Preliminary Earth-Mars Artificial-G NEP Architecture

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 / /

Inter-center Study Team
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
3- 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) 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

Study 1 - Round Trip Earth/Mars Mission
Contents 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 1 Round Trip Earth/Mars Mission

Round Trip Earth/Mars Mission
Assumptions 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

Goals 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 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

Mission Overview >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 LEO (700 km) Rendezvous/Dock Of Crew Taxi and Mars Transfer Vehicle On-orbit Construction of Transfer Vehicle Launch of NEP Transfer Vehicle Launch Of Crew Taxi Launch for Crew Pickup Courtesy: Jerry Condon/JSC

Mission Overview 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

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, (Updated March 3, 2002)

Report out of Quick Turnout Study
Earth-Mars Trajectory Analysis Sensitivity Study Report out of Quick Turnout 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

Mission and System Assumptions
Earth-Mars Trajectory Analysis Sensitivity Study Mission and System Assumptions 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 km Mars parking orbit altitude: vary between LMO of 500 km and aerosynch

Trajectory Analysis Methods
Earth-Mars Trajectory Analysis Sensitivity Study Trajectory 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

Trajectory Sensitivity Analysis Methodology
Earth-Mars Trajectory Analysis Sensitivity Study Trajectory 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 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 500/30,000 Trajectory Point Design
Earth-Mars Trajectory Analysis Sensitivity Study Earth-Mars 500/30,000 Trajectory Point Design 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: mt Total trip time (with Earth spirals): days Earth spiral out/in trip time: / 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 2026 (Earth Return 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% Escape Earth Spiral for days November 1, 2026 Mass after spiral: mt Distance ~ 0.39 AU Close Approach to Sun Earth Begin Spiral Capture at Mars June 27, 2027 Mass before spiral: mt Sun Start at 700 km Earth orbit altitude July 13, 2026 Initial Mass: mt Mercury Finish capture at Mars July 25, 2027 Spiral for days Capture into 500 km orbit Mass after spiral: mt Escape Mars Spiral for 26.3 days September 30, 2027 Mass after spiral: 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

Earth-Mars 16700/90000 Trajectory Point Design
Earth-Mars Trajectory Analysis Sensitivity Study Earth-Mars 16700/90000 Trajectory Point Design 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: mt Total trip time (includes Earth spirals): 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: AU

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% Escape Earth Spiral for 98.5 days November 7, 2026 Mass after spiral: mt Distance ~ 0.39 AU Close Approach to Sun Earth Begin Spiral Capture at Mars June 20, 2027 Mass before spiral: mt Start at 700 km Earth orbit altitude July 31, 2026 Initial Mass: mt Sun Mercury Finish capture at Mars July 27, 2027 Spiral for 6.3 days Capture into 16,700 km orbit Mass after spiral: mt Escape Mars Spiral for 6.1 days Sept. 11, 2027 Mass after spiral: 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 2026 Point Design Bookend Cases Data Table
Earth-Mars Trajectory Analysis Sensitivity Study Earth Mars 2026 Point Design Bookend Cases Data Table Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Sensitivity Analysis Assumptions
Earth-Mars Trajectory Analysis Sensitivity Study Sensitivity 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

IMLEO vs. Earth Return Orbit Altitude
Earth-Mars Trajectory Analysis Sensitivity Study IMLEO vs. Earth Return Orbit Altitude 305 Mars Orbit Mars Stay: 37.0 days Mars Spiral: 54.5 days Altitudes 17200km 300 Mars Stay: 37.0 days 10000km 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 285 Mars Stay: 70.0 days Mars Stay: 60.0 days 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 280 Mars Spiral: 19.8 days Mars Stay: 70.0 days Mars Spiral: 19.7 days Mars Stay: 77.0 days Mars Stay: 70.0 days Mars Spiral: 13.0 days Mars Spiral: 19.6 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

Total and Crewed Mission Time vs. Earth Return Orbit Radius
Earth-Mars Trajectory Analysis Sensitivity Study Total and Crewed Mission Time vs. Earth Return Orbit Radius Courtesy: Melissa McGuire/GRC, Rob Falck/GRC

Low Thrust NEP Trajectory Trade Space Raw Data
Earth-Mars Trajectory Analysis Sensitivity Study Low Thrust NEP Trajectory Trade Space Raw Data Courtesy: Melissa McGuire/GRC Rob Falck/GRC

Observations 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

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

Sun-Earth Libration Point (L2) Mission
Assumptions 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

Mission Overview 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 LEO (700 km) Rendezvous/Dock Of Crew Taxi and Mars Transfer Vehicle On-orbit Construction of Transfer Vehicle Launch of NEP Transfer Vehicle Launch Of Crew Taxi Launch for Crew Pickup Courtesy: Jerry Condon / JSC/EG5

Mission Overview 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

Study Methodology 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

IMLEO and Trip Time vs. Crew Altitude
Sun-Earth Libration Point (L2) Mission IMLEO and Trip Time vs. Crew Altitude

Tabular Trajectory Data
Sun-Earth Libration Point (L2) Mission Tabular Trajectory Data

Complete RAPTOR mission set Review Mars parking orbit parametric study
Future Work 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
Purpose 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

Assumptions 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.

Initial State from JPL Initial State taken from this point

Trajectory tool used: Copernicus
Mars Arrival Parking Orbit Analysis Methodology 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

Prop Usage for Circular and Opt. Ecc Orbits
Mars Arrival Parking Orbit Analysis Prop Usage for Circular and Opt. Ecc Orbits

Optimum Eccentricity and Ha/Hp
Mars Arrival Parking Orbit Analysis Optimum Eccentricity and Ha/Hp

Observations 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 = km SMA = km SMA = km (SMA shown is an altitude)

Tabular Trajectory Data
Mars Arrival Parking Orbit Analysis Tabular Trajectory Data

Mars Parking Orbit Lifetime Carlos Westhelle / EG5 March 3, 2003
Appendix B Mars Parking Orbit Lifetime Carlos Westhelle / EG5 March 3, 2003

Orbit Lifetime at Mars - Introduction
Mars Parking Orbit Lifetime Orbit 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 ( km) for a range of possible vehicle ballistic numbers ( kg/m2)

Orbit Lifetime at Mars - Methodology
Mars Parking Orbit Lifetime Orbit Lifetime at Mars - Methodology 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
Orbit Lifetime at Mars Courtesy: Carlos Westhelle / JSC-EG5

Orbit Lifetime at Mars – Observations
Mars Parking Orbit Lifetime Orbit 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

Integrated Reference Mission – JPL Greg Whiffen/JPL February 23, 2003
Appendix C Integrated Reference Mission – JPL Greg Whiffen/JPL February 23, 2003

Mission Design and Results
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, metric tons, October 24, 2026 Capture Mars-begin spiral, metric tons,July 18, 2027 Areosynchronous orbit 40 days, metric tons, July 30 through Sept 8, 2027 Mars escape, metric tons, September 19, 2027 Earth capture, 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 metric tons Total propellant used is metric tons 5% tankage is metric tons Net Mass without tankage metric tons

Courtesy: Greg Whiffen/JPL

Courtesy: Greg Whiffen / JPL

Effects of Parking Orbit Geometry on Mars Lander Mass
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 Lander Mass
Effects of Mars Parking Orbit Geometry on Mars Lander Mass Effects 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 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
Orbital Maneuvers Drop periapse for aerobraking 1 Parking Orbit Parking Orbit Descent Ascent Raise orbit to PO periapse Deorbit 4 Circularize in 300 X 300 km 2 3 Ascent to 200 X 200 km Entry, Descent, and Landing 1 5 2 Aerobraking 3 Raise orbit to PO apoapse

Dual Lander Case Descent/Ascent Stack Masses: Delta-V’s:
Effects of Mars Parking Orbit Geometry on Mars Lander Mass Dual Lander Case Descent/Ascent Stack Masses: Descent Only Payload = 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.

Lander Mass vs. Mars Parking Orbit Semi-Major Axis
Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander Mass vs. Mars Parking Orbit Semi-Major Axis 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 5000 10000 15000 20000 25000 30000 35000 Mars Parking Orbit Semi-Major Axis (km)

Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander 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 5000 10000 15000 20000 25000 30000 35000 Mars Parking Orbit Semi-Major Axis (km) Vehicle Mass (kg) 28% Dual Lander: Two Stage Ascent

6 crew/30 day case* (staging is different)
Effects of Mars Parking Orbit Geometry on Mars Lander Mass 6 crew/30 day case* (staging is different) Masses: Descent Only Payload = kg Ascent Payload (w/ crew) = 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 Descent/Ascent Stack 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

Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander 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 70000 80000 90000 100000 110000 120000 130000 140000 150000 160000 170000 5000 10000 15000 20000 25000 30000 35000 Mars Parking Orbit Semi-Major Axis (km) Vehicle Mass (kg) Geffre 6 crew/30 day: Single Stage Ascent 35% Courtesy: Dave Lee/JSC

Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander 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 70000 80000 90000 100000 110000 120000 130000 140000 150000 160000 170000 5000 10000 15000 20000 25000 30000 35000 Mars Parking Orbit Semi-Major Axis (km) Vehicle Mass (kg) Geffre 6 crew/30 day: Two Stage Ascent 30% Courtesy: Dave Lee/JSC

Light Descent Payload Case
Effects of Mars Parking Orbit Geometry on Mars Lander Mass Light Descent Payload Case Descent/Ascent Stack Masses: Descent Only Payload = 500 kg Ascent Payload (w/ crew) = 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.

Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander 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 30000 40000 50000 60000 70000 80000 90000 100000 110000 120000 130000 5000 10000 15000 20000 25000 35000 Mars Parking Orbit Semi-Major Axis (km) Vehicle Mass (kg) Light Descent: Single Stage Ascent 37% Courtesy: Dave Lee/JSC

Effects of Mars Parking Orbit Geometry on Mars Lander Mass Lander 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 30000 40000 50000 60000 70000 80000 90000 100000 110000 120000 130000 5000 10000 15000 20000 25000 35000 Mars Parking Orbit Semi-Major Axis (km) Vehicle Mass (kg) Light Descent: Two Stage Ascent 33% Courtesy: Dave Lee/JSC

Effects of Mars Parking Orbit Geometry on Mars Lander Mass
Conclusions 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

Van Allen Radiation Belt Data
Appendix E Van Allen Radiation Belt Data Trapped Proton Belt Data Jerry Condon / JSC/EG5

Trapped Proton Radiation Belt – Dosage vs. Altitude
Van Allen Radiation Belt (Trapped Proton) Data Trapped Proton Radiation Belt – Dosage vs. Altitude Courtesy: Jerry Condon/JSC

Trapped Proton Radiation Belt - Effect of Orbit Orientation
Van Allen Radiation Belt (Trapped Proton) Data Trapped Proton Radiation Belt - Effect of Orbit Orientation Courtesy: Jerry Condon/JSC

Preliminary Earth-Mars Artificial-G NEP Architecture

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