at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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,

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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” Nuclear Thermal Rocket (NTR) Concept Illustration (Expander Cycle, Dual LH 2 Turbopumps) NERVA-derived Carbide Fuel Ceramic Metal (Cermet) Fuel NTP uses high temperature fuel, produces ~525 MWt (for ~25 klb f engine) but operates for < 85 minutes on a round trip mission to Mars (DRA 5.0) 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.” 2

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only The NERVA Experimental Engine (XE) demonstrated 28 start-up / shut-down cycles during tests in Tech Demo System Baseline for NERVA Program Higher Power Fuel Elements Larger Cores for Higher Thrust 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 –25, 50, 75 and 250 klb f H 2 exit temperatures achieved –2,350-2,550 K (in 25 klb f Pewee) I sp capability – sec (“hot bleed cycle” tested on NERVA-XE) – sec (“expander cycle” chosen for NERVA flight engine) Burn duration –~ 62 min (50 klb f NRX-A6 - single burn) –~ 2 hrs (50 klb f NRX-XE: 27 restarts / accumulated burn time) * NERVA: Nuclear Engine for Rocket Vehicle Applications The smallest engine tested, the 25 klb f “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement Rover / NERVA* Program Summary ( ) 3

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only CVD-coated UC 2 Particles Hexagonal FE: 0.75 in across the flats; 35 – 52 in length with 19 coolant channels “Heritage” Rover / NERVA Homogeneous Thermal Reactor Fuel Element and Tie Tube Bundle Arrangement 4

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Key Elements of the NERVA NTR Engine 0.75’’ ~35-52’’ 5

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

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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 7

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Trajectory Options for Human Mars Missions  SUN  Opposition-Class Mission Characteristics (Used in “90-Day” / SEI Mars Studies) –Short Mars stay times (typically days) –Relatively short round-trip times ( 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 Outbound Surface Stay Inbound 8

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Crew Return in MAV Mission Mass Ratio: (R M = M i /M f = exp (  V / g E I sp ) NTR Requires Less Propellant than Chemical Systems M i = M f exp (  V / g E I sp ) “Rocket Equation” where M i is initial total mass of spacecraft in low Earth orbit (LEO), M i = M SC (spacecraft) + M PL (payload) + M prop (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 = m/s 2 ; and I sp = specific impulse (pounds of thrust generated per pound of propellant exhausted per second) NTR is only Propulsion Option besides Chemical to be Tested at Performance Levels needed for a Human Mission to Mars! NTR has 100% Higher Isp than Chemical Propulsion 9

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only “0-g E ” Crewed MTV:  IMLEO ~336.5 t  3 Ares-V Launches Cargo Lander MTV:  IMLEO ~236.2 t  2 Ares-V Launches Habitat Lander MTV:  IMLEO ~236.2 t  2 Ares-V Launches NTR Crewed & Cargo Mars Transfer Vehicles (MTVs) for DRA 5.0: “7-Launch” Strategy  IMLEO = t 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 3 – 25 klb f NDR Engines (I sp ~906 s, T/W eng ~3.5) Common NTR “Core” Propulsion Stages AC/EDL Aeroshell, Surface PL and Lander Mass ~103 t Payload Element ~65 t (6 crew mission) NOTE: With Chemical IMLEO >1200 t Saddle Truss / LH 2 Drop Tank Assembly (Ref: S.K. Borowski, et al., AIAA ) 10

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only NTR Crewed Mars Transfer Vehicle (MTV) Allows NEO Survey and Short Orbital Stay Mars / Phobos Missions 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-2030s, send humans to orbit Mars & return them safely to Earth. DRA 5.0 Crewed MTV Options: “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 No. Restarts: “3-Launch” in-line configuration Ares-V: 140 t; 10 m OD x 30 m L IMLEO: t (6 crew) Total Mission Burn Time: ~79.2 min Largest Single Burn: ~44.6 min No. Restarts: 3 3 – 25 klb f NTRs Phase II Configuration Configuration Used in “7-Launch” Mars Mission Option (ESMD AA Cooke) (Ref: Mars DRA 5.0 Study, NASA-SP , July 2009 ) NTP identified as the preferred propulsion option for DRA

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Growth Paths for DRA 5.0 “Copernicus” NTR Crewed MTV using Modular Components Applications: Fast Conjunction Mars Landing Missions – Expendable “1-yr” Round Trip NEA Missions to 1991 JW (2027), 2000 SG344 (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 MMSEV replaces consumables container for NEO missions 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 Options for Increasing Thrust: Add 4 th Engine, or Transition to LANTR Engines – NTRs with O 2 “Afterburners” 3 – 25 klb f NTRs Transition to “Star Truss” with Drop Tanks to Increase Propellant Capacity “Saddle Truss” / LH 2 Drop Tank Assembly “In-Line” LH 2 Tank Crewed Payload Common NTR “Core” Propulsion Stages 12

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Crewed NTR Asteroid Survey Vehicle (ASV) Earth NEA NEA Rendezvous LH 2 Drop Tank Jettisoned Outbound Transit (A) Earth Entry Velocity <12.5 km/sec Trans-NEA Injection (TNI) MMSEV returns to the ASV Inbound Transit (C) LEO: 407 km circular MMSEV Crew recovery using CEV HEEO: 500 km x 71,136 km Initial ASV capture into a Glenn Research Center Reusable Crewed Near Earth Asteroid (NEA) Survey Mission Using NTR MMSEV detaches from ASV for close-up inspection / sample gathering sorties Direct Entry Water Landing MMSEV attached to ASV’s transfer tunnel NEA Exploration (B) 1991 JW (5/18/27): (112/30/220) 2000 SG344 (4/27/28): (104/ 7/216) Apophis (5/8/28): (268/ 7/ 69) Candidate NEAs (TNI): (A/B/C) days 3 HLV Launches After HEEO LEO insertion, CEV/SM separates from ASV and re-enters Trans-Earth Injection Pre-Decisional, For Discussion Purposes Only 13

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Reusable NTR NEO Survey Mission to 1991 JW Asteroid arrival 9/7/2027  V = km/s Asteroid departure 10/7/2027  V = km/s  Near Earth Asteroid Orbit Earth Orbit Earth departure from 407 km circular orbit 5/18/2027  V = km/s Earth return to 500 km x 71,136 km HEEO 5/14/2028  V = km/s Asteroid 1991 JW: D ~490 m Total  V = km/s Mission Times Outbound 112 days Stay 30 days Return 220 days Total Mission 362 days JSC performed “NEO Accessibility Study” and presented results to ESMD AA on April 7, 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. IMLEO ~316.7 t MMSEV 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 14

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 2033 Mars Orbital Mission Using “Split Mission” Option (RT Time: 545 days with 60 days at Mars) 32.2 m 26.1 m 24.8 m 8.9 m Earth Return Vehicle (ERV) / tanker; uses “minimum energy” outbound trajectory ERV / crewed PL R&D in Mars Orbit LH 2 drop tank jettisoned after TMI Crewed PL element transferred to ERV for trip back to Earth LEO Configuration Consumables canister transfer tunnel & DM jettisoned before TEI LH 2 for Earth return in “core” propulsion stage Crew delivery Orion/SM Outbound crewed MTV; uses higher energy trajectories on “1-way” transit to Mars “Switch-over” 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: 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) 15

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only # Engines / Type:3 / NERVA-derived Engine Thrust:25.1 klbf (Pewee-class) Propellant:LH2 Specific Impulse, Isp:900 sec Cooldown LH2:3% 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 Core Propulsion Stage Star Truss with 2 LH 2 Drop Tanks, Port & Starboard Three 25.1 klb f NTRs 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 100t element is delivered on an SLS LV ( , 10m 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 LH2 drop tanks use a passive TPS. Interface structure includes fluid transfer, electrical, and communications lines. NTP Transfer Vehicle Description: Design 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 MidCrs. Cor.  Vs:65 m/s (in & outbnd) Jettison Both Drop Tanks After TMI-1 Jettison Tunnel, Can & Waste Prior to TEI Mission Constraints / Parameters: In-line Tank Payload: DSH, CEV, Food, Tunnel, etc. Vehicle Mass (mt) / Parameters: Comm. Ant. Nuclear Thermal Propulsion day Mars Transfer Vehicle Core Stage, In-line Tank, & Star Truss w/ 2 LH 2 Drop Tanks 16

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Vehicle rotation about its center-of-mass provides AG environment for the crew out to Mars and back 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 kW e Brayton Rotating Units (~2.63 t) each operating at 2/3 rd of rated power (~17 kW e ) to produce the 50 kW e 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 at 3.0 – 5.2 rpm provides a 0.38 – 1g 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 3 – 25 klb f Cermet-fuel BNTRs (ESCORT) T ex ~2700 K, Isp ~911 s, T/W eng ~5.52 IMLEO ~330 t (6 crew) Total Mission Burn Time: ~77.4 min Largest Single Burn: ~43.5 min No. Restarts: 3 Artificial Gravity Bimodal NTR MTV Option:“Copernicus –B” 17

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only “Revolutionary Capability in an Evolutionary Manner” NTR Options Exist for Power Generation, Thrust Augmentation & Hybrid 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 ~ s) “Bimodal” engines (BNTR) can produce modest electrical power (~15-25 kW e ) to run the space- craft 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 down- stream of the nozzle throat thereby enabling variable thrust and Isp operation depending on the O/H mixture used 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 MW e BNTEP system offers performance comparable to low , 10 MW e “all NEP” system Aerojet / GRC Non-Nuclear O 2 “Afterburner” Nozzle Test Nuclear Thermal Propulsion: “The Next Evolutionary Step” in High Performance Liquid Rocket Propulsion 18

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

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only Size Comparison of RL 10B-2 and Lower Thrust NTR Engine Designs Fvac: 15-klbf ~315 MWt 4.27 m 13.9 ft 0.84 m 2.74 ft 1.87 m 6.13 ft Fvac: 25-klbf ~525 MWt 6.23 m 20.5 ft 4.19 m 13 ft 2.16 m 7 ft RL10B-2 Fvac: klbf Used on Delta IV RL10B-2 KPPs: T ex ~3167 K, p ch ~620 psia, Nozzle AR (  ) ~285:1, Isp ~463 s DRA 5.0 Mission versatility increased with smaller (15-25 klbf) NTR engines; the time and cost to design, build, test and fly is also reduced 5.36 m 17.6 ft 1.45 m 4.75 ft Fvac: klbf ~105 MWt GTD / FTD Engine in 2020 / 2023 NTR “Key Performance Parameters” (KPPs): T ex ~2700 K, p ch ~1000 psia, Nozzle Area Ratio (  ) ~300:1, Isp ~910 s Ref: Russ Joyner, PWR 20

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only NTP Stage Approach for Flight Demo NTP stage concept can be leveraged from Delta 4 DCSS of the same diameter and approx. length ~40-ft (12.2 m) Remove LO2 Tank, Lines, Valves Remove RL10B-2 Use Elements of LO2/LH2 Delta 4 Cryogenic Second Stage (DCSS) Atlas 5 Delta 4 ~16-ft (5 m) Add small NTP with retractable nozzle skirt Increase LH2 lines Similar thrust structure NTP Cryogenic Stage for FTD can Be Made Affordable via Delta 4 Cryogenic Second Stage Components 2012 GLEX Conference, Washington, DC, May

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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 H2 testing under prototypic operating conditions -- p ch, temperature & H2 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 RL10-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 SAFE: Subsurface Active Filtration of Exhaust; also know as “Borehole” 22 Phoebus-2A – 5000 MWt / 250 klbf NTR engine being transported to Test Cell C at NTS in Note technicians riding at the front of the engine NTR Element Environmental Simulator (NTREES) Affordable SAFE Testing 22

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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 klb f “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 klb f – 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 klb f ), affordable engine for ground testing by 2020 followed by a flight technology demonstration mission in 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 ~ 2M$ 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 23

at Lewis Field Glenn Research Center Pre-Decisional, For Discussion Purposes Only 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 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) 24