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Analysis of the Utility of Various Human Landing System Architectures FISO Presentation January 29, 2020 Dr. Tim Kokan, Principal Engineer, Aerojet Rocketdyne Phone: 256-922-2579; E-mail: [email protected] Contents • HLS Architecture Trade Study Overview • Trade Matrix • Ground Rules and Assumptions • Utility Analysis • Analysis Results and Observations • Downselection to Six Architectures • Alternate Weightings Analysis • Uncertainty Analysis • Summary Observations • Backup 2 HLS Architecture Trade Study Overview • Today’s FISO briefing provides an overview of recent Human Landing System (HLS) architecture trade studies Human Landing performed by Aerojet Rocketdyne System • This architecture trade study included an examination of a Ascent range of HLS configurations, launch vehicle options, Element CONOPS options, main propulsion options, and other subsystem design options • Cost, schedule, reliability, extensibility, and performance Descent attributes were assessed for each HLS architecture option Element analyzed • Each HLS architecture option was scored based on these Transfer attribute results utilizing the utility analysis methodology Vehicle Element • High scoring options were further studied with alternate attribute weightings and Monte Carlo uncertainty analyses • Summary observations are provided along the way pointing out various interesting results 3 HLS Architecture Trade Matrix Design Driver Options Transfer Vehicle Element Yes No Refueling Element Yes No TVE Reusability Expendable Reusable Ascent Element Launched Yes No with Descent Element Ascent Element Launch NASA SLS Blue Vulcan NGIS OmegA NASA SLS B1 Delta IVH Ariane 6 SX FH Vehicle Options B1B Origin NG Heavy Heavy Descent Element Launch NASA SLS Blue Vulcan NGIS OmegA NASA SLS B1 Delta IVH Ariane 6 SX FH Vehicle Options B1B Origin NG Heavy Heavy Transfer Vehicle Element NASA SLS Blue Vulcan NGIS OmegA NASA SLS B1 Delta IVH Ariane 6 SX FH Launch Vehicle Options B1B Origin NG Heavy Heavy Lunar Surface Access Polar Global LV Dropoff to Aggregation Weak Stability Ballistic Lunar Orbit Transfer Options - Boundary Ascent LV Dropoff to Aggregation Weak Stability Ballistic Lunar Orbit Transfer Options - Boundary Descent/TVE Low Lunar Low Earth High Earth NRHO High Lunar Orbit Aggregation Orbit Orbit Orbit Orbit NRHO to LLO Transfer Time 0.5 days 1 day 2 days LLO to NRHO Transfer Time 0.5 days 1 day 2 days for Ascent LLO to NRHO Transfer Time Optimal Same for TVE Element Diameter 4.4m 7.0m Ascent Element Non- 2 mT 3 mT 4 mT Propulsion System Dry Mass Ascent Adv Ascent Element Main Engine Evolved Orion Main LMAE (RS-18) OMS AJ10-118K XLR-132 OMS-E RS-72 Main Methane OMAC Engine Options Engine Engine Ascent Adv Descent/TVE Main Engine Evolved LMAE Orion Main AJ10- RL10A-4-2 RL10C-5-1 RL10C-3 RL10C-2-1 OMS XLR-132 OMS-E RS-72 Main Methane OMAC (RS-18) Engine 118K Options Engine Engine Thermal - Ascent Passive Only Passive + Active Thermal - Descent/TVE Passive Only Passive + Active ~20 Billion Potential Options; 326,656 Highlighted Options Run In This Study 4 Ground Rules and Assumptions 5 Launch Vehicles 400 ft 350 ft 325 ft 313 ft ~313 ft 300 ft ~280 ft 250 ft 229 ft 214 ft 200 ft 150 ft 100 ft 50 ft Company/Organization ULA SpaceX NGIS Blue Origin NASA NASA SLS Block 1 SLS Block1B Launch Vehicle Vulcan Heavy Falcon Heavy OmegA Heavy New Glenn Cargo Cargo Configuration Expendable Expendable Expendable Reusable Expendable Expendable Payload to TLI 12 mT(b) 15 mT(a) 12 mT(c) 10 mT(b) 27 mT(a) 40 mT(a) Dynamic Payload Volume 317 m3 145 m3 233 m3 458 m3 230 m3 621 m3 Assumptions: (a) Public; (b) AR internal payload estimate anchored to publically-available GTO payload; (c) AR internal payload estimate anchored to publically-available GEO payload 6 Representative 3-Element HLS CONOPS ~1 Week Surface Stay Descent and Ascent Elements go to Lunar Surface HLS Transfers to LLO Crew Returns (0.5 Days) to NRHO LLO TVE Returns to NRHO Lunar Elements Aggregate (0.5 Days) Gateway at Gateway NRHO Crew Transfers to Lander TVE and Ascent Element Orion Remains in Lunar await refueling for next Orbit w/ Gateway mission Crew Returns to Earth Ascent Descent Transfer Orion/Crew Element Element Vehicle Element AR Architecture 3-Element Study Assumes Launch Order of Ascent/Descent/TVE LEO Comm Comm Comm SLS B1 LV LV LV Crew 7 Representative 2-Element HLS CONOPS Launched Separately ~1 Week Surface Stay Descent and Ascent Elements go to Lunar Crew Returns Surface to NRHO LLO Lunar Elements Aggregate Gateway at Gateway NRHO Crew Transfers to Lander Orion Remains in Ascent Element Lunar Orbit w/ awaits refueling for Gateway next mission Crew Returns to Earth Ascent Descent Orion/Crew Element Element AR Architecture 2-Element Study Assumes a Launch Order of Ascent/Descent if LEO Launched Separately; up to 90 day Comm SLS B1/B1B SLS B1 separation between SLS launches LV Cargo Crew 8 Representative 2-Element HLS CONOPS Launched Together ~1 Week Surface Stay Descent and Ascent Elements go to Lunar Crew Returns Surface to NRHO LLO Lunar Elements Aggregate Gateway at Gateway NRHO Crew Transfers to Lander Orion Remains in Ascent Element Lunar Orbit w/ awaits refueling for Gateway next mission Crew Returns to Earth Ascent + Orion/Crew Descent Up to 90 day separation between SLS launches LEO SLS B1B SLS B1 Cargo Crew 9 Utility Analysis 10 Utility Analysis Background • Utility analysis provides a methodology to make design decisions based on customer preferences – Define Key Decision Attributes (cost, schedule, reliability, performance, etc) – Define Utility Functions and Weightings (functions to translate attribute values to a non-dimensional, normalized utility value to the customer) – Perform Architecture Assessments – Calculate Utility Score for Each Architecture Option – Downselect to Highest Scoring Architecture Options for Further Assessment • A key benefit of utility analysis is the non-linear nature of the utility functions. This feature enables the accurate evaluation of a much larger trade space than with traditional linear methods that are only valid in the local neighborhood of a particular point design. 11 Key Decision Attributes ID KDA Name KDA Type Definition Units Low High Architecture Non-Recurring Non-Recurring Cost of A-1 Cost Delta off Reference Cost Ascent/Descent/TVE relative to baseline FY19$M -160 530 Architecture architecture Recurring Cost of Ascent/Descent/TVE Architecture Recurring (Production Cost + Operations Cost Per A-2 Cost Delta off Reference Cost FY19$M -200 375 Mission + Launch Cost) relative to Architecture baseline architecture HLS Development, Number of Months from ATP to Landing A-3 Production, and Flight Schedule Months 36 84 on Lunar Surface Schedule A-4 HLS Reliability Reliability System Hardware Reliability Per Mission % 95% 100% Level of Extensibility to Achieving A-5 HLS Extensibility Extensibility Non-Dimensional 3 15 Sustained Lunar Architecture Percent Launch Vehicle Payload Mass A-6 LV Payload Mass Margin Performance Margin Relative to Ascent/Descent/TVE % -50% 50% Launch Mass Percent Launch Vehicle Payload Length LV Payload Envelope A-7 Performance Margin Relative to Descent/TVE Stowed % -25% 50% Margin Length 12 How Are We Measuring HLS Architecture Extensibility? • Treated as a non-dimensional parameter for initial trade study • Provides credit for element / mission design choices that align with long-term HLS Architectures – ISRU, Global Surface Access, 4 crew capable HLS designs Ascent KDA Propellant Combination Lunar Surface Element Crew Score Access Ascent – Descent (+TVE) Cabin Mass 1 Storable – Storable Polar 2 mT 2 LOX/CH4 – Storable 3 Storable – LOX/CH4 3 mT LOX/CH4 – LOX/CH4 4 Storable – LOX/LH2 5 LOX/CH4 – LOX/LH2 Global 4 mT Cryogens, Global Lunar Surface Access, and Larger Ascent Elements Receive Higher Utility Scores 13 Utility Functions 1.0 1.0 1.0 0.8 0.8 0.8 0.6 0.6 0.6 u(x) u(x) 0.4 0.4 u(x) 0.4 0.2 0.2 0.2 0.0 0.0 0.0 -160 -22 116 254 392 530 -200 -85 30 145 260 375 36 42 48 54 60 66 72 78 84 HLS Architecture Non- HLS Architecture Recurring HLS Architecture Development Recurring Cost Delta (FY19$M) Cost Delta (FY19$M) Schedule (Montes from ATP) 1.0 1.0 1.0 1.0 0.8 0.8 0.8 0.8 0.6 0.6 0.6 0.6 u(x) u(x) u(x) 0.4 0.4 0.4 u(x) 0.4 0.2 0.2 0.2 0.2 0.0 0.0 0.0 -50% -25% 0% 25% 50% 95% 96% 97% 98% 99% 100% 0.0 -25% 0% 25% 50% HLS System Hardware 3 6 9 12 15 Launch Vehicle Payload Mass Launch Vehicle Envelope Reliability HLS Architecture Extensibility Margin Margin Utility Functions Are Defined for Each Key Decision Attribute Based on Customer Preferences 14 Key Decision Attribute Baseline Weightings Reliability and Development Schedule Weighted Highest; Recurring Cost and Extensibility Weighted Lowest 15 Analysis Results and Observations 16 HLS Element Non-Recurring Price Breakdown Price Breakdown Transfer Vehicle shown for Element Reference 28% Architecture Descent Element 56% Ascent Element 16% Commonality between Descent and TVE reduces development costs of TVE. Ascent Element cost includes only propulsion system. 17 HLS Architecture Recurring Price Breakdown TVE Production 7% Descent Price Breakdown Production 9% shown for Reference Ascent Production 3% Architecture TVE Operations 5% Descent Operations 3% Ascent Operations 6% Total Launch 67% Total Mission Recurring Price Dominated by Launch Vehicle. Ascent Element cost includes only propulsion system. 18 HLS Architecture Hardware Failure Breakdown by Element Failure Probability Breakdown shown for Reference Transfer Vehicle Architecture Element 0.843% Descent Element, 0.824% Ascent Element, 0.077% Failure Breakdown Dominated by Descent Element and TVE; Ascent is Only Human Rated Element.