Nuclear Thermal Rocket Propulsion for Future Human Exploration

Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions presented by Dr. Stanley K. Borowski Chief, Propulsion and Controls Systems Analysis Branch at the Future In-Space Operations (FISO) Colloquium Wednesday, June 27, 2012 Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 1

Nuclear Thermal Rocket (NTR) Concept Illustration (Expander Cycle, Dual LH 2 Turbopumps) NTR: High thrust / high specific impulse (2 x LOX/LH 2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H 2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH 2 tanks, turbopumps, regenerative nozzles and radiation-cooled shirt extensions used -- “NTR is next evolutionary step in high performance liquid rocket engines” During his famous Moon-landing speech in May 1961, President John F. Kennedy also called for accelerated development of the NTR saying this technology “gives promise of some day providing a means of even more exciting and ambitious exploration of space, perhaps beyond the Moon, perhaps to the very end of the solar system itself. ” NERVA-derived Carbide Fuel Ceramic Metal (Cermet) Fuel NTP uses high temperature fuel, produces ~525 MWt (for ~25 klbf engine) but operates for < 85 minutes on a round trip mission to Mars (DRA 5. 0) Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 2

Rover / NERVA* Program Summary (1959 -1972) Tech Demo The smallest engine tested, the 25 klbf “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement • 20 NTR / reactors designed, built and tested at the Nevada Test Site – “All the requirements for a human mission to Mars were demonstrated” • Engine sizes tested – • Larger Cores for Higher Thrust 25, 50, 75 and 250 klb f 2, 350 -2, 550 K (in 25 klbf Pewee) Isp capability – – • Higher Power Fuel Elements H 2 exit temperatures achieved – • System Baseline for NERVA Program 825 -850 sec (“hot bleed cycle” tested on NERVA-XE) 850 -875 sec (“expander cycle” chosen for NERVA flight engine) Burn duration – – ~ 62 min (50 klbf NRX-A 6 - single burn) ~ 2 hrs (50 klbf NRX-XE: 27 restarts / accumulated burn time) --------------* NERVA: Nuclear Engine for Rocket Vehicle Applications Glenn Research Center The NERVA Experimental Engine (XE) demonstrated 28 start-up / shut-down cycles during tests in 1969. Pre-Decisional, For Discussion Purposes Only at Lewis Field 3

“Heritage” Rover / NERVA Homogeneous Thermal Reactor Fuel Element and Tie Tube Bundle Arrangement Hexagonal FE: 0. 75 in across the flats; 35 – 52 in length with 19 coolant channels CVD-coated UC 2 Particles Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 4

Key Elements of the NERVA NTR Engine 0. 75’’ ~35 -52’’ Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 5

Performance Characteristics for Small & Full Size NERVA-Derived Engine Designs – Composite Fuel State-of-the-Art “Pewee” Engine Parameters Glenn Research Center Ref: B. Schnitzler, et al. , “Lower Thrust Engine Options Based on the Small Nuclear Rocket Engine Design”, AIAA-2011 -5846 Pre-Decisional, For Discussion Purposes Only at Lewis Field 6

NOTE: Figure depicts performance regions typically shown for the various fuel options. Fuels can be operated at lower temperature levels to extend fuel life & / or increase engine operational margins. Also, by reducing the fuel loading, higher operating temperatures & specific impulse values are achievable to improve performance Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 7

Trajectory Options for Human Mars Missions SUN Outbound Surface Stay Inbound • Opposition-Class Mission Characteristics (Used in “ 90 -Day” / SEI Mars Studies) – Short Mars stay times (typically 30 - 60 days) – Relatively short round-trip times (400 - 650 days) – Missions always have one short transit leg (either outbound or inbound) and one long transit leg – Long transit legs typically include a Venus swing-by and a closer approach to the Sun (~0. 7 AU or less) – This class trajectory has higher V requirements NOTE: Short orbital stay missions will most likely be chosen for initial human missions to Mars & its moons, Phobos and Deimos • Fast-Conjunction Class Mission Characteristics (Used in DRM 4. 0 & DRA 5. 0) – Long Mars stay times (500 days or more) – Long round trip times (~900 days) – Short “in-space” transit times (~150 to 210 days each way) – Closest approach to the Sun is 1 AU – This class trajectory has more modest V requirements than opposition missions Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 8

NTR Requires Less Propellant than Chemical Systems Mi = Mf exp ( V / g. E Isp) “Rocket Equation” NTR has 100% Higher Isp than Chemical Propulsion where Mi is initial total mass of spacecraft in low Earth orbit (LEO), M i = MSC (spacecraft) + MPL (payload) + Mprop (propellant) and M f = spacecraft mass after a given amount of propellant has been expended in providing a given velocity increment ( V) to the spacecraft, g E = Earth’s gravity = 9. 80665 m/s 2; and I sp = specific impulse (pounds of thrust generated per pound of propellant exhausted per second) Crew Return in MAV Mission Mass Ratio: (RM = Mi /Mf = exp ( V / g. E Isp) NTR is only Propulsion Option besides Chemical to be Tested at Performance Levels needed for a Human Mission to Mars! Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 9

NTR Crewed & Cargo Mars Transfer Vehicles (MTVs) for DRA 5. 0: “ 7 -Launch” Strategy 3 – 25 klbf NDR Engines (Isp ~906 s, T/Weng ~3. 5) Saddle Truss / LH 2 Drop Tank Assembly Payload Element ~65 t (6 crew mission) NOTE: Ares-V Core Stage LH 2 Tank is 10 m D x ~44. 5 m L; two LH 2 tanks cut in ~half with 4 extra end domes provides tanks needed for crewed & 2 cargo MTVs “ 0 -g. E” Crewed MTV: IMLEO ~336. 5 t 3 Ares-V Launches Cargo Lander MTV: IMLEO ~236. 2 t 2 Ares-V Launches Common NTR “Core” Propulsion Stages AC/EDL Aeroshell, Surface PL and Lander Mass ~103 t IMLEO = 808. 9 t Habitat Lander MTV: IMLEO ~236. 2 t 2 Ares-V Launches NOTE: With Chemical IMLEO >1200 t Glenn Research Center (Ref: S. K. Borowski, et al. , AIAA-2009 -5308) Pre-Decisional, For Discussion Purposes Only at Lewis Field 10

NTR Crewed Mars Transfer Vehicle (MTV) Allows NEO Survey and Short Orbital Stay Mars / Phobos Missions 3 – 25 klbf DRA 5. 0 Crewed MTV Options: NTRs • “ 4 -Launch” in-line configuration • Ares-V: 110 t; 9. 1 m OD x 26. 6 m L • IMLEO: ~356. 5 t (6 crew) • Total Mission Burn Time: ~84. 5 min • Largest Single Burn: ~30. 7 min NTP identified as the • No. Restarts: 3 preferred propulsion ---------------------option for DRA 5. 0 • “ 3 -Launch” in-line configuration • Ares-V: 140 t; 10 m OD x 30 m L • IMLEO: 336. 5 t (6 crew) • Total Mission Burn Time: ~79. 2 min • Largest Single Burn: ~44. 6 min • No. Restarts: 3 Phase II Configuration Used in “ 7 -Launch” Mars Mission Option (ESMD AA Cooke) United States’ National Space Policy (June 28, 2010, pg. 11) specifies that NASA shall: By 2025, begin crewed missions beyond the Moon, including sending humans to an asteroid. By the mid-2030 s, send humans to orbit Mars & return them safely to Earth. Glenn Research Center (Ref: Mars DRA 5. 0 Study, NASA-SP-2009 -566, July 2009 ) Pre-Decisional, For Discussion Purposes Only at Lewis Field 11

Growth Paths for DRA 5. 0 “Copernicus” NTR Crewed MTV using Modular Components “Saddle Truss” / LH 2 Drop Tank Assembly 3 – 25 klbf NTRs Common NTR “Core” Propulsion Stages Crewed Payload MMSEV replaces consumables container for NEO missions Applications: • Fast Conjunction Mars Landing Missions – Expendable • “ 1 -yr” Round Trip NEA Missions to 1991 JW (2027), 2000 SG 344 (2028) and Apophis (2028) – Reusable • Propulsion Stage & Saddle Truss / Drop Tank Assembly can also be used as: • Earth Return Vehicle (ERV) / propellant tanker in “Split Mars Mission” Mode – Expendable • Cargo Transfer Vehicle supporting a Lunar Base – Reusable Applications: • Fast Conjunction Mars Landing Missions – Reusable • 2033 Mars Orbital Mission 545 Day Round Trip Time with 60 Days at Mars – Expendable • Cargo & Crew Delivery to Lunar Base – Reusable “In-Line” LH 2 Tank Options for Increasing Thrust: • Add 4 th Engine, or • Transition to LANTR Engines – NTRs with O 2 “Afterburners” Applications: • Faster Transit Conjunction Mars Landing Missions – Reusable • 2033 Mars Orbital Mission 545 Day Round Trip Time with 60 Days at Mars – Expendable • Some LEO Assembly Required – Attachment of Drop Tanks • Additional HLV Launches Transition to “Star Truss” with Drop Tanks to Increase Propellant Capacity Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 12

Reusable Crewed Near Earth Asteroid (NEA) Survey Mission Using NTR Glenn Research Center NEA Exploration (B) MMSEV returns to the ASV NEA Rendezvous NEA Trans-Earth Injection MMSE V Outbound Transit (A) Inbound Transit (C) MMSEV detaches from ASV for close-up inspection / sample gathering Candidate NEAs (TNI): sorties (A/B/C) days • 1991 JW (5/18/27): (112/30/220) • 2000 SG 344 (4/27/28): (104/ 7/216) • Apophis (5/8/28): (268/ 7/ 69) Trans-NEA Injection (TNI) LH 2 Drop Tank Jettisoned After HEEO Initial ASV LEO insertion, capture into a CEV/SM separates from ASV HEEO: 500 km and re-enters x 71, 136 km LEO: 407 km circular Crewed NTR Asteroid Survey Vehicle (ASV) MMSEV attached to ASV’s transfer tunnel Earth Entry Velocity <12. 5 km/sec Crew recovery using CEV 3 HLV Launches Direct Entry Water Landing Earth Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 13

Reusable NTR NEO Survey Mission to 1991 JW Asteroid 1991 JW: D ~490 m Total V = 7. 188 km/s IMLEO ~316. 7 t Mission Times Outbound Stay Return Total Mission 112 days 30 days 220 days 362 days Near Earth Asteroid Orbit • HLV Lift: ~140 t • 10 m OD x 30 m L • Total Mission Burn Time: ~73. 8 min • Largest Single Burn: ~37. 3 min • No. Restarts: 4 6 crew Earth Orbit • Asteroid departure • 10/7/2027 • V = 0. 612 km/s MMSEV • Asteroid arrival • 9/7/2027 • V = 0. 851 km/s • Earth return to 500 km x 71, 136 km HEEO • 5/14/2028 • V = 1. 711 km/s • Earth departure from 407 km circular orbit • 5/18/2027 • V = 4. 014 km/s JSC performed “NEO Accessibility Study” and presented results to ESMD AA on April 7, 2011. Findings: NTR outperformed chemical, SEP/Chemical and all SEP systems, allowing access to more NEOs over larger range of sizes and round trip times for fewer HLV launches. Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 14

2033 Mars Orbital Mission Using “Split Mission” Option (RT Time: 545 days with 60 days at Mars) LEO Configuration Earth Return Vehicle (ERV) / tanker; uses “minimum energy” outbound trajectory LH 2 drop tank jettisoned after TMI Consumables canister transfer tunnel & DM jettisoned before TEI “Switch-over” ERV / crewed PL R&D in Mars Orbit LH 2 for Earth return in “core” propulsion stage LEO Configuration Crewed PL element transferred to ERV for trip back to Earth Outbound crewed MTV; uses higher energy trajectories on “ 1 -way” transit to Mars Crew delivery Orion/SM 32. 2 m 26. 1 m Glenn Research Center 24. 8 m 8. 9 m Earth Return Vehicle (ERV): • IMLEO: ~237. 4 t • HLV Launches: 2 • Total Mission Burn Time: ~64. 2 min • Largest Single Burn: ~25. 4 min • No. Restarts: 3 Outbound Crewed MTV: (6 crew) • IMLEO: ~251. 1 t • HLV Launches: 3 • Total Mission Burn Time: ~47. 7 min • Largest Single Burn: ~25. 2 min • No. Restarts: 3 Total Mission IMLEO: 488. 5 t -------------------------“All-Up” Crewed MTV: (6 crew) • IMLEO: ~429. 4 t • HLV Launches: 4 • Total Mission Burn Time: ~111 min • Largest Single Burn: ~41. 3 min • No. Restarts: 3 (Ref: S. K. Borowski, et al. , 2012 IEEE Aerospace Conference, March 3 -10) Pre-Decisional, For Discussion Purposes Only at Lewis Field 15

Nuclear Thermal Propulsion -- 2033 600 day Mars Transfer Vehicle Core Stage, In-line Tank, & Star Truss w/ 2 LH 2 Drop Tanks Star Truss with 2 LH 2 Drop Tanks, Port & Starboard Core Propulsion Stage Three 25. 1 klbf NTRs Payload: DSH, CEV, Food, Tunnel, etc. Vehicle Mass (mt) / Parameters: In-line Tank Design Constraints / Parameters: • # Engines / Type: 3 / NERVA-derived • Engine Thrust: 25. 1 klbf (Pewee-class) • Propellant: LH 2 • Specific Impulse, Isp: 900 sec • Cooldown LH 2: 3% • Tank Material: Aluminum-Lithium • Tank Ullage: 3% • Tank Trap Residuals: 2% • Truss Material: Graphite Epoxy Composite • RCS Propellants: NTO / MMH • # RCS Thruster Isp: 335 sec (AMBR Isp) • Passive TPS: 1” SOFI + 60 layer MLI • Active CFM: ZBO Brayton Cryo-cooler • I/F Structure: Stage / Truss Docking Adaptor w/ Fluid Transfer Comm. Ant. Mission Constraints / Parameters: • 6 Crew • Outbound time: 183 days (nom. ) • Stay time: 60 days (nom. ) • Return time: 357 days (nom. ) • 1% Performance Margin on all burns • TMI Gravity Losses: 310 m/s total, f(T/W 0) • Pre-mission RCS Vs: 181 m/s (4 burns/stage) • RCS Mid. Crs. Cor. Vs: 65 m/s (in & outbnd) • Jettison Both Drop Tanks After TMI-1 • Jettison Tunnel, Can & Waste Prior to TEI NTP Transfer Vehicle Description: NTP system consists of 3 elements: 1) core propulsion stage, 2) in-line tank, and 3) integrated star truss and dual drop tank assembly that connects the propulsion stack to the crewed payload element for Mars 2033 mission. Each 100 t element is delivered on an SLS LV (178. 35. 01, 10 m O. D. x 25. 2 m cyl. §) to LEO -50 x 220 nmi, then onboard RCS provides circ burn to 407 km orbit. The core stage uses three NERVA-derived 25. 1 klbf engines. It also includes RCS, avionics, power, long-duration CFM hardware (e. g. , COLDEST design, ZBO cryo-coolers) and AR&D capability. The star truss uses Gr/Ep composite material & the LH 2 drop tanks use a passive TPS. Interface structure includes fluid transfer, electrical, and communications lines. Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 16

Artificial Gravity Bimodal NTR MTV Option: “Copernicus –B” • 3 – 25 klbf Cermet-fuel BNTRs (ESCORT) • Tex ~2700 K, Isp ~911 s, T/Weng ~5. 52 • IMLEO ~330 t (6 crew) • Total Mission Burn Time: ~77. 4 min • Largest Single Burn: ~43. 5 min • No. Restarts: 3 • Copernicus – B is an AG version of DRA 5. 0 ” 0 -g. E” NTR crewed MTV that uses its BNTR engines to generate both high thrust & electrical power. No large Sun-tracking PVAs (~3. 5 t) are required • Copernicus – B uses 3 – 25 k. We Brayton Rotating Units (~2. 63 t) each operating at 2/3 rd of rated power (~17 k. We ) to produce the 50 k. We needed to operate the MTV • Brayton units are located within the propulsion stage thrust structure that also supports an ~71 m 2 conical radiator mounted to its exterior Vehicle rotation about its center-of-mass provides AG environment for the crew out to Mars and back • Vehicle rotation at 3. 0 – 5. 2 rpm provides a 0. 38 – 1 g. E AG environment for the crew. A Mars gravity field is provided on the outbound mission leg. On the inbound leg, the rotation rate is gradually increased to help the crew readjust to Earth’s gravity level Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 17

Nuclear Thermal Propulsion: “The Next Evolutionary Step” in High Performance Liquid Rocket Propulsion • NTP provides high thrust (10’s of klbf) with an ~100% increase in Isp over LOX/LH 2 chemical propulsion (from 450 to 900 s) • NTP can transition to higher temperature binary, then ternary carbide fuels (Isp ~950 - 1050 s) Aerojet / GRC Non-Nuclear O 2 “Afterburner” Nozzle Test • “Bimodal” engines (BNTR) can produce modest electrical power (~15 -25 k. We) to run the spacecraft eliminating large Sun-tracking PVAs and allowing AG to improve crew health and fitness • The NTP engine can also be outfitted with an “LOX-Afterburner” nozzle and propellant feed system allowing supersonic combustion downstream of the nozzle throat thereby enabling variable thrust and Isp operation depending on the O/H mixture used NTR Options Exist for Power Generation, Thrust Augmentation & Hybrid Propulsion • The “LOX-Augmented” NTR (LANTR) can utilize extraterrestrial sources of H 2 O, ice to extend the range of human exploration throughout the Solar System without the need for very advanced, lower TRL systems • Coupling higher power BNTRs with EP in “hybrid” ~1. 0 MWe BNTEP system offers performance comparable to low , 10 MWe “all NEP” system “Revolutionary Capability in an Evolutionary Manner” Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 18

Notional NTP Foundational Technology Development and System Technology Demonstration Schedule AES NCPS Project is Focused on Foundational Technology Development Lunar NTR Stage SOTA Reactor Core & Engine Modeling NERVA “Composite” Fuel Crew Return in MAV Small “Fuel-Rich” Engine Hot Gas Source Fuel Element Irradiation Testing in ATR at INL “Cermet” Fuel Ground & Flight Technology Demonstrators Affordable SAFE Ground Testing at the Nevada Test Site (NTS) NTR Element Environmental Simulator (NTREES) Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 19

Size Comparison of RL 10 B-2 and Lower Thrust NTR Engine Designs NTR “Key Performance Parameters” (KPPs): Tex ~2700 K, pch ~1000 psia, Nozzle Area Ratio (�) ~300: 1, Isp ~910 s Mission versatility increased with smaller (15 -25 klbf) NTR engines; the time and cost to design, build, test and fly is also reduced RL 10 B-2 KPPs: Tex ~3167 K, pch ~620 psia, Nozzle AR (�) ~285: 1, Isp ~463 s RL 10 B-2 Fvac: 24. 75 -klbf Used on Delta IV DRA 5. 0 GTD / FTD Engine in 2020 / 2023 Fvac: 15 -klbf ~315 MWt Fvac: 5 - 7. 5 -klbf ~105 MWt 6. 23 m 20. 5 ft 5. 36 m 17. 6 ft 4. 19 m 13 ft Fvac: 25 -klbf ~525 MWt 4. 27 m 13. 9 ft 2. 16 m 7 ft 0. 84 m 2. 74 ft 1. 45 m 4. 75 ft 1. 87 m 6. 13 ft Ref: Russ Joyner, PWR Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 20

NTP Stage Approach for Flight Demo Atlas 5 • Delta 4 NTP stage concept can be leveraged from Delta 4 DCSS of the same diameter and approx. length ~40 -ft (12. 2 m) ~16 -ft (5 m) Use Elements of LO 2/LH 2 Delta 4 Cryogenic Second Stage (DCSS) • Remove LO 2 Tank, Lines, Valves • Remove RL 10 B-2 • Add small NTP with retractable nozzle skirt • Increase LH 2 lines • Similar thrust structure NTP Cryogenic Stage for FTD can Be Made Affordable via Delta 4 Cryogenic Second Stage Components Glenn Research Center 2012 GLEX Conference, Washington, DC, May 22 - 24 Pre-Decisional, For Discussion Purposes Only at Lewis Field 21

Frequently Asked Questions about NTP • Launching Nuclear Systems: • Fission reactor systems (fission surface power or NTR engines) have negligible quantities of radioactive material within them prior to being operated (few 100 Curies vs 400, 000 Curies in Cassini’s 3 RTGs) • Fission product buildup only becomes appreciable at the end of the TMI burn as the MTV is departing Earth orbit for heliocentric space • Fission systems designed to generate thermal power not to explode • Inadvertent criticality accidents prevented by design safety features (e. g. , neutron poison wires, control drum interlocks) or reactor design (e. g. , cermet fuel NTR operating on fast neutrons) • Improvements in fuel element CVD coatings and claddings expected to significantly reduce or eliminate fission product gas release within the engine’s hydrogen exhaust • Cost for Engine Development & Ground Testing will not “break the bank”: • Separate effects tests (non-nuclear, hot H 2 testing under prototypic operating conditions -- pch, temperature & H 2 flow -- in NTREES followed irradiation testing in ATR will validate fuel element design • Small engine (5 klbf) scalable to higher thrust levels will be developed, ground, then flight tested first using a common fuel element design • Lower thrust-class engines (up to ~25 klbf) can use / adapt existing RL 10 -derived engine components (per discussions with PWR) • Small engine size, SAFE ground test approach and use of Nevada Test Site (NTS) assets (e. g. , Device Assembly Facility for 0 -power critical tests), mobile control trailers, etc. , rather than large fixed test structures, indicate lower costs. Recent estimates (Dec. 2011) from the NSTec and the NTS for the SAFE capital cost are ~45 M$ (site and all supporting equipment) with ~2 M$ recurring cost for each additional engine test Phoebus-2 A – 5000 MWt / 250 klbf NTR engine being transported to Test Cell C at NTS in 1968. Note technicians riding at the front of the engine NTR Element Environmental Simulator (NTREES) Affordable SAFE Testing SAFE: Subsurface Active Filtration of Exhaust; also know as “Borehole” Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 22 22

Summary of Results and Key “Take Away” Points on NTP • Nuclear Thermal Propulsion (NTP) is a proven technology; 20 NTR / reactors designed, built and tested at the Nevada Test Site (NTS) in the Rover / NERVA programs • “All the requirements for a human mission to Mars were demonstrated” – thrust level, hydrogen exhaust temperature, max burn duration, total burn time at power, #restarts • The smallest engine tested in the Rover program, the 25 klbf “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement – No major scale ups are required as with other advanced propulsion / power systems • In less than 5 years, 4 different thrust engines tested (50, 75, 250, 25 klbf – in that order) using a common fuel element design – Pewee was the highest performing engine • “Common fuel element” approach used in the AISP / NCPS projects to design a small (~7. 5 klbf), affordable engine for ground testing by 2020 followed by a flight technology demonstration mission in 2023. PWR sees strong synergy between NTP and chemical • SAFE (Subsurface Active Filtration of Exhaust) ground testing at NTS is baseline; capital cost for test HDW is ~45 M$ with ~ 2 M$ for each additional engine test (NTS Dec. 2011) • Cost for engine development and ground testing will not “break the bank” & the system will have broad application ranging from robotic to human exploration missions Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 23

Summary of Results and Key “Take Away” Points on NTP • NTP consistently identified as “preferred propulsion option” for human Mars missions: - NASA’s SEI – Stafford Report (1991) listed NTP as #2 priority after HLV - NASA’s Mars Design Reference Missions (DRMs) 1 (1993) – 4 (1999) - NASA’s Design Reference Architecture (DRA) 5. 0 (2009) • Using NTP, the launch mass savings over “All Chemical” and “Chemical / Aerobrake” systems amounts to 400+ metric tons (~ISS mass) or ~4 or more HLVs. At ~1 B$ per HLV, the launch vehicle cost savings alone can pay for NTP development effort • The DRA 5. 0 crewed MTV “Copernicus” has significant capability allowing reusable “ 1 -yr” NEA missions & short (~1. 5 yrs) Mars / Phobos orbital missions before a landing • JSC’s “NEA Accessibility Study” presented by Bret Drake to Doug Cooke (April 7, 2011). Findings: NTR outperforms chemical, SEP/Chemical & all SEP systems, allowing access to more NEAs over larger range of sizes and round trip times for fewer HLV launches. • With more LH 2, faster “ 1 -way” transit times to from Mars are possible if desired • Lastly, NTP has significant growth capability (other fuels, bimodal & LANTR operation) Glenn Research Center Pre-Decisional, For Discussion Purposes Only at Lewis Field 24