SEPARATIONLANDING ERROR

Flight System 1

Flight System 2

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

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

MONTE CARLO RESULTSSPACECRAFT RANGE AND RANGE SEPARATION DISTRIBUTION

MONTE CARLO RESULTSPRIMARY SPACECRAFT RANGE DISTRIBUTION

MONTE CARLO RESULTSINTEGRATED HEAT LOAD: PRIMARY SPACECRAFT

MONTE CARLO RESULTSINTEGRATED HEAT LOAD: SECONDARY SPACECRAFT

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

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

ACKNOWLEDGMENT

Thanks to the IPPW10 student organizing committee for providing the ‘generous’ scholarship to attend the workshop

QUESTIONS?

BACKUP

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

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!

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

DRAG SKIRT OPTIONS

Heat Shield Extension

Rigid Deployable Decelerator (ADEPT)

Hypersonic Inflatable Aerodynamic Decelerator (HIAD)

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