at lewis field glenn research center pre-decisional, for discussion purposes only nuclear thermal...

at Lewis FieldGlenn Research Center

Pre-Decisional, For Discussion Purposes Only

Nuclear Thermal Rocket Propulsion for Future Human Exploration Missions

presented by

Dr. Stanley K. BorowskiChief, Propulsion and Controls Systems

Analysis Branch

at the

Future In-Space Operations (FISO)Colloquium

Wednesday, June 27, 2012

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NTR: High thrust / high specific impulse (2 x LOX/LH2 chemical) engine uses high power density fission reactor with enriched uranium fuel as thermal power source. Reactor heat is removed using H2 propellant which is then exhausted to produce thrust. Conventional chemical engine LH2 tanks, turbopumps, regenerative nozzles and radiation-cooledshirt extensions used -- “NTR is next evolutionary step in high performance liquid rocket engines”

Nuclear Thermal Rocket (NTR) Concept Illustration(Expander Cycle, Dual LH2 Turbopumps)

NERVA-derivedCarbide 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)

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.”

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The NERVA Experimental Engine (XE) demonstrated 28 start-up / shut-down cycles during tests in 1969.

Tech DemoSystem Baseline for

NERVA ProgramHigher PowerFuel Elements

Larger Coresfor 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 klbf

• H2 exit temperatures achieved– 2,350-2,550 K (in 25 klbf Pewee)

• Isp capability– 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-A6 - single burn)– ~ 2 hrs (50 klbf NRX-XE: 27 restarts

/ accumulated burn time)

-----------------------------* NERVA: Nuclear Engine for Rocket Vehicle Applications

The smallest engine tested, the 25 klbf “Pewee” engine, is sufficient for human Mars missions when used in a clustered engine arrangement

Rover / NERVA* Program Summary (1959-1972)

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CVD-coated UC2 Particles Hexagonal FE: 0.75 in across the flats; 35 – 52 in length with 19 coolant channels

“Heritage” Rover / NERVA Homogeneous Thermal ReactorFuel Element and Tie Tube Bundle Arrangement

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Key Elements of the NERVA NTR Engine

0.75’’

~35-52’’

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Performance Characteristics for Small & Full Size NERVA-Derived Engine Designs – Composite Fuel

Ref: B. Schnitzler, et al., “Lower Thrust Engine Options Basedon the Small Nuclear Rocket Engine Design”, AIAA-2011-5846

State-of-the-Art “Pewee”Engine Parameters

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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

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Trajectory Options for Human Mars Missions

SUN

• 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 requirementsNOTE: 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

OutboundSurface StayInbound

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Crew Return in MAV

Mission Mass Ratio:(RM = Mi /Mf =

exp (V / gE Isp)

NTR Requires Less Propellant than Chemical Systems

Mi = Mf exp (V / gE Isp) “Rocket Equation”

where Mi is initial total mass of spacecraft in low Earth orbit (LEO), Mi = MSC (spacecraft) + MPL (payload) + Mprop (propellant) and Mf = spacecraft mass after a given amount of propellant has been expended in providing a given velocity increment (V) to the spacecraft, gE = Earth’s gravity = 9.80665 m/s2; and Isp = 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 Ispthan Chemical Propulsion

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“0-gE” 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 = 808.9 t

NOTE: Ares-V Core Stage LH2 Tank is 10 m D x ~44.5 m L; two LH2 tanks cut in~half with 4 extra end domes provides tanks needed for crewed & 2 cargo MTVs

3 – 25 klbf NDR Engines(Isp ~906 s, T/Weng ~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 / LH2 Drop Tank Assembly

(Ref: S.K. Borowski, et al., AIAA-2009-5308)

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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 -------------------------------------------• “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

3 – 25 klbf

NTRs

Phase II Configuration

Configuration Used in“7-Launch” Mars MissionOption (ESMD AA Cooke)

(Ref: Mars DRA 5.0 Study, NASA-SP-2009-566, July 2009 )

NTP identified as the preferred propulsion option for DRA 5.0

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Growth Paths for DRA 5.0 “Copernicus” NTR Crewed MTVusing 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 4th Engine, or• Transition to LANTR Engines – NTRs with O2 “Afterburners”

3 – 25 klbf NTRs

Transition to “Star Truss”with Drop Tanks to Increase

Propellant Capacity

“Saddle Truss” / LH2

Drop Tank Assembly

“In-Line” LH2 Tank

Crewed Payload

Common NTR “Core”

Propulsion Stages

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Crewed NTR Asteroid Survey Vehicle (ASV)

Earth

NEA

NEA Rendezvous

LH2 Drop Tank Jettisoned

OutboundTransit (A)

Earth Entry Velocity

<12.5 km/sec

Trans-NEA Injection (TNI)

MMSEV returns to the ASV

InboundTransit (C)

LEO: 407 km circular

MMSEV

Crew recovery using CEV

HEEO: 500 km x 71,136 km

Initial ASVcapture 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 EntryWater Landing

MMSEV attached to ASV’stransfer 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 ASVand re-enters

Trans-Earth Injection

Pre-Decisional, For Discussion Purposes Only13



Reusable NTR NEO Survey Mission to 1991 JW

• Asteroid arrival• 9/7/2027• V = 0.851 km/s

• Asteroid departure• 10/7/2027• V = 0.612 km/s

Near Earth Asteroid Orbit

Earth Orbit

• Earth departure from 407 km circular orbit • 5/18/2027• V = 4.014 km/s

• Earth return to 500 km x 71,136 km HEEO• 5/14/2028• V = 1.711 km/s

Asteroid 1991 JW: D ~490 mTotal V = 7.188 km/s

Mission Times

Outbound 112 daysStay 30 daysReturn 220 daysTotal Mission 362 days

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.

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

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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

LH2 drop tankjettisoned after TMI

Crewed PL elementtransferred to ERVfor trip back to Earth

LEO Configuration

LEO Configuration

Consumables canistertransfer tunnel & DM jettisoned before TEI

LH2 for Earth return in“core” propulsion stage

Crew deliveryOrion/SM

Outbound crewed MTV; useshigher energy trajectorieson “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: 3Total 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)

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• # 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 LH2 Drop Tanks, Port & Starboard

Three 25.1 klbf 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 (178.35.01, 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/W0)• 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:

V (m/s)Burn Time

(min)1st perigee TMI + g-loss 2669 46.8

2nd perigee TMI 1201 15.1MOC 1470 15.5

TEI 3080 24.3Mission Total 8420 101.6

Comm. Ant.

Core Stage mtDry Mass 37.78 Three 25klbg NTP Engines 12.68 Three External Radiation Shields 6.04 Tank Mass 15.34LH2 Mass 56.73RCS Prop Load 5.69Total Launch Mass 100.19Tank Fraction (%) 31.7%Element Length (m) 26.23

Inline Stage mtDry Mass 19.45 Tank Mass 17.05LH2 Mass 70.71RCS Prop Load 9.43Total Launch Mass 99.60Tank Fraction (%) 30.3%Element Length (m) 25.18

Star Truss and Drop Tank (2) mtDry Mass 18.77 Truss Mass 6.07 Tank Mass 14.24LH2 Mass 64.85RCS Prop Load 4.46Total Launch Mass 88.08Tank Fraction (%) 24.9%Element Length (m) 19

Payload mtTotal Launch Mass 71.54Crew 0.79Short Saddle Truss 5.08RCS Prop Load 4.07Element Length (m) 21.65Total Launched Mass 447.55

Nuclear Thermal Propulsion -- 2033 600 day Mars Transfer VehicleCore Stage, In-line Tank, & Star Truss w/ 2 LH2 Drop Tanks

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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-gE” 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 kWe Brayton Rotating Units (~2.63 t) each operating at 2/3rd of rated power (~17 kWe ) to produce the 50 kWe needed to operate the MTV

• Brayton units are located within the propulsion stage thrust structure that also supports an ~71 m2 conical radiator mounted to its exterior

• Vehicle rotation at 3.0 – 5.2 rpm provides a 0.38 – 1gE 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 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

Artificial Gravity Bimodal NTR MTV Option:“Copernicus –B”

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“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/LH2 chemical propulsion (from 450 to 900 s)

• NTP can transition to higher temperature binary, then ternary carbide fuels (Isp ~950 - 1050 s)

• “Bimodal” engines (BNTR) can produce modest electrical power (~15-25 kWe) to run the space-craft eliminating large Sun-tracking PVAs andallowing 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 H2O, 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

Aerojet / GRC Non-Nuclear O2 “Afterburner” Nozzle Test

Nuclear Thermal Propulsion: “The Next Evolutionary Step” in High Performance Liquid Rocket Propulsion

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Crew Return in MAV

NTR Element Environmental Simulator (NTREES)

Ground & Flight Technology Demonstrators

AES NCPS Project is Focused on Foundational Technology Development

Fuel Element IrradiationTesting in ATR at INL

Lunar NTR Stage

Affordable SAFE Ground Testingat 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

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Size Comparison of RL 10B-2 and Lower Thrust NTR Engine Designs

Fvac:15-klbf

~315 MWt

4.27 m13.9 ft

0.84 m2.74 ft

1.87 m6.13 ft

Fvac:25-klbf

~525 MWt

6.23 m20.5 ft

4.19 m13 ft

2.16 m7 ft

RL10B-2Fvac: 24.75-klbf Used on Delta IV

RL10B-2 KPPs: Tex ~3167 K, pch ~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 m17.6 ft

1.45 m4.75 ft

Fvac: 5 - 7.5-klbf~105 MWt

GTD / FTD Enginein 2020 / 2023

NTR “Key Performance Parameters” (KPPs): Tex ~2700 K, pch ~1000 psia, Nozzle Area Ratio () ~300:1, Isp ~910 s

Ref: Russ Joyner, PWR

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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-2Use 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 22 - 24 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 H2 testing under prototypic operating conditions -- pch, 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”

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Phoebus-2A – 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

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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 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 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 ~ 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

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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 LH2, faster “1-way” transit times to from Mars are possible if desired

• Lastly, NTP has significant growth capability (other fuels, bimodal & LANTR operation)

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at lewis field glenn research center pre-decisional, for discussion purposes only nuclear thermal...

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