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

1

Low Thrust Trajectory Team

GRC, JPL, JSC, MSFC

Earth-Mars Artificial-G NEP ArchitectureSun-Earth L2 Architecture

3-Week Parametric Trade Study

Presented to JSC/Exploration OfficeMarch 3, 2003

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

Presentation prepared by: Jerry Condon / JSC / EG5 / 281.483.8173 / [email protected]

Preliminary

2


GRC, JPL, JSC, MSFCInter-center Study Team

GRCMelissa McGuire, Rob Falk

JPLJon Sims, Greg Whiffen

JSCJerry Condon, Ellen Braden, Dave Lee, Kyle Brewer, Carlos Westhelle

Jim Geffre

MSFCReginald Alexander, Larry Kos, Kirk Sorensen

3


GRC, JPL, JSC, MSFC3- Week Study

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) missionDeploy/maintenance of satellite constellation

Dress rehearsal for Mars mission

Due date – March 3, 2003

CustomersJSC/ExPO – Kent Joosten, Bret Drake, Brenda Ward, etc.

HQ/Gary Martin

4


GRC, JPL, JSC, MSFCContents

Study 1 - Round Trip Earth/Mars Mission

Study 2 - Sun-Earth L2 Libration Point Mission

AppendixA. Mars Arrival Parking Orbit Analysis

B. Mars Parking Orbit Lifetime

C. Integrated Reference Mission

D. Effects of Parking Orbit Geometry on Mars Lander Mass

E. Trapped Proton Belt Data

5


GRC, JPL, JSC, MSFC

Study 1Round Trip Earth/Mars Mission

6


GRC, JPL, JSC, MSFCAssumptions

Two vehiclesNEP 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 constraintsDetermine vehicle thrust vector steering requirements unconstrained by Artificial Gravity (AG) vehicle configurationsResults may influence AG vehicle configurations

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

Round Trip Earth/Mars Mission

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

Fire BatonArtificial-G NEP

Mars Transfer Vehicle

7


GRC, JPL, JSC, MSFCGoals and Objectives

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 altitudeDetermine 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

8


GRC, JPL, JSC, MSFCMission Overview

LEO (700 km)

Helio

centr

ic F

light

Eart

h -

Mars

Launch

CrewReturn

HEO30,000 –> 90,000 km

(Circular Orbits)

Pre-deployedMars Lander

500 -> 90,000 km(Elliptical or Circular Orbits)

Launch of NEPTransfer Vehicle

Launch OfCrew Taxi

On-orbit Construction of Transfer Vehicle

Launch forCrew Pickup

Landing

Rendezvous/Dock Of Crew Taxi and Mars Transfer Vehicle

Crew Delivery Taxi(Possible Emergency Return Vehicle)

Helio

centric Flig

ht

Mars - E

arth

Rendezvous/DockOf Descent/Ascent VehicleAnd Mars Transfer Vehicle

>30 Day Surface Stay

Mars Crew Transfer VehicleConstant ThrustPower = 6 MWEfficiency = 60%Isp = 4000 secMass Return to Earth = 89 mt



Courtesy: Jerry Condon/JSC

9



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 beltCrew will, however, spiral through the larger trapped electron belt

Mars transfer vehicle spirals from crew transfer orbit to heliocentric orbit targeted to MarsMars 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 returnMars 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


10


GRC, JPL, JSC, MSFC

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)

11


GRC, JPL, JSC, MSFCReport out of Quick Turnout Study

Trajectory Analysis Methods

Trajectory Sensitivity Study Analysis Methods

Point design case Data and Trajectory Plots

Sensitivity Study resultsIMLEO and Total trip time as a function of Mars/Earth orbital altitudes

Table of raw data for sensitivity study

Earth-Mars Trajectory Analysis Sensitivity Study

12


GRC, JPL, JSC, MSFCMission and System Assumptions

System AssumptionsPower: 6 MWSpecific Impulse (Isp): 4000 sec Thruster efficiency: 60%Tankage Fraction: 5%

Mission Assumptions Mass returned to Earth: 89 mtLaunch Date: 2026Stay 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 kmEarth return orbit altitude: vary between 30,000 - 90,000 kmMars parking orbit altitude: vary between LMO of 500 km and aerosynch


13


GRC, JPL, JSC, MSFCTrajectory Analysis Methods

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


14


GRC, JPL, JSC, MSFCTrajectory Sensitivity Analysis Methodology

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 trajectoryRefined 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 altitudeThe 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


15


GRC, JPL, JSC, MSFCEarth-Mars 500/30,000 Trajectory Point Design

Point Design Assumptions: Earth Departure Orbit: 700 km altitudeEarth Return Orbit: 30,000 km altitudeMars Parking Orbit: 500 km altitudeStay Time in Mars Orbit: 40 daysTotal Trip time includes LEO to high Earth orbit spiral time

Point Design Result Highlights (see Table for further details)IMLEO: 303.7 mtTotal trip time (with Earth spirals): 744.8 daysEarth spiral out/in trip time: 110.7 / 9.6 daysEarth spiral out/in propellant cost: 44.5 / 3.9 mtMars spiral in/out trip time: 28.4 / 26.3 daysMars spiral in/out propellant cost: 11.4 / 10.6 mtTime in Mars Vicinity: 94.7 daysClosest approach of trajectory to Sun: 0.39 AU


16


GRC, JPL, JSC, MSFCEarth-Mars 2026 (Earth Return 30000 km, Mars Parking

Orbit 500 km) Point Design Trajectory Plot

Earth

Mars

Sun

Escape EarthSpiral for 110.7 daysNovember 1, 2026Mass after spiral: 259.1 mt

Capture at Earth July 27, 2028Orbit altitude 30,000 kmSpiral for 9.6 days to captureMass after spiral: 89 mt

Begin Spiral at Earth return July 17, 2028Mass before spiral: 92.9 mt

Begin Spiral Capture at Mars June 27, 2027Mass before spiral: 183.5 mt

Finish capture at Mars July 25, 2027Spiral for 28.4 daysCapture into 500 km orbitMass after spiral: 172.1 mt

Stay time 40 days in Mars orbitBegin Spiral Escape of MarsSeptember 3, 2027

Escape MarsSpiral for 26.3 days September 30, 2027Mass after spiral: 161.5 mt

Close Approach to SunDistance ~ 0.39 AU

Start at 700 km Earth orbit altitudeJuly 13, 2026Initial Mass: 303.7 mt

Mercury

•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%

Courtesy: Melissa McGuire/GRC, Rob Falck/GRC


17


GRC, JPL, JSC, MSFCEarth-Mars 16700/90000 Trajectory Point Design

Point Design Assumptions: Earth Departure Orbit: 700 km altitudeEarth Return Orbit: 90,000 km altitudeMars Parking Orbit: 16,700 km altitudeStay Time in Mars Orbit: 70 daysTotal Trip time includes LEO to high Earth orbit spiral time

Point Design Result Highlights (see Table for further details)IMLEO: 271.6 mtTotal trip time (includes Earth spirals): 692.9 daysEarth spiral out/in trip time: 98.5 / 2.1 daysEarth spiral out/in propellant cost: 40 / 0.86 mtMars spiral in/out trip time: 6.23/ 6.06 daysMars spiral in/out propellant cost: 2.5 / 2.4 mtTime in Mars Vicinity: 82.3 daysClosest approach of trajectory to Sun: 0.398 AU


18


GRC, JPL, JSC, MSFCEarth-Mars 2026 (90,000 km Earth return, 16,700 km Mars

Parking Orbit)Point Design Trajectory Plot

Start at 700 km Earth orbit altitudeJuly 31, 2026Initial Mass: 271.6 mt

Earth

Mars

Sun

Escape EarthSpiral for 98.5 daysNovember 7, 2026Mass after spiral: 232.0 mt

Capture at Earth June 23, 2028Orbit altitude 90,000 kmSpiral for 2.1 days to captureMass after spiral: 89 mt

Begin Spiral at Earth return July 21, 2028Mass before spiral: 89.6 mt

Begin Spiral Capture at MarsJune 20, 2027 Mass before spiral: 160.8 mt

Finish capture at MarsJuly 27, 2027 Spiral for 6.3 daysCapture into 16,700 km orbitMass after spiral: 158.3 mt

Stay time 70 days in Mars orbitBegin Spiral Escape of MarsSept. 5, 2027

Escape MarsSpiral for 6.1 daysSept. 11, 2027Mass after spiral: 155.9 mt

Close Approach to SunDistance ~ 0.39 AU

Mercury

•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%



19


GRC, JPL, JSC, MSFCEarth Mars 2026 Point Design Bookend Cases Data Table



20


GRC, JPL, JSC, MSFCSensitivity Analysis Assumptions

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


21


GRC, JPL, JSC, MSFCIMLEO vs. Earth Return Orbit Altitude

270

275

280

285

290

295

300

305

30000 40000 50000 60000 70000 80000 90000

Earth Departure/Return Orbit Altitude (km)

IMLEO

(m

t)

17200km

10000km

5000km

500km

Mars Stay: 37.0 daysMars Spiral: 54.5 days







Mars Stay: 60.0 daysMars Spiral: 29.8 days Mars Stay: 60.0 days

Mars Spiral: 29.6 daysMars Stay: 37.0 daysMars Spiral: 29.4 days








Mars Stay: 77.0 daysMars Spiral: 12.6 days Mars Stay: 77.0 days

Mars Spiral: 12.5 days Mars Stay: 77.0 daysMars Spiral: 12.4 days

Mars Orbit Altitudes



22


GRC, JPL, JSC, MSFCTotal and Crewed Mission Time

vs. Earth Return Orbit Radius



23


GRC, JPL, JSC, MSFCLow Thrust NEP Trajectory Trade Space Raw Data

Courtesy: Melissa McGuire/GRCRob Falck/GRC


24


GRC, JPL, JSC, MSFCObservations

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


25


GRC, JPL, JSC, MSFC

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

26



Satellite constellation deploy/maintenance missionAlso, dress rehearsal for Mars mission

Single vehicle - NEP Mars transfer vehicleNo 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

MissionOpportunity 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

27



LEO (700 km)

Tra

ns

SE-L

2 F

light

CrewReturn

HEO30,000 –> 90,000 km

(Circular Orbits)

Launch of NEPTransfer Vehicle

Launch OfCrew Taxi

On-orbit Construction of Transfer Vehicle

Launch forCrew Pickup

Rendezvous/Dock Of Crew Taxi and Mars Transfer Vehicle

Crew Delivery Taxi(Possible Emergency Return Vehicle)

Tra

ns-E

arth

Flight



Sun-Earth L2 Libration Point (SE-L2)SE-L2 Operations

Courtesy: Jerry Condon / JSC/EG5


28



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


29


GRC, JPL, JSC, MSFCStudy Methodology

Trajectory tool used: CopernicusMulti-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


30


GRC, JPL, JSC, MSFCIMLEO and Trip Time vs. Crew Altitude

Earth-Sun L2 Mission LEO Mass Requirements and Crew Transfer Times

120000

120500

121000

121500

122000

122500

123000

123500

124000

30000 40000 50000 60000 70000 80000 90000

Crew Transfer Altitude (km)

Init

ial

Mas

s at

700

km L

EO

(k

g)

30.0

35.0

40.0

45.0

50.0

55.0

60.0

65.0

70.0

Ro

un

dtr

ip C

rew

Tra

nsf

er T

ime

(day

s)

IMLEO

Crew Transfer Time

Power = 6000 kWIsp = 4000 sec

Efficiency = 0.60


31


GRC, JPL, JSC, MSFCTabular Trajectory Data

Crew Transfer Altitude km 30000 50000 70000 90000

IMLEO kg 123496 121775 120831 120254Spiral Time to Boarding Altitude days 31.0 35.0 37.3 38.8Outbound Crewed Transfer Time days 31.6 26.6 23.9 22.1Inbound Crewed Transfer Time days 28.5 24.5 22.3 20.9Total Crewed Transfer Time days 60.1 51.1 46.2 43.0


32


GRC, JPL, JSC, MSFCFuture Work

Complete RAPTOR mission setCompare and contrast results with VariTOP

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

33


GRC, JPL, JSC, MSFC

Appendices

34


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

35


GRC, JPL, JSC, MSFCPurpose

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


36



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.


37


GRC, JPL, JSC, MSFCInitial State from JPL

Initial State taken from this point


38


GRC, JPL, JSC, MSFCMethodology

Trajectory tool used: CopernicusMulti-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


39


GRC, JPL, JSC, MSFCProp Usage for Circular and Opt. Ecc Orbits

Propellent Requirement Comparison for Circular Orbits vs. Optimum Eccentricity Orbits

Prop usage shown is for final burn arc and spiral down to given SMA

16000

18000

20000

22000

24000

26000

28000

30000

32000

0 10000 20000 30000 40000 50000 60000 70000 80000 90000 100000

Orbit Semi-Major Axis - Altitude (km)

To

tal

Pro

p U

sag

e (k

g)

0

150

300

450

600

750

900

1050

1200

Pro

p U

sag

e D

iffe

ren

ce (

kg)

Total Propellant - Circular Orbits

Total Propellant - Ellipse orbits

Prop. Usage Difference

Power = 6000 KwIsp = 4000sEff = 0.60Initial state in heliocentric space provided by JPL for 2026 mission oportunity. The JPL data was optimized for a Mars orbit altitude of 17048 km.


40


GRC, JPL, JSC, MSFCOptimum Eccentricity and Ha/Hp

Optimum Orbit Eccentricity vs. Semi-Major Axisand Corresponding Apoapse and Periapse Altitudes

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

0 10000 20000 30000 40000 50000 60000 70000 80000 90000 100000

Semi-Major Axis - Altitude (km)

Orb

it E

ccen

tric

ity

0

20000

40000

60000

80000

100000

120000

140000

160000

180000

200000

Ha/

Hp

Alt

itu

de

(km

)

Optimal Eccentricity

Periapse Altitude

Apoapse Altitude

Power = 6000 KwIsp = 4000sEff = 0.60Initial state in heliocentric space provided by JPL for 2026 mission oportunity. The JPL data was optimized for a Mars orbit altitude of 17048 km.


41


GRC, JPL, JSC, MSFCObservations

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 = 39600 kmSMA = 42000 km SMA = 30000 km

(SMA shown is an altitude)


42


GRC, JPL, JSC, MSFCTabular Trajectory Data

Mars Orbit Altitude km 500 2000 5000 10000 20000 30000Eccentricity - 0.0008 0.0002 0.0003 0.0013 0.0049 0.0017Final Mass kg 164865 166975 169369 171410 173314 174316Mars Approach and Spiral Time days 78.2 73.0 67.8 62.0 57.3 54.9Propellant kg 31643 29533 27139 25098 23194 22192

Eccentricity - 0.0438 0.0452 0.1434 0.2156 0.4158 0.4483Mars Approach and Spiral Time days 78.2 73.0 67.1 62.0 57.2 54.5Final Mass kg 164869 166977 169380 171434 173379 174450Propellant kg 31639 29531 27128 25074 23129 22058Circle cost kg 4 2 11 24 65 134

Circ

ular

O

rbit

Opt

imal

E

ccen

tric

ity

Mars Orbit Altitude km 40000 50000 60000 70000 80000 90000Eccentricity - 0.0001 0.0001 0.0002 0.0008 0.0001 0.0009Final Mass kg 174875 175389 175778 176061 176245 176041.73Mars Approach and Spiral Time days 53.5 52.2 51.2 50.5 50.1 50.6Propellant kg 21633 21119 20730 20447 20263 20466

Eccentricity - 0.7044 0.5652 0.7211 0.7991 0.8463 0.8731Mars Approach and Spiral Time days 53.2 51.7 50.3 49.3 48.7 48.2Final Mass kg 174984 175609 176173 176552 176818 177013Propellant kg 21524 20899 20335 19956 19690 19495Circle cost kg 109 220 395 491 573 971

Circ

ular

O

rbit

Opt

imal

E

ccen

tric

ity


43


GRC, JPL, JSC, MSFC

Appendix B

Mars Parking Orbit Lifetime

Carlos Westhelle / EG5

March 3, 2003

44


GRC, JPL, JSC, MSFCOrbit Lifetime at Mars - Introduction

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)


45


GRC, JPL, JSC, MSFCOrbit Lifetime at Mars - Methodology

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


46


GRC, JPL, JSC, MSFC

Orbit Lifetime at Mars

0.0

50.0

100.0

150.0

200.0

250.0

300.0

350.0

400.0

120 130 140 150 160 170 180 190 200 210 220

Initial Orbit Altitude [km]

BN = 150 kg/m^2 BN = 324 kg/m^2 BN = 600 kg/m^2 BN = 900 kg/m^2 BN = 1500 kg/m^2

Mars AscentParking Orbit Target

Orbit Lifetime at Mars

365+ Day Propagation Initial and Final Altitudes

BN = 150 kg/m2

Init. Orb. = 200 km at 1 yr = 161.9 kmInit. Orb. = 210 km at 1 yr = 199.1 km

BN = 324 kg/m2


BN = 600 kg/m2


BN = 900 kg/m2


BN = 1500 kg/m2

Init. Orb. = 170 km at 1 yr = 152.2 km

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)

Courtesy: Carlos Westhelle / JSC-EG5


47


GRC, JPL, JSC, MSFCOrbit Lifetime at Mars – Observations

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


48


GRC, JPL, JSC, MSFC

Appendix C

Integrated Reference Mission – JPL

Greg Whiffen/JPL

February 23, 2003

49


GRC, JPL, JSC, MSFCMission Design and Results

Single end to end multi-body integrated trajectory using MysticTrajectory characteristics:

Start escape spiral at 30,000 km altitude Earth orbit, 224 metric tons, September 8, 2026Escape Earth, 209.9 metric tons, October 24, 2026Capture Mars-begin spiral, 178.1 metric tons,July 18, 2027Areosynchronous orbit 40 days, 173.3 metric tons, July 30 through Sept 8, 2027Mars escape, 171.4 metric tons, September 19, 2027Earth capture, 104.1 metric tons, July 10, 2028Earth 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 orbitTime spent in low mars orbit is 40 days. Dry mass with tankage is 97.567 metric tonsTotal propellant used is 126.433 metric tons5% tankage is 6.322 metric tonsNet Mass without tankage 91.245 metric tons

Courtesy: Greg Whiffen/JPL

51


GRC, JPL, JSC, MSFC


52


GRC, JPL, JSC, MSFC


53


GRC, JPL, JSC, MSFC


54


GRC, JPL, JSC, MSFC


55


GRC, JPL, JSC, MSFC

Courtesy: Greg Whiffen / JPL

56


GRC, JPL, JSC, MSFC

Appendix D

Effects of Parking Orbit Geometry on Mars Lander Mass

Dave Lee JSC/EG5

March 3, 2003

57


GRC, JPL, JSC, MSFCEffects of Mars Parking Orbit Geometry on Lander Mass

Comparison of lander mass trends for circular vs. elliptical orbits

Payload mass cases based on:Previous Dual Lander StudyJSC/EX/Jim Geffre 6 crew/30 day caseLight descent payload case for illustration

Delivery method not consideredDelivery method would amplify mass trendsNo periapse raise after aerobrake budgetedHigh ellipse more suited to aerobrake delivery

Effects of Mars Parking Orbit Geometry on Mars Lander Mass

58


GRC, JPL, JSC, MSFCOrbital Maneuvers

1

2

3

4

5

Drop periapse for aerobraking

Aerobraking

Circularize in 300 X 300 km

Deorbit

Entry, Descent, and Landing

2

3

Raise orbit to PO periapse

Ascent to200 X 200 km

Raise orbit to PO apoapse

1

Descent Ascent

ParkingOrbit

ParkingOrbit


59


GRC, JPL, JSC, MSFCDual Lander Case

Masses:Descent Only Payload = 15314 kgAscent Payload (w/ crew) = 2624 kg6 Crew (93 kg each) = 558 kg totalAeroshell mass 10% of total vehicle mass

Delta-V’s:Terminal descent = 632 m/sAscent to 200 km circ = 3900 m/sRendezvous = 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

Aeroshell

Ascent Stage

Circ/Deorbit Stage

Descent Payload

Descent Stage

Ascent Payload

Descent/Ascent Stack

Figure intended to show payloads and staging order only.

No relative scale should be inferred.Stage location and orientation should not

be inferred.


60


GRC, JPL, JSC, MSFCLander Mass vs. Mars Parking Orbit Semi-Major Axis

400 km periapse

Circular Orbits

10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

40000

50000

60000

70000

80000

90000

100000

110000

0 5000 10000 15000 20000 25000 30000 35000

Mars Parking Orbit Semi-Major Axis (km)

Ve

hic

le M

as

s (

kg

)

Dual Lander:Single Stage Ascent

34%


61



400 km periapse

Circular Orbits

10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

40000

50000

60000

70000

80000

90000

100000

110000

0 5000 10000 15000 20000 25000 30000 35000


Ve

hic

le M

as

s (

kg

)

28%

Dual Lander:Two Stage Ascent


62


GRC, JPL, JSC, MSFC6 crew/30 day case* (staging is different)

Masses:Descent Only Payload = 17266.8 kgAscent Payload (w/ crew) = 5345.5 kg6 Crew (82 kg each) = 492 kg totalAeroshell mass 14% of total vehicle mass





Aeroshell

Ascent Stage

Circ/Deorbit Stage

Descent Payload

Descent Stage

Ascent Payload




be inferred.*Based on JSC/EX/Jim Geffre design


63



400 km periapse

Circular Orbits

10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

70000

80000

90000

100000

110000

120000

130000

140000

150000

160000

170000

0 5000 10000 15000 20000 25000 30000 35000


Ve

hic

le M

as

s (

kg

)

Geffre 6 crew/30 day:Single Stage Ascent

35%

Courtesy: Dave Lee/JSC


64



400 km periapse

Circular Orbits

10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

70000

80000

90000

100000

110000

120000

130000

140000

150000

160000

170000

0 5000 10000 15000 20000 25000 30000 35000


Ve

hic

le M

as

s (

kg

)

Geffre 6 crew/30 day:Two Stage Ascent

30%



65


GRC, JPL, JSC, MSFCLight Descent Payload Case

Masses:Descent Only Payload = 500 kgAscent Payload (w/ crew) = 5345.5 kg6 Crew (82 kg each) = 492 kg totalAeroshell mass 10% of total vehicle mass





Aeroshell

Ascent Stage

Circ/Deorbit Stage

Descent Payload

Descent Stage

Ascent Payload




be inferred.


66



400 km periapse

Circular Orbits10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

30000

40000

50000

60000

70000

80000

90000

100000

110000

120000

130000

0 5000 10000 15000 20000 25000 30000 35000


Ve

hic

le M

as

s (

kg

)

Light Descent:Single Stage Ascent

37%



67



400 km periapse

Circular Orbits

10000 km periapse

2000 km periapse

20000 km periapse

5000 km periapse

30000

40000

50000

60000

70000

80000

90000

100000

110000

120000

130000

0 5000 10000 15000 20000 25000 30000 35000


Ve

hic

le M

as

s (

kg

)

Light Descent:Two Stage Ascent

33%



68


GRC, JPL, JSC, MSFCConclusions

Elliptical orbits offer major mass advantages for large SMAs as compared to circular orbits

Up to 37% lander mass savings for some large SMA casesMost 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 orbitsOver 25% lander mass difference for some higher orbit casesLess than 10% for lowest orbitsMost 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


69


GRC, JPL, JSC, MSFC

Appendix E

Van Allen Radiation Belt DataTrapped Proton Belt Data

Jerry Condon / JSC/EG5

70


GRC, JPL, JSC, MSFCTrapped Proton Radiation Belt – Dosage vs. Altitude

Dose Rate for Circular Orbits

0

50000

100000

150000

200000

250000

300000

350000

400000

450000

500000

0.0 2000.0 4000.0 6000.0 8000.0 10000.0 12000.0 14000.0 16000.0

Circular Orbit Altitude (km)

Do

se

(re

m/y

r)

(Tra

pp

ed P

roto

ns

On

ly)

Solar Minimum0.1" Aluminum Shielding28.5 deg. inclination240 hours in each orbit

Van Allen Radiation Belt (Trapped Proton) Data


71


GRC, JPL, JSC, MSFCTrapped Proton Radiation Belt - Effect of Orbit Orientation

Radiation Dose for a 400x35790 km EllipseVarying Angles Between Perigee and the Ascending Node

0

10000

20000

30000

40000

50000

60000

-180 -135 -90 -45 0 45 90 135 180

Perigee Angle from Ascending Node (deg)

Do

se

(re

m/y

r)

(Tra

pp

ed

Pro

ton

s O

nly

)

51.6 deg. incl.

28.5 deg. incl.

Solar Minimum0.1" Aluminum Shielding240 hours in each orbit407x35790 km ellipse

Van Allen Radiation Belt (Trapped Proton) Data


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

Documents

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