Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Preliminary
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
1
Inter-center Study Team
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
3- Week Study
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
Deploy/maintenance of satellite constellation
Dress rehearsal for Mars mission
Due date – March 3, 2003
Customers
JSC/ExPO – Kent Joosten, Bret Drake, Brenda Ward, etc.
HQ/Gary Martin
3
Contents
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Study 1 - Round Trip Earth/Mars Mission
Study 2 - Sun-Earth L2 Libration Point Mission
Appendix
A.
B.
C.
D.
Mars Arrival Parking Orbit Analysis
Mars Parking Orbit Lifetime
Integrated Reference Mission
Effects of Parking Orbit Geometry on Mars
Lander Mass
E. Trapped Proton Belt Data
4
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Study 1
Round Trip Earth/Mars Mission
5
Round Trip Earth/Mars Mission
Assumptions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Two vehicles
NEP Mars Transfer Vehicle (MTV)
Object of parametric study
Lander/Ascent Vehicle (LAV)
Previously deployed at Mars
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
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
6
Goals and Objectives
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
7
Round Trip Earth/Mars Mission
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Mission Overview
>30 Day Surface Stay
Landing
Launch
Pre-deployed
Mars Lander
500 -> 90,000 km
(Elliptical or Circular Orbits)
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
Crew
Return
30,000 –> 90,000 km
(Circular Orbits)
Rendezvous/Dock Of Crew Taxi and
Mars Transfer Vehicle
LEO (700 km)
On-orbit Construction
of Transfer Vehicle
Launch of NEP Launch Of
Transfer Vehicle Crew Taxi
Launch for
Crew Pickup
8
Courtesy: Jerry Condon/JSC
Round Trip Earth/Mars Mission
Mission Overview
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
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
9
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Earth-Mars Trajectory Analysis Sensitivity Study
Exploration 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)
10
Earth-Mars Trajectory Analysis Sensitivity Study
Report out of Quick Turnout Study
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
11
Earth-Mars Trajectory Analysis Sensitivity Study
Mission and System Assumptions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
12
Earth-Mars Trajectory Analysis Sensitivity Study
Trajectory Analysis Methods
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
13
Earth-Mars Trajectory Analysis Sensitivity Study
Trajectory Sensitivity Analysis Methodology
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
14
Earth-Mars Trajectory Analysis Sensitivity Study
Earth-Mars 500/30,000 Trajectory Point Design
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
15
Earth-Mars Trajectory Analysis Sensitivity Study
Earth-Mars 2026 (Earth Return 30000 km, Mars Parking
Orbit 500 km) Point Design Trajectory Plot
•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%
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Escape Earth
Spiral for 110.7 days
November 1, 2026
Mass after spiral: 259.1 mt
Close Approach to Sun
Distance ~ 0.39 AU
Begin Spiral Capture at Mars
June 27, 2027
Mass before spiral: 183.5 mt
Earth
Start at 700 km Earth orbit altitude
July 13, 2026
Initial Mass: 303.7 mt
Sun
Mercury
Finish capture at Mars
July 25, 2027
Spiral for 28.4 days
Capture into 500 km orbit
Mass after spiral: 172.1 mt
Mars
Stay time 40 days in Mars orbit
Begin Spiral Escape of Mars
September 3, 2027
Escape Mars
Spiral for 26.3 days
September 30, 2027
Mass after spiral: 161.5 mt
Begin Spiral at Earth return
July 17, 2028
Mass before spiral: 92.9 mt
Capture at Earth
July 27, 2028
Orbit altitude 30,000 km
Spiral for 9.6 days to capture
Mass after spiral: 89 mt
16
Courtesy: Melissa McGuire/GRC, Rob Falck/GRC
Earth-Mars Trajectory Analysis Sensitivity Study
Earth-Mars 16700/90000 Trajectory Point Design
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
17
Earth-Mars Trajectory Analysis Sensitivity Study
Earth-Mars 2026 (90,000 km Earth return, 16,700 km Mars
Parking Orbit)Point Design Trajectory Plot
•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%
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Escape Earth
Spiral for 98.5 days
November 7, 2026
Mass after spiral: 232.0 mt
Close Approach to Sun
Distance ~ 0.39 AU
Begin Spiral Capture at Mars
June 20, 2027
Mass before spiral: 160.8 mt
Earth
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
Mars
Stay time 70 days in Mars orbit
Begin Spiral Escape of Mars
Sept. 5, 2027
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
Begin Spiral at Earth return
July 21, 2028
Mass before spiral: 89.6 mt
18
Courtesy: Melissa McGuire/GRC, Rob Falck/GRC
Earth-Mars Trajectory Analysis Sensitivity Study
Earth Mars 2026 Point Design Bookend Cases Data Table
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
19
Courtesy: Melissa McGuire/GRC, Rob Falck/GRC
Earth-Mars Trajectory Analysis Sensitivity Study
Sensitivity Analysis Assumptions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
20
Earth-Mars Trajectory Analysis Sensitivity Study
IMLEO vs. Earth Return Orbit Altitude
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
305
Mars Orbit
Altitudes
17200km
Mars Stay: 37.0 days
Mars Spiral: 54.5 days
300
Mars Stay: 37.0 days
Mars Spiral: 53.6 days
10000km
Mars Stay: 37.0 days
Mars Spiral: 53.0 days
IMLEO (mt)
295
290
5000km
500km
Mars Stay: 37.0 days
Mars Spiral: 52.7 days
Mars Stay: 37.0 days
Mars Spiral: 52.4 days
Mars Stay: 60.0 days
Mars Spiral: 30.6 days
Mars Stay: 60.0 days
Mars Spiral: 30.1 days
285
Mars Stay: 70.0 days
Mars Spiral: 20.4 days
Mars Stay: 60.0 days
Mars Spiral: 29.8 days
Mars Stay: 70.0 days
Mars Spiral: 20.0 days
280
Mars Stay: 70.0 days
Mars Spiral: 19.8 days
Mars Stay: 77.0 days
Mars Spiral: 13.0 days
Mars Stay: 77.0 days
Mars Spiral: 12.8 days
275
Mars Stay: 77.0 days
Mars Spiral: 12.6 days
Mars Stay: 60.0 days
Mars Spiral: 29.6 days
Mars Stay: 37.0 days
Mars Spiral: 29.4 days
Mars Stay: 70.0 days
Mars Spiral: 19.7 days
Mars Stay: 77.0 days
Mars Spiral: 12.5 days
Mars Stay: 70.0 days
Mars Spiral: 19.6 days
Mars Stay: 77.0 days
Mars Spiral: 12.4 days
270
30000
40000
50000
60000
70000
80000
Earth Departure/Return Orbit Altitude (km)
Courtesy: Melissa McGuire/GRC, Rob Falck/GRC
90000
21
Earth-Mars Trajectory Analysis Sensitivity Study
Total and Crewed Mission Time
vs. Earth Return Orbit Radius
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
22
Courtesy: Melissa McGuire/GRC, Rob Falck/GRC
Earth-Mars Trajectory Analysis Sensitivity Study
Low Thrust NEP Trajectory Trade Space Raw Data
Courtesy:
Melissa McGuire/GRC
Rob Falck/GRC
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
23
Earth-Mars Trajectory Analysis Sensitivity Study
Observations
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
24
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Study 2
Sun-Earth L2 Libration Point (SE-L2) Mission
25
Sun-Earth Libration Point (L2) Mission
Assumptions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
26
Sun-Earth Libration Point (L2) Mission
Mission Overview
Sun-Earth L2 Libration Point (SE-L2)
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
SE-L2 Operations
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
Crew
Return
30,000 –> 90,000 km
(Circular Orbits)
Rendezvous/Dock Of Crew Taxi and
Mars Transfer Vehicle
LEO (700 km)
On-orbit Construction
of Transfer Vehicle
Launch of NEP Launch Of
Transfer Vehicle Crew Taxi
Launch for
Crew Pickup
27
Courtesy: Jerry Condon / JSC/EG5
Sun-Earth Libration Point (L2) Mission
Mission Overview
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
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
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
28
Sun-Earth Libration Point (L2) Mission
Study Methodology
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
29
Sun-Earth Libration Point (L2) Mission
IMLEO and Trip Time vs. Crew Altitude
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Earth-Sun L2 Mission
LEO Mass Requirements and Crew Transfer Times
70.0
Power = 6000 kW
Isp = 4000 sec
Efficiency = 0.60
Initial Mass at 700km LEO
(kg)
123500
123000
65.0
60.0
IMLEO
Crew Transfer Time
122500
55.0
122000
50.0
121500
45.0
121000
40.0
120500
35.0
120000
30000
40000
50000
60000
70000
Crew Transfer Altitude (km)
80000
Roundtrip Crew Transfer Time
(days)
124000
30.0
90000
30
Sun-Earth Libration Point (L2) Mission
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Tabular Trajectory Data
Crew Transfer Altitude
IMLEO
Spiral Time to Boarding Altitude
Outbound Crewed Transfer Time
Inbound Crewed Transfer Time
Total Crewed Transfer Time
km
kg
days
days
days
days
30000
123496
31.0
31.6
28.5
60.1
50000
121775
35.0
26.6
24.5
51.1
70000
120831
37.3
23.9
22.3
46.2
90000
120254
38.8
22.1
20.9
43.0
31
Future Work
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
32
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Appendices
33
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
34
Mars Arrival Parking Orbit Analysis
Purpose
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
35
Mars Arrival Parking Orbit Analysis
Assumptions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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.
36
Mars Arrival Parking Orbit Analysis
Initial State from JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Initial State
taken from
this point
37
Mars Arrival Parking Orbit Analysis
Methodology
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
38
Mars Arrival Parking Orbit Analysis
Prop Usage for Circular and Opt. Ecc Orbits
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Propellent Requirement Comparison for Circular Orbits vs. Optimum Eccentricity
Orbits
Prop usage shown is for final burn arc and spiral down to given SMA
32000
1200
Power = 6000 Kw
Isp = 4000s
Eff = 0.60
Initial state in heliocentric space provided by JPL for 2026
mission oportunity. The JPL data was optimized for a
Mars orbit altitude of 17048 km.
28000
1050
900
26000
750
Total Propellant - Circular Orbits
Total Propellant - Ellipse orbits
24000
600
Prop. Usage Difference
22000
450
20000
300
18000
150
16000
0
10000
20000
30000
40000
50000
60000
70000
80000
90000
0
100000
Orbit Semi-Major Axis - Altitude (km)
39
Prop Usage Difference (kg)
Total Prop Usage (kg)
30000
Mars Arrival Parking Orbit Analysis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Optimum Eccentricity and Ha/Hp
Optimum Orbit Eccentricity vs. Semi-Major Axis
and Corresponding Apoapse and Periapse Altitudes
1
200000
0.9
180000
0.8
160000
0.7
140000
Apoapse Altitude
0.6
120000
0.5
100000
Power = 6000 Kw
Isp = 4000s
Eff = 0.60
Initial state in heliocentric space provided by JPL for
2026 mission oportunity. The JPL data was optimized
for a Mars orbit altitude of 17048 km.
0.4
0.3
0.2
80000
60000
40000
Periapse Altitude
0.1
20000
0
0
10000
20000
30000
40000
50000
60000
70000
80000
90000
0
100000
Semi-Major Axis - Altitude (km)
40
Ha/Hp Altitude (km)
Orbit Eccentricity
Optimal Eccentricity
Mars Arrival Parking Orbit Analysis
Observations
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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 = 42000 km
(SMA shown is an altitude)
SMA = 39600 km
SMA = 30000 km
41
Mars Arrival Parking Orbit Analysis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Circular
Orbit
Mars Orbit Altitude
km
Eccentricity
Final Mass
kg
Mars Approach and Spiral Time days
Propellant
kg
500
0.0008
164865
78.2
31643
2000
0.0002
166975
73.0
29533
5000
0.0003
169369
67.8
27139
10000
0.0013
171410
62.0
25098
20000
0.0049
173314
57.3
23194
30000
0.0017
174316
54.9
22192
Optimal
Eccentricity
Eccentricity
Mars Approach and Spiral Time days
Final Mass
kg
Propellant
kg
Circle cost
kg
0.0438
78.2
164869
31639
4
0.0452
73.0
166977
29531
2
0.1434
67.1
169380
27128
11
0.2156
62.0
171434
25074
24
0.4158
57.2
173379
23129
65
0.4483
54.5
174450
22058
134
Circular
Orbit
Mars Orbit Altitude
km
Eccentricity
Final Mass
kg
Mars Approach and Spiral Time days
Propellant
kg
40000
0.0001
174875
53.5
21633
50000
0.0001
175389
52.2
21119
60000
0.0002
175778
51.2
20730
70000
0.0008
176061
50.5
20447
80000
0.0001
176245
50.1
20263
90000
0.0009
176041.73
50.6
20466
Optimal
Eccentricity
Tabular Trajectory Data
Eccentricity
Mars Approach and Spiral Time days
Final Mass
kg
Propellant
kg
Circle cost
kg
0.7044
53.2
174984
21524
109
0.5652
51.7
175609
20899
220
0.7211
50.3
176173
20335
395
0.7991
49.3
176552
19956
491
0.8463
48.7
176818
19690
573
0.8731
48.2
177013
19495
971
42
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Appendix B
Mars Parking Orbit Lifetime
Carlos Westhelle / EG5
March 3, 2003
43
Mars Parking Orbit Lifetime
Orbit Lifetime at Mars - Introduction
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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)
44
Mars Parking Orbit Lifetime
Orbit Lifetime at Mars - Methodology
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
45
Mars Parking Orbit Lifetime
Orbit Lifetime at Mars
365+ Day Propagation
Initial and Final Altitudes
BN = 150 kg/m2
Init. Orb. = 200 km at 1 yr = 161.9 km
Init. Orb. = 210 km at 1 yr = 199.1 km
BN = 324 kg/m2
Init. Orb. = 190 km at 1 yr = 164.6 km
Init. Orb. = 200 km at 1 yr = 192.3 km
BN = 600 kg/m2
Init. Orb. = 180 km at 1 yr = 150.8 km
Init. Orb. = 190 km at 1 yr = 181.5 km
BN = 900 kg/m2
Init. Orb. = 175 km at 1 yr = 150.3 km
Init. Orb. = 180 km at 1 yr = 168.2 km
BN = 1500 kg/m2
Init. Orb. = 170 km at 1 yr = 152.2 km
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Candidate
Descent/Ascent
Vehicle Design
-Propagation limited to 365 days
-Orbit is considered decayed at 125.0
km altitude.
-CR (coefficient of reflectivity) assumed to
be 0.0 (study shows that CR = 2.0
doesn't change results)
46
Courtesy: Carlos Westhelle / JSC-EG5
Mars Parking Orbit Lifetime
Orbit Lifetime at Mars – Observations
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
47
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Appendix C
Integrated Reference Mission – JPL
Greg Whiffen/JPL
February 23, 2003
48
Mission Design and Results
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
49
Courtesy: Greg Whiffen/JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
51
Courtesy: Greg Whiffen/JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
52
Courtesy: Greg Whiffen/JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
53
Courtesy: Greg Whiffen/JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
54
Courtesy: Greg Whiffen/JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
55
Courtesy: Greg Whiffen / JPL
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Appendix D
Effects of Parking Orbit Geometry on
Mars Lander Mass
Dave Lee JSC/EG5
March 3, 2003
56
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Effects of Mars Parking Orbit Geometry on Lander Mass
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
57
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Orbital Maneuvers
1
Drop periapse
for aerobraking
Parking
Orbit
Parking
Orbit
Descent
Ascent
Raise orbit to
PO periapse
Deorbit
4
Circularize in
300 X 300 km
2
3
5
Ascent to
200 X 200 km
Entry, Descent,
and Landing
1
2
Aerobraking
3
Raise orbit to
PO apoapse
58
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Dual Lander Case
Masses:
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Descent/Ascent Stack
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
Ascent
Payload
Ascent
Stage
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
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 59
should not
be inferred.
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
110000
Dual Lander:
Single Stage Ascent
100000
Circular Orbits
20000 km periapse
10000 km periapse
Vehicle Mass (kg)
90000
34%
5000 km periapse
80000
2000 km periapse
70000
400 km periapse
60000
50000
40000
0
5000
10000
15000
20000
25000
Mars Parking Orbit Semi-Major Axis (km)
30000
60
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
110000
100000
Dual Lander:
Two Stage Ascent
Vehicle Mass (kg)
90000
80000
Circular Orbits
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
Mars Parking Orbit Semi-Major Axis (km)
30000
61
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
6 crew/30 day case*
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
(staging is different)
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
Ascent
Payload
Ascent
Stage
Delta-V’s:
Terminal descent = 632 m/s
Ascent to 200 km circ = 3931 m/s
Rendezvous = 45 m/s
Descent
Payload
Descent
Stage
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
*Based on JSC/EX/Jim Geffre design
Circ/Deorbit
Stage
Aeroshell
Figure intended to show payloads and
staging order only.
No relative scale should be inferred.
62
Stage location and orientation should not
be inferred.
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
170000
160000
150000
Geffre 6 crew/30 day:
Single Stage Ascent
Circular Orbits
20000 km periapse
10000 km periapse
35%
Vehicle Mass (kg)
140000
5000 km periapse
130000
120000
2000 km periapse
110000
400 km periapse
100000
90000
80000
70000
0
5000
10000
15000
20000
25000
Mars Parking Orbit Semi-Major Axis (km)
Courtesy: Dave Lee/JSC
30000
63
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
170000
160000
150000
Geffre 6 crew/30 day:
Two Stage Ascent
Vehicle Mass (kg)
140000
130000
Circular Orbits
20000 km periapse
120000
10000 km periapse
30%
110000
5000 km periapse
100000
2000 km periapse
90000
400 km periapse
80000
70000
0
5000
10000
15000
20000
25000
Mars Parking Orbit Semi-Major Axis (km)
Courtesy: Dave Lee/JSC
30000
64
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Light Descent Payload Case
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
Ascent
Payload
Ascent
Stage
Delta-V’s:
Terminal descent = 632 m/s
Ascent to 200 km circ = 3931 m/s
Rendezvous = 45 m/s
Descent
Payload
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
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 65
should not
be inferred.
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
130000
120000
110000
Light Descent:
Single Stage Ascent
20000 km periapse
Circular Orbits
10000 km periapse
Vehicle Mass (kg)
37%
5000 km periapse
100000
90000
2000 km periapse
80000
400 km periapse
70000
60000
50000
40000
30000
0
5000
10000
15000
20000
25000
Mars Parking Orbit Semi-Major Axis (km)
Courtesy: Dave Lee/JSC
30000
66
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
130000
120000
110000
Light Descent:
Two Stage Ascent
Vehicle Mass (kg)
100000
90000
Circular Orbits
20000 km periapse
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
Mars Parking Orbit Semi-Major Axis (km)
Courtesy: Dave Lee/JSC
30000
67
35000
Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Conclusions
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
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
68
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Appendix E
Van Allen Radiation Belt Data
Trapped Proton Belt Data
Jerry Condon / JSC/EG5
69
Van Allen Radiation Belt (Trapped Proton) Data
Trapped Proton Radiation Belt – Dosage vs. Altitude
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Dose Rate for Circular Orbits
500000
Solar Minimum
0.1" Aluminum Shielding
28.5 deg. inclination
240 hours in each orbit
Dose (rem/yr) (Trapped Protons Only)
450000
400000
350000
300000
250000
200000
150000
100000
50000
0
0.0
2000.0
4000.0
6000.0
8000.0
10000.0
Circular Orbit Altitude (km)
Courtesy: Jerry Condon/JSC
12000.0
14000.0
16000.0
70
Van Allen Radiation Belt (Trapped Proton) Data
Trapped Proton Radiation Belt - Effect of Orbit Orientation
Low Thrust Trajectory Team
GRC, JPL, JSC, MSFC
Radiation Dose for a 400x35790 km Ellipse
Varying Angles Between Perigee and the Ascending Node
60000
Solar Minimum
0.1" Aluminum Shielding
240 hours in each orbit
407x35790 km ellipse
Dose (rem/yr)
(Trapped Protons Only)
50000
40000
30000
20000
28.5 deg. incl.
10000
51.6 deg. incl.
0
-180
-135
-90
-45
0
45
Perigee Angle from Ascending Node (deg)
Courtesy: Jerry Condon/JSC
90
135
180
71