Alternative Strategies for Exploring Mars and the Moons of Mars

GLEX-2012.08.2.2x12575-Presentation Drake, Baker, Hoffman, Landau, Voels

National Aeronautics and Space Administration National Aeronautics and Space Administration

Alternative Strategies for Exploring Mars and the Moons of Mars

Bret G. Drake / NASA Johnson Space Center, USA John D. Baker / Jet Propulsion Laboratory, California Institute of Technology, USA Stephen J. Hoffman, PhD. / Science Applications International Corporation, USA Damon Landau, PhD. / Jet Propulsion Laboratory, California Institute of Technology, USA Stephen A. Voels, PhD. / Science Applications International Corporation, USA

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https://ntrs.nasa.gov/search.jsp?R=20120009026 2019-04-05T01:46:55+00:00Z


“Set far-reaching exploration milestones. By 2025, begin crewed missions beyond the moon, including sending humans to an asteroid. By the mid-2030s, send humans to orbit Mars and return them safely to Earth”

2010 National Space Policy

“Set far-reaching exploration milestones. By 2025, begin crewed missions beyond the moon, including sending humans to an asteroid. By the mid-2030s, send humans to orbit Mars and return them safely to Earth”

• Recent discussions have focused on the prospect of conducting a human mission to Mars orbit as a validation test prior to the surface mission

• These strategies are drawn from the historical precedence of the Apollo test missions

• Apollo 8 and 10 flew very similar mission profiles to the eventual surface missions

• But this Apollo analogy may not apply for much harder and longer Mars orbital missions

Careful examination of the required capabilities and knowledge needed is necessary to fully understand the key issues and applicability of a human mission to Mars orbit prior to a surface mission

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Mar

s Moo

n Su

rfac

e (P

hobo

s/De

imos

)

Mission Sequence

Mission Summary

Mar

s Orb

it

Day 1

Mission Site: Phobos / Deimos

Capture into High-Mars Orbit

Deimos: 20,063 km circular 0.9 deg, 1.26 day period

Phobos: 5981 km circular 1 deg, 0.32 day period

Departure Trajectory

• Deimos survey • SEV anchoring

• Test payload anchoring methods. • Collect initial samples

• Science package deployment • Drilling operations

Day 60

SEV 1 + CPS 1

Crew:

• Test EVA procedures & mobility • Test payload anchoring methods.

SEV 2 + CPS 2

Deimos Phobos

HMO: 1-sol 250 x 33,813 km

Assumed Mars Orbit Strategy 1. Capture into a 1-sol parking orbit with proper

plane change to match departure asymptote 2. Leave Mars Transfer Vehicle in 1-sol parking

orbit 3. Prepare for orbital operations 4. Utilize SEV-1 to explore Deimos for 14 days

(1,370-2,770 m/s delta-v required)

5. Utilize SEV-2 to explore Phobos for 14 days (1,700-3,170 m/s delta-v required)

6. Prepare for Mars departure 7. Trans-Earth Injection

Jettison

Jettison

Rescue CPS

Short Stay Mars Vicinity Operations

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Mars Ballistic Trajectory Classes

EARTH DEPARTURE

MARS ARRIVAL

γ

MARS DEPARTURE

VENUS SWING-BY

SUN

EARTH RETURN EARTH

DEPARTURE

MARS ARRIVAL

γ

MARS DEPARTURE

SUN

EARTH RETURN

Long-Stay Missions (Conjunction) • Variations about the minimum

energy mission • Long-stays at Mars (~500 days)

and long overall duration (900-1000 days)

Short-Stay Missions (Opposition) • Variations of missions with short

Mars surface stays (20-60 days) and may include Venus swing-by

• Total mission duration typically 540-840 days

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Example Delta-v versus Mission Duration

0

10

20

30

40

50

60

- 200 400 600 800 1,000

Tota

l Del

ta-v

(km

/s)

Total Mission Duration (Days)

Crew Vehicle Total Delta-VOpposition Class - 2033 "Good" Opportunity

20 Day Stay40 Day Stay60 Day Stay80 Day Stay100 Day StayConjunction

Trajectory Set: 27 January 2012

ORBIT ASSUMPTIONSEarth Departure Orbit = 400 X 400 kmMars Arrival Oribt = 250 X 33,813 kmMars Departure Oribt = 250 X 33,813 kmDirect Entry at Earth Return

PLANETARY ARRIVAL ASSUMPTIONSMars Propulsive CaptureCapture Plane: As isDirect Earth Entry @ 13 km/s

Opposition Class “Short-Stay” Conjunction Class “Long-Stay”

Stay Time Varies (550-730 Days)

60-Day One-Way Transits

200-Day One-Way Transits

No Venus Swing-by

0

200

400

600

800

1000

1200

Tota

l Miss

ion

Dura

tion

(Day

s)

Earth Departure Opportunity

Opposition Class (60 Day Stay) MissionsMission Duration

Inbound (130 - 560 d) At Mars (60 - 60 d) Outbound (140 - 530 d)

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GLEX-2012.08.2.2x12575-Presentation Drake, Baker, Hoffman, Landau, Voels 6

Total Crew Mission ∆V Sensitivity Co-planar Trajectories

Mission opportunities (Earth departure date) occur approximately every 26 months

Due to the difference in orbits of both the Earth and Mars, the required trajectories vary for each Earth departure date

Short-stay (opposition) missions demonstrate significantly variation

Less sensitivity occurs for long-stay (conjunction) missions

-

5

10

15

20

200 400 600 800 1,000

Tota

l Miss

ion

Delta

-V (

km/s

)


202820302033203520372039204120432045

Opposition Class “Short-Stay”

(60 Days at Mars)

Conjunction Class “Long-Stay” (60-200 Day Transits)

Mission Characteristics • LEO: 400 km x 400 km • HMO: 250 km x 33,813 km • Direct Earth Entry: 13 km/s


TRANSLATING ∆V TO MASS

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Transportation and Exploration Systems Assumptions

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Multi Purpose Crew Vehicle

• Crew delivery to Earth orbit and high-energy Earth return

• CM inert = 9.8 t

• SM inert = 4.5 t

• SM specific impulse = 328 s

Deep Space Habitat

• Support a crew of 4-6 from 365 – 1100 days

• Mass Range : 28-65 t

• Consumables loaded based on crew size & mission duration

Solar Electric Propulsion

• Low-thrust with solar power

• Spacecraft alpha ~30 kg/kw

• Specific impulse = 1800-4000 s

• Xe tank fraction = 5%

• Total power varies

Space Exploration Vehicle • Primary purpose is for

exploration of the moons

• Crew of 2 for 14 days

• Nominal mass = 6.7 t

• CH4 Stage when needed:

• Stage Fraction: 15%

• Isp: 355 s

Chemical Propulsion Stage

• High-thrust propulsive stage

• Parametric design with each stage optimized

• Zero-boiloff cryo management

• Stage fraction ~ 23%

• Specific impulse = 465 s

Nuclear Electric Propulsion

• Low-thrust with nuclear power

• Spacecraft alpha ~20 kg/kw

• Specific impulse = 1800-4000 s

• Xe tank fraction = 5%

• Total power varies

Mars Landers

• Parametric sizing of entry and landing systems

• Inflatable (HIAD) entry system assumed

• Wet lander mass: 89-113 t

Nuclear Thermal Propulsion • High-thrust nuclear propulsion

• NERVA-derived common core propulsion (20 t core)

• 3 x 111 kN engines

• Specific Impulse = 900 s

• All LH2 fuel with zero boil-off

• Drop tanks @ 27% tank fraction

Space Launch System

• Gross Performance ~ 130 t

• Net Performance ~ 120.4 t

• Performance estimates to negative perigee conditions: (-87 km x 241 km)


ISS ISS

ISS

Opposition Class Missions Crew Vehicle Mass With 60-Days at Mars

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0

200

400

600

800

1,000

1,200

0 200 400 600 800 1,000

Tota

l Mas

s in

LEO

(t)


Nuclear Electric Propulsion

202820302033203520372039204120432045

0

200

400

600

800

1,000

1,200

0 200 400 600 800 1,000

Tota

l Mas

s in

LEO

(t)


202820302033203520372039204120432045

Nuclear Electric

Nuclear Thermal Chemical

Solar Electric

ISS


Conjunction Class Missions Crew Vehicle Mass Total Mission Durations Approximately 1,100 Days

10

Nuclear Electric

Nuclear Thermal Chemical

Solar Electric

0

200

400

600

800

1,000

1,200

300 400 500 600 700 800

Tota

l Mas

s in

LEO

(t)

Useful Exploration Time at Mars (Days)

202820302033203520372039204120432045

0

200

400

600

800

1,000

1,200

300 400 500 600 700 800

Tota

l Mas

s in

LEO

(t)


Nuclear Thermal Propulsion

202820302033203520372039204120432045

0

200

400

600

800

1,000

1,200

300 400 500 600 700 800

Tota

l Mas

s in

LEO

(t)


Nuclear lectric Propulsion

202820302033203520372039204120432045

0

200

400

600

800

1,000

1,200

300 400 500 600 700 800

Tota

l Mas

s in

LEO

(t)


202820302033203520372039204120432045

ISS ISS

ISS

Nuclear Electric Solar Electric

ISS


ISS

For the Conjunction Class missions, the stay at Mars can be lengthened to allow faster one-way transits to and from Mars

Practical limits exist, due to the physics of the trajectories

The limits are dependent on the propulsion technology choice

The range of practical transits for the Nuclear Thermal Propulsion depicted here

Transit times also dependent on Earth departure year

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Conjunction Class Missions Crew Vehicle Mass Shortening the One-Way Transit Times – Nuclear Thermal Propulsion

-

200

400

600

800

1,000

1,200

0 50 100 150 200 250 300

Tota

l Mas

s in

LEO

(t)

One-Way Tranist Time (Days)

202820302033203520372039204120432045


ADDITIONAL MISSION DESIGN CONSIDERATIONS

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Short Stay Orbital Operations Concept High-Thrust Missions

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Phobos 0.32 day period

Deimos 1.26 day period

MTV Remains in Parking Orbit

1 sol period Inclination Varies

Arrival Trajectory Declination Varies

Arrival Asymptote Declination Varies

Departure Asymptote Declination Varies

Deimos Inclination = 0.9 deg Phobos Inclination = 1.1 deg

Assumed Mars Orbit Strategy 1. Cargo pre-deployed into high-Mars parking orbit 2. Crew captures into high-Mars parking orbit 3. Crew transfer vehicle remains in high-Mars

parking orbit 4. Crew prepares for orbital operations 5. Space Exploration Vehicle #1 used to explore

Deimos for 14 days (1,370-2,770* m/s delta-v required)

6. Space Exploration Vehicle #2 used to explore Phobos for 14 days (1,700-3,170* m/s delta-v required)

7. Prepare for Mars departure 8. Trans-Earth Injection

Departure Trajectory Declination Varies

MTV Remains in Parking Orbit

1 sol period

* These values still under review


Mars Orbital Missions • Mars orbital missions, including exploration of the moons of

Mars, are conducted entirely in deep-space • Reducing the exposure of the mission crew to the hazards of

deep-space is of prime concern for these missions • Practical considerations (transportation technology and number

of launches) will limit mission durations to not much less than 600 days. Thus, human health issues cannot be obviated by propulsion technology alone

• If there is no true difference between 600 and 900 days from a human health perspective, then long-stay (conjunction class missions) should be used

Mars Surface Missions • Application of short-stay opposition class missions is not so clear • Short-surface stay alone is insufficient to ameliorate the human

health concerns (zero-g and radiation) • It is anticipated, though yet to be confirmed, that the surface

environment of Mars (partial gravity and radiation) may provide sufficient human health mitigation for long-stay missions

• Landing large payloads remains a key challenge for Mars surface missions, both short and long stay

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Key Takeaways

Alternative Strategies for Exploring Mars and the Moons of Mars

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