Outline Aerocapture Technology Developments by the In-Space Introduction to Aerocapture Propulsion Program Application at Titan, Venus, and Neptune Current Development Status Next Steps

Michelle M. Munk In-Space Propulsion Aerocapture Manager

Planetary Science Subcommittee Meeting | October 3, 2008 1 2 www.nasa.gov

Aerobraking vs Aerocapture Aerocapture Benefits for Robotic Missions

Entry targeting burn Atmospheric entry Nominal Best A/C Best Atmospheric Drag Energy Orbit A/C % Best non- Aerocapture Mission - Science Orbit Mass, non-A/C Reduces Orbit dissipation/ Insertion V, Increase A/C Option ! kg Mass, kg Period Autonomous km/s guidance Periapsis Venus V1 - 300 km circ 4.6 5078 2834 79 All-SEP raise Aerobraking maneuver Venus V2 - 8500 x 300 km 3.3 5078 3542 43 All-SEP (propulsive) Target orbit Mars M1 - 300 km circ 2.4 5232 4556 15 Aerobraking ~300 Passes Controlled exit Mars M2 - ~1 Sol ellipse 1.2 5232 4983 5 Chem370 Hyperbolic Through Upper Jettison Aeroshell Approach Atmosphere Jupiter J1 - 2000 km circ 17.0 2262 <0 Infinite N/A Aerocapture: A vehicle uses active control to autonomously guide itself to an atmospheric exit target, establishing a final, low Jupiter J2 - Callisto ellipse 1.4 2262 4628 -51 Chem370 orbit about a body in a single atmospheric pass. Orbit Insertion Saturn S1 - 120,000 km circ 8.0 494 <0 Infinite N/A Burn Pros Cons Pros Cons Uses very little fuel--significant Needs protective aeroshell Titan T1 - 1700 km circ 4.4 2630 691 280 Chem370 mass savings for larger vehicles Little spacecraft design Still need ~1/2 propulsive fuel Uranus U1 - Titania ellipse 4.5 1966 618 218 Chem370 impact load Establishes orbit quickly (single One-shot maneuver; no turning pass) back, much like a lander Gradual adjustments; can Hundreds of passes = more Neptune N1 - Triton ellipse 6.0 1680 180 832 Chem370 pause and resume as chance of failure Has high heritage in prior Fully dependent on flight needed (with fuel) hypersonic entry vehicles software Aerocapture offers significant increase in delivered payload: Operators make decisions Months to start science Flies in mid-atmosphere where ENHANCING missions to Venus, Mars Operational distance limited dispersions are lower by light time (lag) Adaptive guidance adjusts to STRONGLY ENHANCING to ENABLING missions to Titan, and Uranus At mercy of highly variable day-of-entry conditions ENABLING missions to Jupiter, Saturn, and Neptune upper atmosphere Fully autonomous so not distance-limited 3 Ref.: Hall, J. L., Noca, M. A., and Bailey, R. W. “Cost-Benefit Analysis of the Aerocapture Mission Set,” Journal of Spacecraft and Rockets, 4 Vol. 42, No. 2, March-April 2005 Titan Aerocapture Reference Concept Titan Systems Definition Study Results

Mass (kg) Subsystem Rack-up

Current Best % System Component Estimate Contingency Growth Allocation Lander 280.2 29.8% 363.8 400.0 Orbiter/Lander Interface 47.5 30.0% 61.8 61.8 Orbiter 883.6 24.2% 1097.7 1200.0 Prop Mod/Orbiter Interface 47.3 30.0% 61.4 61.4 SEP Prop Module 1084.0 21.4% 1316.5 1450.0 Launch/Prop Mod Interface 60.0 30.0% 78.0 78.0 Stack Total 2402.6 24.0% 2979.2 3251.2 Launch Vehicle Capability 3423 Lander System Level Mass Margin 29.8% ( LV Cap - CBE ) / LV Cap System Reserve 13.0% ( LV Cap - Growth ) / LV Cap “Titan Explorer”-type mission based on Launch Veh: 4450 4450 Delta IV H Delta IV H SSE Roadmap circa 2001 2.4 m diameter HGA Gravity Assist: EGA VGA EGA VEEGA Orbiter Detailed analysis by multi-Center team of Upper Stage: SEP SEP SEP Chem Capture Type: Chem discipline experts (many papers) Aero Aero Aero Trip Time: 6 yrs 6 yrs 6 yrs 12 yrs* Delta 4450, SEP, EGA, aerocapture has SEP Prop 30% system level margin, >10% system * Includes 2-yr moon tour used to reduce propellant requirements for all propulsive capture reserve Module Aerocapture mass fraction = 39% of orbiter • Aerocapture/SEP is Enabling to Strongly Enhancing, dependent on Titan mission requirements launch wet mass Solar • Aerocapture/SEP results in ~2.4x more payload at Titan compared to all-propulsive mission for 3.75 m diameter Arrays Aeroshell same launch vehicle 5 6 Aerocapture can be used with a chemical ballistic trajectory: Delta IV H, 7.1 year trip, EGA, 32% margin Ref: “Aerocapture Systems Analysis for a Titan Mission”, NASA TM 2006-214273, March 2006.

Titan Aerothermal Updates Since 2002 Study Titan-GRAM Model vs Cassini-Huygens Data

Cassini-Huygens provided: • Improved ephemeris data for reduced flight path angle uncertainty • Improved atmospheric density measurement accuracy • Improved atmospheric constituent data (less than 2% CH4 vs 5% assumed in 2002 study) Aerothermal modeling investments and testing provided improved aeroheating estimates and less critical need for TPS Aerocapture Minimum Alt Range Aerocapture to Orbit Alt Range development • Reduced heating estimates Observations from HASI and INMS are well within Titan-GRAM result in 75-100 kg less TPS Ref: Mike max/min estimates mass than sized during the Wright 2002 study (Laub and Chen, 2005) 7 Ref.: Justh and Justus, “Comparisons of Huygens Entry Data and Titan Remote Sensing Observations with the 8 Titan Global Reference Atmospheric Model (Titan-GRAM)” Titan Aerocapture Technologies - Ready! Aerocapture Benefit for a Venus Mission

Enabling Technologies - No new enabling technology required 1165 kg Launch Vehicle Capability Strongly Enhancing Technologies Delta 2925H-10, C3 = 8.3 km2/s2 Mass savings will ! Aeroheating methods development, validation scale up for • Large uncertainties currently exist, improved prediction capability could result in reduced TPS mass Flagship-class ! TPS Material Testing • TPS materials proposed and other TPS options exist today, but are not tested against expected mission radiative heating at Titan !Atmosphere Modeling Venus Orbiter Enhancing Technologies (OML Design Only) ! Aeroshell lightweight structures - reduced aerocapture mass Guidance - Existing guidance algorithms have been demonstrated to provide acceptable performance, improvements could provide increased robustness ! Simulation - Huygens trajectory reconstruction, statistics and modeling upgrades Mass properties/structures tool - systems analysis capability improvement, concept trades Deployable high gain antennae – increased data return 300 x 300 km Ø 2.65 m The following technologies provide significant benefit to the mission but are already in a funded development cycle for TRL 6 • MMRTG (JPL sponsored AO in proposal phase, First flight MSL) Into 300 x 300 km Venus orbit with same launch vehicle, Aerocapture delivers: • SEP engine (Glenn Research Center engine development complete in ‘10) • 1.8x more mass into orbit than aerobraking • Second Generation AEC-Able UltraFlex Solar Arrays (175 W/kg) • 6.2x more mass into orbit than all chemical • Optical navigation to be demonstrated on MRO 9 10 ! Reference: “Systems Analysis for a Venus Aerocapture Mission”, NASA TM 2006-214291, April 2006

Example Monte Carlo Simulation Results: Venus Aerocapture Venus Aerocapture Technology - In Good Shape Venus Aerocapture Systems Analysis Study, 2004 2 Vehicle L/D = 0.25, m/CDA = 114 kg/m Target orbit: 300 km circ., polar • Aerocapture is feasible and robust at Venus with high heritage low All-propulsive !V required for orbit insertion: 3975 m/s L/D configuration !V provided by aerocapture: 3885 m/s (97.7% of total) • 100% of Monte Carlo cases capture successfully 30 deg/sec bank rate, 5 deg/sec2 bank acceleration • TPS investments could enable more mass-efficient ablative, insulating TPS; accompanying aerothermal analysis investments 100% successful capture would enable prediction of ablation, potential shape change • Additional guidance work would increase robustness for small scale height of Venus atmosphere • Mass savings will scale up for a Flagship-class mission, so Aerocapture provides a way to achieve the challenging science return that is desired • Possible orbiter + lander/probe on 1 launch

90 m/s of post- aerocapture propulsive !V

11 12 1-sigma variations at 100 km = ~8%; 3" = ~24% Neptune Orbiter Aerocapture Reference Concept Neptune Aeroheating Challenges

Mass (kg) • Vehicle divided into 4 zones for TPS sizing. Launch Mass Summary Subsystem Rack-up System TPS selected/sized for max heating point in 35% Dry Margin Carried at Current Allocation Zone 4 Orbiter and SEP Level each zone. Flight Best % Wet minus Fuel Dry Mass (includes Component Units Estimate Growth Growth Allocation Fuel Load Load Margin Zone 3 Orbiter Launch Dry Mass 269 518.2 28.5% 666.0 1081.4 282.5 798.9 35.1% base) • TPS interfaces will require a significant effort Aeroshell/TPS Dry Mass 34 681.0 30.0% 885.2 885.3 0.0 885.3 • No facility exists for testing these heating Probes (2) 2 159.3 30.0% 207.1 228.6 0.0 228.6 30.3% Zone 1 levels SEP Stage Dry Mass 197 1133.8 29.7% 1469.7 2899.2 1154.5 1744.7 35.0% Zone 2 Launch/Prop Mod Interface 1 49.0 30.0% 70.0 70.0 0.0 70.0 • Combination radiative and convective Stack Total 503 2541.3 29.8% 3298.0 5164.5 1437.0 3727.5 31.8% ~0.74 m 2.88 m environment is analysis and testing challenge Launch Vehicle Capability 5964

Unallocated Launch Reserve 13.4% Unallocated Reserve / LV Cap JPL System Dry Mass Margin 31.8% ( Dry Alloc - Dry CBE ) / Dry Alloc 5m Fairing ( Dry Growth - Dry CBE ) / Dry CBE (Measure of component maturity) NASA Dry Mass Contingency 29.8% Zone 2 – Wind High

1.6 ) NASA Dry Mass Margin 13.0% ( Dry Alloc - Dry Growth ) / Dry Growth (Measure of system maturity) 16 2 Orbiter Zone 1 – Nose ) TPS Sizing High 2 4.0 40 1.4 14 ) Delta IV H, 5m Fairing, 5964 kg, C3 = 18.44 Radiative 2 TPS Sizing Solar 3.5 35 31.8% System Dry Mass Margin; 13% 1.2 Convective 12 Radiative Unallocated Launch Reserve (800 kg) arrays 3.0 Convective 30 ) Mass margin provides opportunity for 1.0 10 2 • Third probe 2.5

Med 25 • Increased aeroshell size for possible reduction .8 Ref 8 Med in aeroheating rates/loads, TPS thickness 2.0 20 (kW/cm requirements, surface recession SEP Prop .6 6 Ref

1.5 Low Module 15 Outside of Outside

Orbiter of Outside

~57% aerocapture mass fraction Low Total Peak Heat Rate 2.88 m Length Flattened Total Heat Load (kJ/cm .4 4

(includes aerocapture propellant) Total Heat Load (kJ/cm 1.0 10

Ellipsled Aeroshell Range Expected Total Peak Heat Rate (kW/cm Expected Range Expected ~48% structure/TPS mass fraction .2 2 .5 5 2 Probes 13 0 0 0 0 14 Reference: “Systems Analysis for a Neptune Aerocapture Mission”, NASA TM 2006-214300, May 2006

Neptune Aerocapture Technologies - Need Work ISPT’s Low-Risk Aeroshell Mass Improvements

Enabling Technologies Warm Structure System Model - based on MER, MPF, validated with testing TPS Manufacturing For environments up to 300 W/cm2 • TPS thicknesses are beyond current manufacturing experience for carbon phenolic for this shape/acreage HT-424 Aerothermodynamic methods and validation for adhesive • Aerothermodynamics characterized by high radiative and convective aeroheating, coupled convection/radiation/ablation, significant surface recession Aluminum Modified RS9 14% honeycomb for adhesive • Coupled convection/radiation/ablation capability for three-dimensional flowfields Improvement • Approach needed to determine and represent aerodynamics/uncertainties on resultant time varying Graphite polycyanate path dependent shapes in aero database/simulation honeycomb • Testing facilities and methods MER SLA-561V System 250 deg C Warm Structure SLA-561V System 316 deg C Areal Density = 2.07 lb/ft2 Areal Density = 1.78 lb/ft2 Strongly Enhancing Technologies Guidance Algorithm - Existing guidance algorithms provide adequate performance; Improvements Hot Structure System Model - based on Genesis, validated with testing possible to determine ability to reduce heat loads for given heat rate; accommodate time varying, path 2 dependent shape and ballistic coefficient change For environments up to 700 W/cm Flight Control Algorithm - Accommodate shape change uncertainties Atmosphere Modeling - Neptune General Circulation Model output to represent dynamic variability of Final System includes 31% atmosphere Final System included 11-layer MLI Reduced Mass TPS - Lower mass TPS concepts, ex. Reduced density carbon phenolic backup aluminum and high-temp coating Improvement Honeycomb structure Alpha Modulation Lower Mass and Power Science Instruments Dual Stage MMRTGs Hot Structure Carbon-Carbon/Calcarb Deployable Ka-Band HGA Genesis Carbon-Carbon 15 Areal Density = 3.65 lb/ft2 Areal Density = 2.50 lb/ft2 16 Higher Bondlines and Efficient Ablators Reduce Mass Aerocapture Flight Validation Concept ARA Material Density Heating Range New Missions Features Mission Parameters Vehicle Type 60° sphere-cone aeroshell Hyperlite-A 0.21 g/ cm2 55 – 115 W/cm2 MPF-type High Efficiency Vehicle Mass (CBE) 148 kg, 1.2 m diameter Access to space Delta-II dual launch to 13000 km 2 2 SRAM-17 0.27 g/cm 115 – 210 W/cm CEV, MSL Robust Char Mission Duration 9.1 hours SRAM-20 0.32 g/cm2 140 – 260 W/cm2 CEV, MSL, RTF repair Low Recession Atmospheric Entry Speed 9.6 km/s Atmospheric !V 1.7 km/s PhenCarb-20 0.32 g/cm2 200 – 500 W/cm2 CEV, Titan High Heating Nominal Launch June 2010 ST9 Vehicle Concept NMP ST9 Funding $85 M 2 2 PhenCarb-32 0.51 g/cm 500 – 1,100 W/cm Venus, Neptune Severe Heating ISP ST9 Funding $22 M • Aerocapture System Technology for Planetary Missions Mission Sequence was one of five competitors for NASA’s New Millennium SRAM and PhenCarb Sizing SRAM-20 (forebody, Titan Reference Mission) Program Space Technology-9 mission (2006) 1 SRAM-17 2.TCM to set up for entry SRAM-14 • The ST9 Aerocapture concept would have validated: PC-20PC-20 0.8 To high apoapse – Aerocapture as a system technology for immediate use in (13000 km) 130% decrease future missions to Solar System destinations possessing 4.Periapse raise 0.6 significant atmospheres TPS Thickness – The performance of the autonomous Aerocapture guidance (in) 0.4 0.8 system based on bank angle control

1-m SRAM-20 aeroshell test at Solar Tower – Efficient and robust new TPS for multiple applications 0.2 • Feedback on technology element readiness was very 0.6 favorable 1. Launch 0 250 325 400 • ISPT’s recent maturation plans largely guided by work 5.Parking orbit(s) (current SOA) 17 3. Aerocapture18 Allowable Bondline Temperature (deg C) defined in this proposal for data download 0.4

0.2

Current (and Final) ISPT Aerocapture Tasks Current (and Final) ISPT Aerocapture Tasks (cont’d) (through0 FY09) 250 325 400 • Manufacture “large scale” (2.65-m) aeroshell • Verify guidance software operation in “hardware-in-the-loop” ground • Advanced, high-temperature structure by ATK testbed SRAM-20 ablator applied using “modular” approach • • Verify timing and control interfaces • Sensor/repair plugs included • Perform Space Environmental Effects testing on promising materials for both rigid aeroshells and inflatable decelerators (TPS, structure, adhesive, sensors) • Impact • Space Radiation • Cold Soak

• Followed by arcjet testing • Continue aerothermal modeling efforts • Spectrometer measurements of ablation products • Surface catalysis analysis

• CO2 EAST tests to verify shock chemistry

19 20 Aerocapture Technology Subsystem Readiness Aerocapture Development Summary

Destination Venus Earth Mars Titan Neptune • Aerocapture is Enabling or Strongly Enhancing Subsystem for many of the destinations in the Solar System, Venus-GRAM (2004) Earth-GRAM (1974) Mars-GRAM (1988) Titan-GRAM (2002) based on Neptune-GRAM (2003) Atmosphere based on world-wide validated by Space Shuttle continuously updated with Yelle atmosp. Accepted developed from Voyager, saving launch mass, trip time, and cost VIRA. latest mission data. worldwide to be updated with other observations Goal: Capture Physics Cassini-Huygens data Heritage shape, well Heritage shape, well Heritage shape, well Heritage shape, well New shape; aerodynamics • Aerocapture is made of flight system elements Aerodynamics understood aerodynamics understood aerodynamics understood aerodynamics understood aerodynamics to be established. that have Strong Heritage and firm CA=±3%, CN=±5%, # CA=±3%, CN=±5%, # CA=±3%, CN=±5%, # CA=±3%, CN=±5%, # CA=±8%, CN=±8%, # Goal: Errors ! 2% =±2% =±2% =±2% =±2% =±10% TRIM TRIM TRIM TRIM TRIM computational basis GN&C APC algorithm captures Small delivery errors. APC Small delivery errors using Ephemeris accuracy APC algorithm with # Goal: Robust 96% of corridor algorithm captures 97% of !DOR. APC algorithm improved by Cassini- control captures 95% of performance for 4-6 corridor captures 99% of corridor Huygens. APC algorithm corridor. • ISPT investments in modeling and test DOF simulations captures 98% of corridor capabilities are Benefiting Current NASA More testing needed on Technology ready for ST9. ISPT investments have ISPT investments have Zoned approach for mass TPS efficient mid-density TPS. LMA hot structure ready for provided more materials provided more materials efficiency. Needs more projects Goal: Reduce SOA by Combined convective arrivals > 10.5 km/s. ready for application to ready for application. investment. 30%+, expand TPS and radiative facility slow arrivals, and new choices needed. ones for faster entries. • ISPT investments have readied Multiple Structures High-temp systems will High-temp systems will High-temp systems will High-temp systems will Complex shape, large Heatshield Components for Mission Infusion reduce mass by 31%. reduce mass by 14%-30%. reduce mass by 14%- reduce mass by 14%-30%. scale. Extraction difficult. Goal: Reduce SOA 30%. mass by 25% • 2 warm structure systems Convective models Environment fairly well- Convective models agree Convective models agree Conditions cannot be within 15%. Radiative: Aerothermal match within 20% known from Apollo, Shuttle. within 15%. Radiative duplicated on Earth in • Hot structure system laminar, 45% with Models match within 15% predict models will agree models agree within 35- existing facilities. More Goal: Models match turbulence. Radiative within 50% where radiation 300% work on models needed. within 15% models agree within 50% is a factor. • Multiple new charring ablators System Accomplishes 97.7% of ! Accomplishes 97.2% of !V Accomplishes 97.8% of ! Accomplishes 95.8% of !V Accomplishes 96.9% of ! Goal: Robust V to achieve 300 x 300 to achieve 300 x 130 km V to achieve 1400 x 165 to achieve 1700 x 1700 km V to achieve Triton • Sensors performance with km orbit. No known orbit. No known technology km orbit. No known orbit. No known tech gaps. observ. orbit. ENABLING ready technology technology gaps. gaps. technology gaps. ENABLING • Aerothermal tools and methods 21 22 Ready for Infusion Some Investment Needed Significant Investment Needed

What’s Next?

• Finish what we started within ISPT (shown in “Current Tasks”) • Continue to support (likely only through advocacy) model improvements • Aerothermal and atmospheric • Gather validation data through flight tests; sensor development important (currently unfunded) • Educate about mission benefits and advocate for use • Continue New Frontiers incentive discussions • Request involvement in Titan Flagship Study • Is ISPT ground development + MSL hypersonic guidance + CEV skip entry = Aerocapture validation? • Pursue TPS flight test or Aerocapture flight validation opportunity? • ARMD/ISPT partnership? • New Millennium Program restart? " Bottom line facing Aerocapture: Is flight validation NECESSARY?

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