Utilizing Lunar Architecture Transportation Elements for Mars Exploration

Utilizing Lunar Architecture Transportation Elements for Mars Exploration 19 September 2007 Brad St. Germain, Ph.D. Director of Advanced Concepts brad.stgermain@sei.aero 1+770.379.8010 1

Introduction Architecture Overview Interplanetary Trajectory Analysis Architecture Transportation Elements Affordability and Reliability Analysis Findings Table of Contents 2

Introduction 3

Overview This presentation summarizes the efforts of SpaceWorks Engineering s advanced design team to develop a viable Mars architecture based on Project Constellation launch vehicles and related lunar transportation technologies. Allchemical LOX/LH2 transfer vehicles are used and no in-situ resource utilization (propellant manufacture) has been assumed. An overall concept of operations is outlined and details are provided on element masses, Earth-Mars transfer times, development and operations costs, and estimated mission reliability. Analysis Team Several engineers at SEI participated in this project. Each contributed to one or more of the disciplinary analysis areas below: - Architecture integration & trades - EDL, descent stages, ascent stages, atmospheric trajectories - Surface habitats and transfer habitats - Chemical in-space stages & propulsion - Surface mobility/rovers - In-space trajectories - Cost & campaign analysis - Reliability - Graphic/illustration support Introduction 4

GUIDING DESIGN PHILOSOPHIES - Synergy with Project Constellation flight elements and experiences - Minimum new technology investment REQUIREMENTS (CONSTRAINTS) - Mission Sortie-class surface missions (no permanent base) Three crew members Post-Lunar exploration timeframe (~ 2030 and beyond) - In-Space Propulsion All-chemical in-space propulsion (LOX/LH2) - Launch Utilize standard Ares V and Ares I launch vehicle configurations - Surface Assume no In-situ Resource Utilization (ISRU) Pressurized rover provides surface mobility - Technology Assumptions Zero-boiloff cryogenic storage Aerocapture / Aerobraking of large payloads Use of inflatable Transfer Habitats (TransHabs) for outbound and return crew legs Major Architecture Assumptions 5

DISCIPLINE Atmospheric Trajectories In-Space Trajectories Habitats (Mass, Volumes, and Power) Propulsive Stages (Mass, Volume, and Thrust) Thermal Protection System (TPS) sizing for aerocapture maneuvers (aerobrakes) CAD, Mass Properties, C.G. estimation Liquid Rocket Engine performance, reliability, and weight estimation Cost Estimating (element level DDT&E, TFU) Campaign Analysis (life-cycle cost integration) Mission Reliability estimation Visualization, Graphics SOFTWARE TOOL POST (3D) Bullseye HabSizer StageSizer Sentry Solid Edge REDTOP-2 NAFCOM, cost analogy Cost Analysis Module (CAM) Fault Tree Analysis, Reliability Block Diagrams Canvas, Maya, Photoshop COMMENTS NASA LaRC trajectory analysis code available to U.S. industry. Modified for Mars atmosphere and gravity. SEI commercial software product. Suitable for selection of highthrust interplanetary trajectories. Fast-acting Lambert s solver. SEI internal Microsoft Excel model. Anchored to habitat models and data used for LAT Phase 2. SEI internal Microsoft Excel model. Anchored to Ares I and Ares V upper stage and core mass estimates. SEI commercial software product for TPS stackup optimization. 1- D transient thermal analysis. Insulative and ablative materials database available. EDS commercial CAD software product SEI commercial product for power balance analysis of future rocket engines of various cycles and propellants. ITAR restricted code. NASA MSFC / SAIC element cost estimation software available to U.S. industry. Augmented by comparative analogy for certain mission elements. SEI custom approach to phases and aggregate costs and year-byyear cost distributions for an entire human campaign including technology, development, manufacture, and operations costs. Excel-based instantiation of FTA and RBD approaches. Run probabilistically. Anchored to ESAS reliability results for lunar missions. Commercial software products used for configuration drawings, dimensional layouts, and artist s renderings Note: DENOTES SEI-developed tool SEI s Primary Software Tools for Assessing Human Mars Missions 6

Architecture Overview 7

Overall Architecture Representation 8

Return TransHab Earth Return Vehicle (ERV) Trans Mars Injection (TMI) Stage In-Space Propulsion Stage (ISPS) Surface Habitat Mars Entry Vehicle (MEV) Outbound TransHab Crewed Mars Entry Vehicle (MEV) PRE-DEPLOY CARGO LEG Mission 1.A.1 TMI Stage Mission 1.A.2 ISPS, Return TransHab, ERV Launch Manifest Details In Order of Launch (Nominal Two Year Separation Between Pre-Deployed Cargo Leg and Crewed Leg) CREW/MEV LEG Mission 1.B.1 Surface Habitat MEV Mission 1.B.2 Crewed MEV, Outbound TransHab Mission 1.B.3 Crew to LEO Prior To TMI 9

140 130 120 110 100 90 3.1 t 123.2 t Total = 126.3 t EDS Payload Adapter TMI Stage Assumed Ares V LEO Payload Capability = 130.0 t IMLEO Mass [t] 80 70 60 50 40 1.8 t 6.0 t 10.9 t 53.3 t Total = 72 t EDS Payload Adapter Earth Return Vehicle Return TransHab ISPS 30 20 10 Mission Payload Masses 0 Cargo Launch #1: TMI Stage Cargo Launch #2: ISPS, ERV, Return TransHab 10

50 45 Total = 42.5 t Assumed Ares V C3 Requirement: C3 = 19.4 km 2 /s 2 Corresponds to ~ 43 t Payload 40 1.1 t 0.3 t IMLEO Mass [t] 35 30 25 20 0.8 t 20.3 t Total = 34.5 t EDS Payload Adapter Surface Habitat MEV - Power & Propulsion Module MEV - Descent Stage MEV - Heatshield 10.7 t 4.0 t 14.4 t EDS Payload Adapter Crew Outbound TransHab MEV - Pressurized Rover MEV - Ascent Stage MEV - Descent Stage MEV - Heatshield 15 10 2.9 t 5 8.4 t 9.9 t 0 2.1 t Crew Launch #1: Surface Habitat MEV Mission Payload Masses (Not Including EDS Masses) 2.1 t Crew Launch #2: Crewed MEV, TransHab 11

Interplanetary Trajectory Analysis 12

Heliocentric Transfer Trajectory: Pre-Deploy Cargo Leg PRE-DEPLOY CARGO LEG - Pre-Deploy ISPS, ERV, and Return TransHab - Departure Date = 12/15/2030 - C3 Departing Earth = 12.04 km 2 /s 2 - Earth Departure Delta-V = 3.76 km/s - Time of Flight (TOF) = 285 days - Propulsive Capture Delta-V at Mars = 2.51 km/s Heliocentric Transfer Trajectory: Crew/MEV Leg CREW/MEV LEG - Crew Outbound Leg with MEV and Outbound TransHab - Departure Date = 4/21/2033 - C3 Departing Earth = 9.40 km 2 /s 2 - Earth Departure Delta-V = 3.65 km/s - Time of Flight (TOF) = 205 days - Aerocapture Entry Velocity at Mars = 5.96 km/s (at 125 km) - Surface Stay = 500 days Note: Using SEI- developed Bullseye trajectory code - Crew Return Leg with ERV and Return TransHab - Departure Date = 3/27/2035 - C3 Departing Mars = 10.36 km 2 /s 2 - Mars Departure Delta-V = 2.38 km/s - Time of Flight (TOF) = 230 days - Direct Entry Velocity at Earth = 11.5 km/s (at 125 km) - Crew Return Date = 11/12/2035 - Total Crew Mission Time = 935 days Representative Trajectories for First Opportunity (Minimum Total C3 Solutions) 13

Maximum Earth Departure C3 of 14.88 km 2 /s 2 (Available C3 from TMI Stage 15.32 km 2 /s 2 ) Maximum Mars Arrival C3 of 11.88 km 2 /s 2 (Available C3 from ISPS 12.24 km 2 /s 2 ) 20.0 2030 Cargo Launch Opportunities 12.2 16.0 2033 Cargo Launch Opportunities 12.2 Earth Departure C3 (km 2 /s 2 ) 16.0 12.0 8.0 4.0 6 days 0.0 12/7/2030 12/11/2030 12/15/2030 12/19/2030 Earth Departure Date 12.1 12.0 11.9 11.8 11.7 Mars Arrival C3 (km 2 /s 2 ) Earth Departure C3 (km 2 /s 2 ) 12.0 8.0 4.0 12.0 11.8 11.6 11.4 11.2 11.0 10.8 0.0 33 days 10.6 4/1/2033 4/11/2033 4/21/2033 5/1/2033 5/11/2033 Earth Departure Date Mars Arrival C3 (km 2 /s 2 ) 20.0 2035 Cargo Launch Opportunities 30.0 15.2 2037 Cargo Launch Opportunities 15.0 16.0 24.0 15.1 13.0 Earth Departure C3 (km 2 /s 2 ) 12.0 8.0 4.0 18.0 12.0 6.0 Mars Arrival C3 (km 2 /s 2 ) Earth Departure C3 (km 2 /s 2 ) 15.0 14.9 14.8 11.0 9.0 7.0 5.0 Mars Arrival C3 (km 2 /s 2 ) 0.0 60 days 0.0 5/11/2035 5/31/2035 6/20/2035 7/10/2035 7/30/2035 Earth Departure Date 6 days 14.7 9/2/2037 9/5/2037 9/8/2037 9/11/2037 Earth Departure Date 3.0 Earth Departure C3 Requirement Earth Departure C3 Available Earth Departure Window Mars Arrival C3 Requirement Mars Arrival C3 Available Cargo Mission Departure Windows 14

Maximum Earth Departure C3 of 19.39 km 2 /s 2 (Available C3 from Ares V 19.97 km 2 /s 2 w/ 42.5t Payload) Maximum Mars Departure C3 of 12.7 km 2 /s 2 (Available C3 from ISPS 13.09 km 2 /s 2 ) 2033 Crew Launch Opportunites Mars Departure Date 2035 Crew Launch Opportunites Mars Departure Date 2/12/2035 3/24/2035 5/3/2035 6/12/2035 30.0 109 days 15.0 7/7/2037 7/12/2037 7/17/2037 50.0 7 days 12.8 Earth Departure C3 (km 2 /s 2 ) 25.0 20.0 15.0 10.0 13.0 11.0 9.0 7.0 Mars Departure C3 (km 2 /s 2 ) Earth Departure C3 (km 2 /s 2 ) 40.0 30.0 20.0 10.0 12.7 12.6 12.5 12.4 Mars Departure C3 (km 2 /s 2 ) 5.0 100 days 5.0 12/27/2032 2/15/2033 4/6/2033 5/26/2033 Earth Departure Date 2037 Crew Launch Opportunities 0.0 32 days 6/20/2035 7/5/2035 7/20/2035 8/4/2035 Earth Departure Date 2039 Crew Launch Opportunities 12.3 Earth Departure C3 (km 2 /s 2 ) 7/17/2039 7/24/2039 7/31/2039 8/7/2039 25.0 23.0 21.0 19.0 17.0 Mars Departure Date 12 days 15.0 13.0 11.0 25 days 15.0 5.0 8/3/2037 8/15/2037 8/27/2037 9/8/2037 Earth Departure Date 9.0 7.0 Mars Departure C3 (km 2 /s 2 ) Earth Departure C3 (km 2 /s 2 ) 7/7/2041 7/23/2041 8/8/2041 8/24/2041 30.0 26.0 22.0 18.0 14.0 Mars Departure Date 37 days 15.0 12.0 10.0 57 days 0.0 8/3/2039 9/2/2039 10/2/2039 11/1/2039 Earth Departure Date 9.0 6.0 3.0 Mars Departure C3 (km 2 /s 2 ) Crew Mission Departure Windows Earth Departure C3 Requirement Earth Departure C3 Available Earth Departure Window Mars Departure C3 Requirement Mars Departure C3 Available Mars Departure Window 15

Architecture Transportation Elements 16

TMI Stage - Isp = 464 sec. - Propellants = LOX/LH2 - Mixture Ratio = 5.85 - Inert Mass = 10.53 t - Propellant = 117.67 t Trans-Mars Injection (TMI) Stage 17

ISPS (with Drop Tanks) - Isp = 464 sec. - Propellants = LOX/LH2 - Mixture Ratio = 5.85 - Core Inert Mass = 4.2 t - Core Propellant = 17.2 t - Drop Tanks Inert Mass (total) = 1.6 t - Drop Tanks Propellant (total) = 30.3 t In-Space Propulsion Stage (ISPS) 18

In-Space Inflatable Transfer Habitats (TransHabs) TransHabs Common Mission Requirements - Crew Size = 3 - Closed ECLSS - Inflatable External Shell - Deployable Solar Arrays (x4) - Discarded Prior to Entry (each arrival) Outbound TransHab - Maximum Mission Duration = 355 days - Inert Mass = 8.9 t - Gross Mass (w/ provisions, w/o crew) = 10.7 t - Habitable Volume = 44.4 m 3 - Pressurized Volume = 84.2 m 3 - Average Power = 7,730 W Inbound TransHab - Maximum Mission Duration = 265 days - Has Extra Provisions For Crew Abort From Surface - Inert Mass = 8.8 t - Gross Mass (w/ provisions, w/o crew) = 10.9 t - Habitable Volume = 43.6 m 3 - Pressurized Volume = 84.2 m 3 - Average Power = 7,660 W 19

Surface Habitat MEV - Maximum Diameter = 10 m - Surface Habitat = 20.3 t - Power & Propulsion Module = 2.9 t - Descent Stage = 8.4 t - Forward Heatshield = 2.1 t - Total MEV Mass = 33.7 t Crewed MEV - Maximum Diameter = 10 m - Pressurized Rover = 4.1 t - Ascent Stage = 14.4 t - Descent Stage = 9.9 t - Forward Heatshield = 2.1 t - Total MEV Mass = 30.5 t Mars Excursion Vehicles (MEV) - Integrated Configuration 20

Surface Habitat - Inhabited Mission Duration = 500 days - Crew Size = 3 - Payload to Surface = 500 kg - Closed ECLSS - Rigid Structure - Suitlock Airlock - Carried to Surface on MEV Descent Stage - Inert Mass: 15.9 t (without consumables) - Gross Mass: 20.3 t - Habitable Volume: 84.4 m 3 - Pressurized Volume: 105.8 m 3 - Average Power: 8,070 W Surface Habitat Detail (3 Crew for 500 days) 21

Crewed MEV Entry / Ascent Profile (After Aerocapture into Low Mars Orbit) 22

Affordability and Reliability Analysis 23

Life Cycle Cost (LCC) estimation methodology - Use of historically-based Cost Estimating Relationship (CER) equations where available - Additional estimates derived from previous SEI cost analyses and NASA budget line items - Attempt to include all relevant Mars exploration costs (including test flights, operations costs, sustaining engineering costs, etc.) - Use of SEI s Stack em cost integration tool - Assumed Significant leveraging of launch vehicle development from lunar campaign and minimal leveraging of lunar developments for other elements (MEV, ISPS, and habitats) Element cost estimate assumptions - Earth entry vehicle based upon steady state Orion cost estimate plus 20% - Ground facilities costs estimated using SEI s Facilities and Ground Support Equipment Operations Analysis (FGOA) Model which draws from a database of KSC facilities costs - Mission operations cost based on analogy to International Space Station steady state mission operations - Assume 92-97% learning curve effects for production costs of all in-space elements Reductions by comparison to cost of completely new development realized by leveraging existing systems - Use of existing Ares I and Ares V launch vehicles has advantage of zero development cost - Assume minor engineering modifications with associated cost for Earth Return Vehicle (ERV) - 20% cost savings (DDT&E and TFU) for surface habitats Cost wrap assumptions - 15% margin on all cost estimates - NASA program-level cost wrap of 12% Supporting project cost assumptions - Pre-phase A costs equivalent to 5% of DDT&E costs - $1B (FY2007) total for technology development (FY2025-FY2028) - $1B (FY2007) total for robotic missions (FY2025-FY2028) - Sustaining engineering nominally 5% of DD cost for most elements Cost Analysis Approach 24

Pre-Campaign DEV (2025-2030) OPERATIONS (2031+) FY25 FY26 FY27 FY28 FY29 FY30 FY31 FY32 FY33 FY34 FY35 FY36 FY37 FY38 FY39 FY40 FY41 FY42 FY43 FY44 FY45 FY46 FY247 FY48 Technology Maturation Development and Design Robotic Missions System/Flight Tests Full Up Cargo Leg Test LEO Demo # 1 Full Up Crew Leg Test LEO Demo # 2 (uncrewed) DEMO 1 DEMO 2 Campaign Cargo Leg Ares V Cargo Ares V - TMI Crew/MEV Leg Ares V MEV: Habitat Ares V MEV: Crew Ares I + Orion: Crew 1 2 3 4 Mission Operations In-Space / Surface Ops IOC LEGEND Milestones Ares V-Cargo 1 Ares V-Cargo 2 Ares V-Crew 1 Ares V-Crew 2 Ares I Crew/CEV Flight 1 Sortie # Notional Development Schedule and Operations Campaign 25

TOTAL LIFE CYCLE COST BREAKOUT Cost Item LCC [FY2007] % of Total LCC Trans-Mars Injection Stage In-Space Propulsion Stage Crewed MEV Ascent Stage $4,630 M $3,930 M $7,580 M 4.8% 4.1% 7.8% TOTAL LCC BREAKOUT Crewed MEV Descent Stage $4,450 M 4.6% Trans Mars Injection Stage, 4.8% Pressurized Rover Surface Habitat MEV Propulsion Module Surface Habitat MEV Descent Stage Surface Habitat TransHabs Earth Return Vehicle Operations (Mission, Ground, EVA) Facilities: Launch $2,970 M $510 M $4,220 M $9,840 M $6,930 M $510 M $9,430 M $450 M 3.1% 0.5% 4.4% 10.2% 7.2% 0.5% 9.7% 0.5% Program Integration (Govt.), 9.3% Robotic Missions, 1.0% Technology Development, 1.0% Reserves, 13.0% In-Space Transfer Stage, 4.1% Crewed MEV Ascent Stage, 7.8% Crewed MEV Descent Stage, 4.6% Pressurized Rover, 3.1% Surface Habitat MEV Propulsion Module, 0.5% Surface Habitat MEV Descent Stage, 4.4% Facilities: Mission Operations $2,450 M 2.5% Launch Vehicles, 14.6% Surface Systems Launch Vehicles $1,130 M $14,130 M 1.2% 14.6% Surface Systems, 1.2% Surface Habitat, 10.2% Technology Development* Robotic Missions* Program Integration (Govt.) $1,000 M $1,000 M $9,020 M 1.0% 1.0% 9.3% Facilities: Mission Operations, 2.5% Facilities: Launch, 0.5% Operations (Mission, Ground, EVA), 9.7% TransHab, 7.2% Earth Return Vehicle, 0.5% Reserves $12,630 M 13.0% Total $96,810 M 100.0% *Denotes Fixed Cost Campaign Life Cycle Cost (LCC) 2025-2040 26

Total Life Cycle Cost (LCC) for FY2025 to FY2040 = $97 B $18,000 Reserves $16,000 Program Integration (Govt.) Robotic Missions $14,000 Technology Development Launch Vehicles US$M [FY2007] $12,000 $10,000 $8,000 $6,000 NASA FY2007 Space Operations and Exploration Budgets = US$10,261 M Surface Systems Facilities: Mission Operations Facilities: Launch Operations (Mission, Ground, EVA) Earth Return Vehicle TransHab Surface Habitat Surface Habitat MEV Descent Stage $4,000 Surface Habitat MEV Propulsion Module Pressurized Rover Crewed MEV Descent Stage $2,000 Crewed MEV Ascent Stage In-Space Transfer Stage $- 2025 2026 2027 2028 2029 2030 2031 2032 2033 2034 2035 2036 2037 2038 2039 2040 Trans Mars Injection Stage Fiscal Year Annual Program Cost: Sand Chart 27

Reliability analysis performed using Fault Tree and Event Sequence Diagram (ESD) approach with basic events at the subsystem level Continuous operating time reliability (1-e -λt ) applied for subsystems with significant operating times Subsystem redundancy and engine-out capability modeled Sources include NASA Exploration Systems Architecture Study (ESAS), published data of commercial companies (e.g. P&W RL-10), SEI internal models (e.g. GT-Safety), AIAA and other technical papers Loss of Crew (LOC) reliability calculated by applying likelihood of failure resolution to each major loss of mission event in ESDs Probability distributions generated from 20,000 Monte Carlo trials of input variables with triangular distributions (+/- 20%) Mars Entry Vehicle (MEV) Entry Fault Tree Reliability Analysis Approach 28

400 350 Mean: 11.5% 350 300 Mean: 38.6% Frequency 300 250 200 150 100 50 90th: 12.4% Frequency 250 200 150 100 50 90th: 39.6% 0 0 10.2% 10.5% 10.8% 11.1% 11.5% 11.8% 12.1% 12.4% 12.8% 13.1% 13.4% 36.1% 36.6% 37.0% 37.5% 37.9% 38.4% 38.8% 39.3% 39.7% 40.2% 40.6% 41.1% LOC Probability (% ) LOM Probability (%) LOC LOM LOM 0% 5% 10% 15% 20% 25% 30% 35% 40% LOM or LOC (%) Ares V w/ EDS (4) Earth Orbit Rendezvous Ares I TransHab (2) Aerocapture & MOR MEV (2) Surface Hab Ascent ISPS and TMI Stage ERV ============== Cargo leg Crew leg Return leg LOC once every 8.5 missions (11.5 % mean LOC) LOM once every 2.6 missions (38.6 % mean LOM) Mars entry and long duration operation of the TransHabs and surface habitats are the chief contributors to LOC Ares V launches, TransHab operation, and MEV entry account for most of the LOM Reliability Analysis Results 29

Findings 30

While NASA s current focus is on the Ares I and Orion programs and lunar exploration initiatives, it is important even at this early stage of development to see how resources produced by these programs translate to future missions. This presentation has outlined a feasible Mars exploration architecture that is predicated on using multiple elements and subsystems that are expected to be developed and matured during NASA s lunar exploration campaign. The Mars architecture outlined in this study uses traditional LOX/LH2 propulsion systems along with Ares I and Ares V launch vehicles. The architecture was sized to provide four sortie-style Mars exploration missions between 2030 and 2040. Observations 31

In general, this architecture has very few large-scale requirements for new technologies beyond those being developed for the lunar campaign. However, key technology needs still exist in the following areas: Aerocapture - Aerocapture techniques at Mars with large payloads and entry vehicle masses - Ablative TPS materials for Mars aerobrake and Earth Return Vehicle. Baseline is PICA, but Boeing's Phenolic Ablator (BPA), SLA-561, SIRCA, and even Avcoat are feasible alternatives. Habitats - Inflatable TransHab technology for long duration space flight (note that we may be able to leverage NASA and Bigelow s earlier investment in inflatables) - Closed water and air ECLSS for TransHabs and surface habitat (unless already developed for the lunar campaign) Cryogenic Storage - Zero/low boiloff technology for LOX and LH2 tanks on ISPS stage (~4 year mission time) - Zero/low boiloff technology for the smaller MEV ascent and descent stage tanks (~2.5 year mission time) Key Technology Development Needs 32

Business Address: 1200 Ashwood Parkway Suite 506 Atlanta, GA 30338 U.S.A. Phone: 770-379-8000 Fax: 770-379-8001 Internet: WWW: E-mail: info@sei.aero 33