Earth-Mars Artificial-G NEP Architecture Sun-Earth L2 Architecture 3-Week Parametric Trade Study

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Earth-Mars Artificial-G NEP Architecture

Sun-Earth L2 Architecture

3-Week Parametric Trade Study

Presented to JSC/Exploration Office

March 3, 2003

Low Thrust Trajectory Team – GRC, JPL, JSC, MSFC

Presentation prepared by: Jerry Condon / JSC / EG5 / 281.483.8173 / gerald.l.condon@nasa.gov

Preliminary

- GRC
- Melissa McGuire, Rob Falk

- JPL
- Jon Sims, Greg Whiffen

- JSC
- Jerry Condon, Ellen Braden, Dave Lee, Kyle Brewer, Carlos Westhelle
- Jim Geffre

- MSFC
- Reginald Alexander, Larry Kos, Kirk Sorensen

- 2 Studies – NEP parametric mission design trades
Study 1 - Round trip Earth/Mars mission

- Augment results from NEP (EM-L1 departure) study done last year at JSC
- Determine cost (mass, time) to depart from Earth orbit and spiral to/from selected Mars parking orbits for Earth return
Study 2 - Sun-Earth libration point (L2) mission

- Determine cost (mass, time) to depart from Earth orbit and spiral to/from selected Mars parking orbits for Earth return
- Deploy/maintenance of satellite constellation
- Dress rehearsal for Mars mission

- Augment results from NEP (EM-L1 departure) study done last year at JSC

- JSC/ExPO – Kent Joosten, Bret Drake, Brenda Ward, etc.
- HQ/Gary Martin

- Study 1 - Round Trip Earth/Mars Mission
- Study 2 - Sun-Earth L2 Libration Point Mission
- Appendix
- Mars Arrival Parking Orbit Analysis
- Mars Parking Orbit Lifetime
- Integrated Reference Mission
- Effects of Parking Orbit Geometry on Mars Lander Mass
- Trapped Proton Belt Data

Study 1Round Trip Earth/Mars Mission

Round Trip Earth/Mars Mission

- Two vehicles
- NEP Mars Transfer Vehicle (MTV)
- Object of parametric study

- Lander/Ascent Vehicle (LAV)
- Previously deployed at Mars

- NEP Mars Transfer Vehicle (MTV)
- Use same vehicle specifications as last year (2002) study for Artificial Gravity Mars transfer vehicle*:
- Power = 6 MW, Engine efficiency = 60%, Isp = 4000 sec, Tankage fraction = 5%
- Final mass target (back at Earth) = 89mt
- No thrust vector turning constraints
- Determine vehicle thrust vector steering requirements unconstrained by Artificial Gravity (AG) vehicle configurations
- Results may influence AG vehicle configurations

- No thrust vector turning constraints

- 2026 opportunity, <90 day stay in Mars vicinity >30 days surface stay
- Initial Earth orbit 700 km circular LEO
- Crew taxi transfers crew from ground to crew transfer altitude (30,000 – 90,000 km)
- No constraint on heliocentric closest approach to Sun

Fire Baton

Artificial-G NEP

Mars Transfer Vehicle

* Preliminary Assessment of Artificial Gravity Impacts to Deep-Space Vehicle Design, JSC/EX Document No. EX-02-50, 2002

- Perform parametric study to enhance understanding of propellant and trip time requirements for both a round trip Earth-Mars mission and a Sun-Earth L2 Libration Point mission
- Compare results generated using different tools (e.g., VariTOP, RAPTOR, Copernicus, Mystic)

- Minimize initial mass in low Earth orbit (IMLEO)
- Crewed trip time <700 days
- Perform parametric assessment of Mars parking orbit altitude
- Determine preferred (minimum propellant mass) orbit apoapse and periapse altitudes for selected semi-major axis altitude targets
- Compare against circular orbit altitudes for same semi-major axis target

- Understand effect of parking orbit geometry on lander vehicle mass

Round Trip Earth/Mars Mission

>30 Day Surface Stay

Landing

Launch

Pre-deployed

Mars Lander

500 -> 90,000 km

(Elliptical or Circular Orbits)

Heliocentric Flight

Earth - Mars

Heliocentric Flight

Mars - Earth

Rendezvous/Dock

Of Descent/Ascent Vehicle

And Mars Transfer Vehicle

Mars Crew Transfer Vehicle

Constant Thrust

Power = 6 MW

Efficiency = 60%

Isp = 4000 sec

Mass Return to Earth = 89 mt

Crew Delivery Taxi

(Possible Emergency Return Vehicle)

HEO

30,000 –> 90,000 km

(Circular Orbits)

Crew

Return

Rendezvous/Dock Of Crew Taxi and

Mars Transfer Vehicle

LEO (700 km)

On-orbit Construction

of Transfer Vehicle

Launch of NEP

Transfer Vehicle

Launch Of

Crew Taxi

Launch for

Crew Pickup

Courtesy: Jerry Condon/JSC

Round Trip Earth/Mars Mission

- Spiral NEP Mars transfer vehicle from LEO (700 km) to selected crew transfer orbit (flight crew not onboard)
- Crew taxi launches from ground to Mars transfer vehicle (30,000 – 90,000 km)
- Crewed mission begins with crew transferred to Mars transfer vehicle above the trapped proton radiation belt
- Avoids crew spiral through proton radiation belt
- Crew will, however, spiral through the larger trapped electron belt

- Crewed mission begins with crew transferred to Mars transfer vehicle above the trapped proton radiation belt
- Mars transfer vehicle spirals from crew transfer orbit to heliocentric orbit targeted to Mars
- Mars transfer vehicle transitions from heliocentric space to selected Mars parking orbit (semi-major axis) altitude target (500-90,000 km)
- Mars surface stay (>30 days)
- After surface mission complete, Mars transfer vehicle spirals from Mars parking orbit (500-90,000 km) to heliocentric space targeted to Earth return
- Mars transfer vehicle transitions from heliocentric space to original crew transfer orbit at Earth (30,000 – 90,000 km) for crew pick-up with crew taxi

Earth-Mars Trajectory Analysis Sensitivity StudyExploration Study 1 Follow-on(Three week Quick Study preliminary results)

Melissa L. McGuire

Robert D. Falck

NASA Glenn Research Center

7820 / Systems Analysis Branch

February 28, 2003 (Updated March 3, 2002)

Earth-Mars Trajectory Analysis Sensitivity Study

- Trajectory Analysis Methods
- Trajectory Sensitivity Study Analysis Methods
- Point design case Data and Trajectory Plots
- Sensitivity Study results
- IMLEO and Total trip time as a function of Mars/Earth orbital altitudes
- Table of raw data for sensitivity study

System Assumptions

Power: 6 MW

Specific Impulse (Isp): 4000 sec

Thruster efficiency: 60%

Tankage Fraction: 5%

Mission Assumptions

Mass returned to Earth: 89 mt

Launch Date: 2026

Stay time in Mars space: approx 90 days

Resulted in stay times at Mars in orbit from 37 to 77 days

Mission Total Trip time goal: 700 days

Limiting Orbit Assumptions (for sensitivity trade)

Earth departure orbit altitude : LEO of 700 km

Earth return orbit altitude: vary between 30,000 - 90,000 km

Mars parking orbit altitude: vary between LMO of 500 km and aerosynch

Earth-Mars Trajectory Analysis Sensitivity Study

Varitop, JPL low thrust trajectory analysis code

Trajectories contain spiral escape at Earth, spiral capture/escape at Mars, spiral capture into Earth orbit upon return

Set the final mass at Earth return to 89 mt

Set launch date guess to generate a 2026 launch opportunity

Earth orbits modeled as circular

No constraints on heliocentric orbit proximity to Sun

No propellant allotted for Mars orbit operations (eccentricity, inclination, etc. corrections)

Four bookend point design cases used Mars stay times of 40 and 70 days for low and high Mars parking Orbit altitude cases respectively

These stay times allow for approximately 90 days in Mars vicinity.

More refined Mars stay time choices in sensitivity cases

Earth-Mars Trajectory Analysis Sensitivity Study

First: Ran a series of Mars parking orbit altitudes from 500 to 17,200 km

Second: For each Mars parking orbit, ran a series of Earth return orbits from 30,000 km to 90,000 km altitude

For Each trajectory

Refined guess for stay time in Mars orbit such that the sum of stay time plus spiral capture time and spiral escape time approximately 90 days

Start from a 700 km LEO departure orbit altitude

The NEP vehicle flies the whole trajectory from LEO to Earth return capture

Total trajectory time includes the spiral from LEO to the high earth orbit altitude (I.e., crew delivery altitude) through Earth escape

Earth-Mars Trajectory Analysis Sensitivity Study

Point Design Assumptions:

Earth Departure Orbit: 700 km altitude

Earth Return Orbit: 30,000 km altitude

Mars Parking Orbit: 500 km altitude

Stay Time in Mars Orbit: 40 days

Total Trip time includes LEO to high Earth orbit spiral time

Point Design Result Highlights (see Table for further details)

IMLEO: 303.7 mt

Total trip time (with Earth spirals): 744.8 days

Earth spiral out/in trip time: 110.7 / 9.6 days

Earth spiral out/in propellant cost: 44.5 / 3.9 mt

Mars spiral in/out trip time: 28.4 / 26.3 days

Mars spiral in/out propellant cost: 11.4 / 10.6 mt

Time in Mars Vicinity: 94.7 days

Closest approach of trajectory to Sun: 0.39 AU

Earth-Mars Trajectory Analysis Sensitivity Study

Earth-Mars Trajectory Analysis Sensitivity Study

- Mission Assumptions:
- Earth Departure Orbit: 700 km altitude
- Earth Return Orbit: 30,000 km altitude
- Mars Parking Orbit: 500 km altitude
- Stay Time in Mars Orbit: 40 days
- System Assumptions
- Power: 6 MW
- Specific Impulse (Isp): 4000 sec
- Thruster efficiency: 60%
- Tankage Fraction: 5%

Escape Earth

Spiral for 110.7 days

November 1, 2026

Mass after spiral: 259.1 mt

Close Approach to Sun

Distance ~ 0.39 AU

Earth

Begin Spiral Capture at Mars

June 27, 2027

Mass before spiral: 183.5 mt

Sun

Start at 700 km Earth orbit altitude

July 13, 2026

Initial Mass: 303.7 mt

Mercury

Finish capture at Mars

July 25, 2027

Spiral for 28.4 days

Capture into 500 km orbit

Mass after spiral: 172.1 mt

Escape Mars

Spiral for 26.3 days

September 30, 2027

Mass after spiral: 161.5 mt

Capture at Earth

July 27, 2028

Orbit altitude 30,000 km

Spiral for 9.6 days to capture

Mass after spiral: 89 mt

Mars

Stay time 40 days in Mars orbit

Begin Spiral Escape of Mars

September 3, 2027

Begin Spiral at Earth return

July 17, 2028

Mass before spiral: 92.9 mt

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Point Design Assumptions:

Earth Departure Orbit: 700 km altitude

Earth Return Orbit: 90,000 km altitude

Mars Parking Orbit: 16,700 km altitude

Stay Time in Mars Orbit: 70 days

Total Trip time includes LEO to high Earth orbit spiral time

Point Design Result Highlights (see Table for further details)

IMLEO: 271.6 mt

Total trip time (includes Earth spirals): 692.9 days

Earth spiral out/in trip time: 98.5 / 2.1 days

Earth spiral out/in propellant cost: 40 / 0.86 mt

Mars spiral in/out trip time: 6.23/ 6.06 days

Mars spiral in/out propellant cost: 2.5 / 2.4 mt

Time in Mars Vicinity: 82.3 days

Closest approach of trajectory to Sun: 0.398 AU

Earth-Mars Trajectory Analysis Sensitivity Study

Earth-Mars Trajectory Analysis Sensitivity Study

- Mission Assumptions:
- Earth Departure Orbit: 700 km altitude
- Earth Return Orbit: 90,000 km altitude
- Mars Parking Orbit: 16,700 km altitude
- Stay Time in Mars Orbit: 70 days
- System Assumptions
- Power: 6 MW
- Specific Impulse (Isp): 4000 sec
- Thruster efficiency: 60%
- Tankage Fraction: 5%

Escape Earth

Spiral for 98.5 days

November 7, 2026

Mass after spiral: 232.0 mt

Close Approach to Sun

Distance ~ 0.39 AU

Earth

Begin Spiral Capture at Mars

June 20, 2027

Mass before spiral: 160.8 mt

Start at 700 km Earth orbit altitude

July 31, 2026

Initial Mass: 271.6 mt

Sun

Mercury

Finish capture at Mars

July 27, 2027

Spiral for 6.3 days

Capture into 16,700 km orbit

Mass after spiral: 158.3 mt

Escape Mars

Spiral for 6.1 days

Sept. 11, 2027

Mass after spiral: 155.9 mt

Capture at Earth

June 23, 2028

Orbit altitude 90,000 km

Spiral for 2.1 days to capture

Mass after spiral: 89 mt

Mars

Stay time 70 days in Mars orbit

Begin Spiral Escape of Mars

Sept. 5, 2027

Begin Spiral at Earth return

July 21, 2028

Mass before spiral: 89.6 mt

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Earth-Mars Trajectory Analysis Sensitivity Study

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Earth Departure Orbit: 700 km altitude

Earth Return Orbit: vary from 30,000 to 90,000 km altitude

Mars Parking Orbit: vary from 500 to 17,200 km altitude

Stay Time in Mars Orbit: calculated to sum time in Mars vicinity to approximately 90 days

Resulted in stay times at Mars in orbit from 37 to 77 days

Total Trip time includes spiral time from LEO to high Earth orbit

Earth-Mars Trajectory Analysis Sensitivity Study

Earth-Mars Trajectory Analysis Sensitivity Study

305

Mars Orbit

Mars Stay: 37.0 days

Altitudes

Mars Spiral: 54.5 days

17200km

10000km

300

Mars Stay: 37.0 days

Mars Spiral: 53.6 days

5000km

Mars Stay: 37.0 days

Mars Spiral: 53.0 days

500km

Mars Stay: 37.0 days

Mars Spiral: 52.7 days

295

Mars Stay: 37.0 days

Mars Spiral: 52.4 days

Mars Stay: 60.0 days

290

Mars Spiral: 30.6 days

IMLEO (mt)

Mars Stay: 60.0 days

Mars Spiral: 30.1 days

Mars Stay: 70.0 days

Mars Stay: 60.0 days

285

Mars Spiral: 20.4 days

Mars Spiral: 29.8 days

Mars Stay: 60.0 days

Mars Spiral: 29.6 days

Mars Stay: 70.0 days

Mars Stay: 37.0 days

Mars Spiral: 20.0 days

Mars Spiral: 29.4 days

Mars Stay: 70.0 days

Mars Spiral: 19.8 days

280

Mars Stay: 70.0 days

Mars Spiral: 19.7 days

Mars Stay: 70.0 days

Mars Stay: 77.0 days

Mars Spiral: 19.6 days

Mars Spiral: 13.0 days

Mars Stay: 77.0 days

275

Mars Spiral: 12.8 days

Mars Stay: 77.0 days

Mars Spiral: 12.6 days

Mars Stay: 77.0 days

Mars Spiral: 12.5 days

Mars Stay: 77.0 days

Mars Spiral: 12.4 days

270

30000

40000

50000

60000

70000

80000

90000

Earth Departure/Return Orbit Altitude (km)

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Earth-Mars Trajectory Analysis Sensitivity Study

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Earth-Mars Trajectory Analysis Sensitivity Study

Courtesy:

Melissa McGuire/GRC

Rob Falck/GRC

Missions of 700 round trip are possible with limits on Earth and Mars orbit altitude choices

Total trip time does not equal total crew time

Note: The astronauts will ascend to the NEP vehicle once it’s in the high earth altitude via a crew taxi

Trade studies needed to evaluate choice of Mars parking orbit with respect to Ascent/Descent vehicle versus NEP vehicle performance

Note: Appendix D provides some preliminary data

Further analysis needed to evaluate proximity to Sun on return leg

Earth-Mars Trajectory Analysis Sensitivity Study

Study 2Sun-Earth L2 Libration Point (SE-L2) Mission

Sun-Earth Libration Point (L2) Mission

- Satellite constellation deploy/maintenance mission
- Also, dress rehearsal for Mars mission

- Single vehicle - NEP Mars transfer vehicle
- No rendezvous at SE-L2

- Target => SE-L2
- Use same vehicle specifications as last year study for Mars transfer vehicle
- Power = 6 Mw
- Engine efficiency = 0.6
- Isp = 4000 sec
- No thrust vector turning constraints

- Final mass target (back at Earth) =89mt
- Mission
- Opportunity independent - selectable stay time at SE-L2 (independent of Earth departure time)
- Crew transfer altitude designed to keep crew out of trapped proton radiation belt

Sun-Earth Libration Point (L2) Mission

SE-L2 Operations

Sun-Earth L2 Libration Point (SE-L2)

Mars Crew Transfer Vehicle

Constant Thrust

Power = 6 MW

Efficiency = 60%

Isp = 4000 sec

Mass Return to Earth = 89 mt

Trans SE-L2 Flight

Trans-Earth Flight

Crew Delivery Taxi

(Possible Emergency Return Vehicle)

HEO

30,000 –> 90,000 km

(Circular Orbits)

Crew

Return

Rendezvous/Dock Of Crew Taxi and

Mars Transfer Vehicle

LEO (700 km)

On-orbit Construction

of Transfer Vehicle

Launch of NEP

Transfer Vehicle

Launch Of

Crew Taxi

Launch for

Crew Pickup

Courtesy: Jerry Condon / JSC/EG5

Sun-Earth Libration Point (L2) Mission

- Spiral NEP ‘Mars’ transfer vehicle from LEO (700 km) to selected crew transfer orbit (flight crew not onboard)
- Note: The Mars transfer vehicle is used for this mission to Sun-Earth L2 (SE-L2)
- In addition to meeting planned objectives, the SE-L2 mission could provide a proving ground for future Mars missions

- Note: The Mars transfer vehicle is used for this mission to Sun-Earth L2 (SE-L2)
- Crew taxi launches from ground to Mars transfer vehicle (30,000 – 90,000 km)
- Crewed mission begins with crew transferred to Mars transfer vehicle above the trapped proton radiation belt
- Avoids crew spiral through proton radiation belt
- Crew will, however, spiral through the larger trapped electron belt

- Crewed mission begins with crew transferred to Mars transfer vehicle above the trapped proton radiation belt
- Mars transfer vehicle spirals from crew transfer orbit to SE-L2
- Variable stay time at L2
- Mars transfer vehicle returns crew from SE-L2 to original crew transfer orbit at Earth (30,000 – 90,000 km) for crew pick-up with crew taxi

Sun-Earth Libration Point (L2) Mission

- Trajectory tool used: Copernicus
- Multi-body, multi-spacecraft, continuous thrust trajectory tool in development at University of Texas – Center for Space Research

- Mission - trajectories were solved backwards (from end of mission to beginning) in order to determine required IMLEO needed to conclude mission with an 89 mt mass
- Mission segments:
- Return trip from SE-L2 to crew transfer altitude (30,000 – 90,000 km)
- Outbound trip from 100,000 km to SE-L2
- Spiral up from 700 km initial circular Earth parking orbit to 100,000 km circular orbit
- Mass matching performed for the vehicle at 100,000 km altitude

- Mission segments:

Sun-Earth Libration Point (L2) Mission

Sun-Earth Libration Point (L2) Mission

- Complete RAPTOR mission set
- Compare and contrast results with VariTOP

- Review Mars parking orbit parametric study
- Evaluate sudden change in eccentricity at 38,000 km altitude range

Appendices

Appendix A

Mars Arrival Parking Orbit Analysis

Earth-Mars Round Trip Mission

Comparison of Elliptical vs. Circular Mars Parking Orbit Arrival

Kyle Brewer / JSC/EG5

March 3, 2003

Mars Arrival Parking Orbit Analysis

- Provide a comparison of insertion into Circular vs. Elliptical orbits at Mars based on a state vector from a fully integrated roundtrip mission provided by JPL

Mars Arrival Parking Orbit Analysis

- Same Vehicle specifications as previous study
- The JPL mission is optimized for the following roundtrip mission:
- Depart 30,000 km Earth orbit
- Arrive/Stay Depart Aerosynchronous (17,048 km alt) orbit
- Arrive 30,000 km Earth orbit

- Initial state vector and mass taken from beginning of Mars approach burn (see next slide)
- Given that the state and mass are not optimized for the variety of orbits analyzed, the resulting data should be considered for comparative purposes only.

Mars Arrival Parking Orbit Analysis

Initial State taken from this point

Mars Arrival Parking Orbit Analysis

- Trajectory tool used: Copernicus
- Multi-body, multi-spacecraft, continuous thrust trajectory tool in development at University of Texas – Center for Space Research

- Trajectories to circular orbits were computed by specifying the desired orbit radius and constraining the eccentricity to 0.0 and solving for minimum thrusting time
- Optimum eccentricity orbits were determined by holding only the desired Semi-Major Axis constant and solving for minimum thrusting time to meet that SMA constraint

Mars Arrival Parking Orbit Analysis

Mars Arrival Parking Orbit Analysis

SMA = 30000 km

SMA = 39600 km

SMA = 42000 km

Mars Arrival Parking Orbit Analysis

- A large jump in optimum eccentricity is seen around the target SMA of 39,000 km
- This is the target about which the powered trajectory makes it’s first complete pass around the planet

(SMA shown is an altitude)

Mars Arrival Parking Orbit Analysis

Appendix B

Mars Parking Orbit Lifetime

Carlos Westhelle / EG5

March 3, 2003

Mars Parking Orbit Lifetime

- Current Mars ascent vehicle targeted to 200 km temporary parking orbit
- Off-nominal situations (e.g. failure of subsequent engine firing) may require extended stay in this orbit
- This lifetime study takes a quick look at the parking orbit lifetime as a function of altitude range (130-200 km) for a range of possible vehicle ballistic numbers (150-1500 kg/m2)

Mars Parking Orbit Lifetime

- STK-Astrogator was used to propagate the vehicle with a Mars GRAM atmosphere model
- Orbit was propagated until it decayed to a 125 km altitude (Mars entry interface) up to a maximum time cutoff of 365 days
- For orbit propagations reaching this 365 day limit, the resulting orbit altitudes are noted on the plot on the next slide

Mars Parking Orbit Lifetime

Courtesy: Carlos Westhelle / JSC-EG5

Mars Parking Orbit Lifetime

- A 200 km circular Mars parking orbit provides sufficient time (> 365 days) for an extended stay for a worst-case ballistic number (i.e., 150 kg/m2)
- Note: For this case the vehicle will decay to Mars entry interface (125 km) in approximately another 40 days

Appendix C

Integrated Reference Mission – JPL

Greg Whiffen/JPL

February 23, 2003

- Single end to end multi-body integrated trajectory using Mystic
- Trajectory characteristics:
- Start escape spiral at 30,000 km altitude Earth orbit, 224 metric tons, September 8, 2026
- Escape Earth, 209.9 metric tons, October 24, 2026
- Capture Mars-begin spiral, 178.1 metric tons,July 18, 2027
- Areosynchronous orbit 40 days, 173.3 metric tons, July 30 through Sept 8, 2027
- Mars escape, 171.4 metric tons, September 19, 2027
- Earth capture, 104.1 metric tons, July 10, 2028
- Earth 30,000 km altitude orbit, 97.6 metric tons, July 26, 2028

- Vehicle characteristics:
- Power = 6 MW, Efficiency = 60%, Isp = 4000 seconds

- Trajectory results:
- Total flight time is 687 days from 30,000 km altitude Earth orbit to a return 30,000 km altitude Earth orbit
- Time spent in low mars orbit is 40 days.
- Dry mass with tankage is 97.567 metric tons
- Total propellant used is 126.433 metric tons
- 5% tankage is 6.322 metric tons
- Net Mass without tankage 91.245 metric tons

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen / JPL

Appendix D

Effects of Parking Orbit Geometry on Mars Lander Mass

Dave Lee JSC/EG5

March 3, 2003

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

- Comparison of lander mass trends for circular vs. elliptical orbits
- Payload mass cases based on:
- Previous Dual Lander Study
- JSC/EX/Jim Geffre 6 crew/30 day case
- Light descent payload case for illustration

- Delivery method not considered
- Delivery method would amplify mass trends
- No periapse raise after aerobrake budgeted
- High ellipse more suited to aerobrake delivery

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Drop periapse for aerobraking

1

Parking

Orbit

Parking

Orbit

Descent

Ascent

Raise orbit to PO periapse

Deorbit

Circularize in 300 X 300 km

4

2

3

Ascent to

200 X 200 km

Entry, Descent, and Landing

1

5

2

Aerobraking

3

Raise orbit to PO apoapse

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Descent/Ascent Stack

- Masses:
- Descent Only Payload = 15314 kg
- Ascent Payload (w/ crew) = 2624 kg
- 6 Crew (93 kg each) = 558 kg total
- Aeroshell mass 10% of total vehicle mass

- Delta-V’s:
- Terminal descent = 632 m/s
- Ascent to 200 km circ = 3900 m/s
- Rendezvous = 45 m/s

- Single stage and two stage ascent modeled (same delta-V)
- Stage Mass fractions calculated per historical model
- except terminal descent stage (Mass Fraction = 0.58)

- Specific Impulse for all stages 379 s

Ascent Payload

Ascent Stage

Descent Payload

Descent Stage

Circ/Deorbit Stage

Aeroshell

Figure intended to show payloads and staging order only.

No relative scale should be inferred.

Stage location and orientation should not be inferred.

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

110000

Dual Lander:

Single Stage Ascent

100000

Circular Orbits

20000 km periapse

10000 km periapse

90000

34%

5000 km periapse

80000

Vehicle Mass (kg)

2000 km periapse

70000

400 km periapse

60000

50000

40000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

110000

Dual Lander:

Two Stage Ascent

100000

90000

80000

Circular Orbits

Vehicle Mass (kg)

20000 km periapse

70000

10000 km periapse

28%

5000 km periapse

60000

2000 km periapse

400 km periapse

50000

40000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Descent/Ascent Stack

- Masses:
- Descent Only Payload = 17266.8 kg
- Ascent Payload (w/ crew) = 5345.5 kg
- 6 Crew (82 kg each) = 492 kg total
- Aeroshell mass 14% of total vehicle mass

- Delta-V’s:
- Terminal descent = 632 m/s
- Ascent to 200 km circ = 3931 m/s
- Rendezvous = 45 m/s

- Single stage and two stage ascent modeled (same delta-V)
- Stage Mass fractions calculated per historical model
- except terminal descent stage (Mass Fraction = 0.58)

- Specific Impulse for all stages 379 s

Ascent Payload

Ascent Stage

Descent Payload

Descent Stage

Circ/Deorbit Stage

Aeroshell

Figure intended to show payloads and staging order only.

No relative scale should be inferred.

Stage location and orientation should not be inferred.

*Based on JSC/EX/Jim Geffre design

170000

Geffre 6 crew/30 day:

Single Stage Ascent

160000

Circular Orbits

20000 km periapse

10000 km periapse

150000

35%

140000

5000 km periapse

130000

Vehicle Mass (kg)

120000

2000 km periapse

110000

400 km periapse

100000

90000

80000

70000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Courtesy: Dave Lee/JSC

170000

Geffre 6 crew/30 day:

Two Stage Ascent

160000

150000

140000

130000

Circular Orbits

20000 km periapse

Vehicle Mass (kg)

120000

30%

10000 km periapse

110000

5000 km periapse

100000

2000 km periapse

90000

400 km periapse

80000

70000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Courtesy: Dave Lee/JSC

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Descent/Ascent Stack

- Masses:
- Descent Only Payload = 500 kg
- Ascent Payload (w/ crew) = 5345.5 kg
- 6 Crew (82 kg each) = 492 kg total
- Aeroshell mass 10% of total vehicle mass

- Delta-V’s:
- Terminal descent = 632 m/s
- Ascent to 200 km circ = 3931 m/s
- Rendezvous = 45 m/s

- Single stage and two stage ascent modeled (same delta-V)
- Stage Mass fractions calculated per historical model
- except terminal descent stage (Mass Fraction = 0.58)

- Specific Impulse for all stages 379 s

Ascent Payload

Ascent Stage

Descent Payload

Descent Stage

Circ/Deorbit Stage

Aeroshell

Figure intended to show payloads and staging order only.

No relative scale should be inferred.

Stage location and orientation should not be inferred.

130000

Light Descent:

Single Stage Ascent

120000

20000 km periapse

Circular Orbits

10000 km periapse

110000

37%

5000 km periapse

100000

90000

2000 km periapse

Vehicle Mass (kg)

80000

400 km periapse

70000

60000

50000

40000

30000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Courtesy: Dave Lee/JSC

130000

Light Descent:

Two Stage Ascent

120000

110000

100000

90000

Circular Orbits

20000 km periapse

Vehicle Mass (kg)

80000

10000 km periapse

33%

5000 km periapse

70000

2000 km periapse

60000

400 km periapse

50000

40000

30000

0

5000

10000

15000

20000

25000

30000

35000

Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

Courtesy: Dave Lee/JSC

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

- Elliptical orbits offer major mass advantages for large SMAs as compared to circular orbits
- Up to 37% lander mass savings for some large SMA cases
- Most pronounced for Single Stage Ascent (but still significant for Two Stage)
- If aerobraking delivery were desired, elliptical orbits would offer additional mass advantage

- Two stage ascent offers major mass advantages for high orbits
- Over 25% lander mass difference for some higher orbit cases
- Less than 10% for lowest orbits
- Most pronounced for Light Descent case and Circular orbits

- If we consider the mass impact of delivering the lander/ascent vehicle to the Mars parking orbit, these mass trends would be amplified

Appendix E

Van Allen Radiation Belt Data

Trapped Proton Belt Data

Jerry Condon / JSC/EG5

Van Allen Radiation Belt (Trapped Proton) Data

Courtesy: Jerry Condon/JSC

Van Allen Radiation Belt (Trapped Proton) Data

Courtesy: Jerry Condon/JSC