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Nuclear Thermal Propulsion National Aeronautics and -~- Space Administration ~ . :· Nuclear Thermal Propulsion: An Overview of NASA Development Efforts Ryan Wilkerson| Presented at Missouri S&T | 1031.19 Nuclear Thermal Propulsion Background How do we assess engine performance? Thrust Specific Impulse Rocket Thrust Equation Total Impulse 푭 = 풎ሶ 푽풆 + ( 풑풆 − 풑풐) 푨풆 푰 = 푭∆풕 = 풎푽풆 Equivalent Velocity Specific Impulse ( 풑풆 − 풑풐) 푨풆 푰 푽풆풒 푭 푽풆풒 = 푽풆 + 푰풔풑 = = = 풎ሶ 풎품풐 품풐 풎ሶ 품풐 푭 = 풎ሶ 푽풆풒 Thrust is an indicator of how Specific impulse indicates how hard an engine can push efficiently an engine uses propellant Current engine architectures Chemical Engines Advanced Propulsion SRB-SS SRB-SLS F-1 J-2 RS-25 Ion NEP NTP Sea Level Thrust 2800 3600 1800 109.3 418 - - (klbf) Vacuum Thrust (klbf) - - 2020.7 232.3 512.3 2E-5 - 2E-2 25-250 Sea Level ISP (s) 242 269 269.7 200 366 - - Vacuum ISP (s) - - 303.1 421 452.3 2000-8000 800-1000 Propellant PBAN-APCP PBAN-APCP LO2/RP-1 LO2/LH2 LO2/LH2 Xe LH2 Saturn V Current engine architectures Chemical Engines Advanced Propulsion SRB-SS SRB-SLS F-1 J-2 RS-25 Ion NEP NTP Sea Level Thrust 2800 3600 1800 109.3 418 - - (klbf) Vacuum Thrust (klbf) - - 2020.7 232.3 512.3 2E-5 - 2E-2 25-250 Sea Level ISP (s) 242 269 269.7 200 366 - - Vacuum ISP (s) - - 303.1 421 452.3 2000-8000 800-1000 Propellant PBAN-APCP PBAN-APCP LO2/RP-1 LO2/LH2 LO2/LH2 Xe LH2 L-- Space Shuttle SOLID ROCKET BOOSTER-SRI Current engine architectures Chemical Engines Advanced Propulsion SRB-SS SRB-SLS F-1 J-2 RS-25 Ion NEP NTP Sea Level Thrust 2800 3600 1800 109.3 418 - - (klbf) Vacuum Thrust (klbf) - - 2020.7 232.3 512.3 2E-5 - 2E-2 25-250 Sea Level ISP (s) 242 269 269.7 200 366 - - Vacuum ISP (s) - - 303.1 421 452.3 2000-8000 800-1000 Propellant PBAN-APCP PBAN-APCP LO2/RP-1 LO2/LH2 LO2/LH2 Xe LH2 Space Launch System Current engine architectures Chemical Engines Advanced Propulsion SRB-SS SRB-SLS F-1 J-2 RS-25 Ion NEP NTP Sea Level Thrust 2800 3600 1800 109.3 418 - - (klbf) Vacuum Thrust (klbf) - - 2020.7 232.3 512.3 2E-5 - 2E-2 25-250 Sea Level ISP (s) 242 269 269.7 200 366 - - Vacuum ISP (s) - - 303.1 421 452.3 2000-8000 800-1000 Propellant PBAN-APCP PBAN-APCP LO2/RP-1 LO2/LH2 LO2/LH2 Xe LH2 Ion Engine Mars Transit Vehicle There are many uses for nuclear power in space Radioisotope Thermal Electric Generators Nuclear Electric Propulsion Nuclear Thermal Propulsion Direct Heating of Propellant Electricity Generation to Power Thruster to Provide Thrust )He alpha patl!Cle 8 Nuclear thermal rocket engine NUCLEAR REAClOR 950 900 Note: results based on 850 800 • 1000 psia chamber 750 pressure 700 • 2700 K chamber ~ 650 .§ temperature 0 600 = 550 • 150:1 nozzle area ratio "§ g 500 ~ 450 i 400 350 300 250 CONlROL DRUM Specific Impulse (revisited) 1500Nuclear Thermal Propulsion with Various Propellants 휸−ퟏ 푭 ퟏ ퟐ휸 푹푻 풑 휸 - H2 푰 = = ퟏ − 풆 풔풑 1300 풎ሶ 품풐 -J--H품풐 휸 − ퟏ 푴 -1-풑풄 - NH 3 - H20 푻 1100 푰 ∝ - CO2 풔풑 푴 f ~ 900 Cl. -NTP engines produce thrust by heating ~ 700 propellant using a nuclear core -propellant temperature directly correlates 500 300 to Isp -core power and temperature determine 100 exhaust temperature and therefore Isp NTP mission proposals, from the moon to Mars Design Transition from Single Large NTR to Clustered Smaller Engines Supplying Modest Electrical Power Expendable TLI Stage for First Lunar Outpost Mission using Clustered 25 klbf Engines -- “Fast Track Study” (1992) Zero-Gravity Crewed MTV uses 3 - 25 klbf NTR Engines & PVA Auxiliary Power – Mars (2009) Reusable Lunar Transfer Vehicle using Single 75 klbf Engine -- SEI (1990-91) Artificial Gravity BNTR Crewed Transfer Vehicle also using “Bimodal” NTR Earth Return Vehicle using Clustered Clustered 15 klbf / 25 kWe Engines -- Mars DRM 4.0 (1999) 15 klbf / 25 kWe Engines -- Mars DRM 1.0 (1993) Historic Nuclear Thermal Propulsion Efforts Rover/NERVA* era (1955-1972) • 20 Rocket/reactors designed, built & tested at cost of ~1.4 B$ • Engine sizes tested 25, 50, 75 and 250 klbf • H2 exit temperatures achieved 2,350-2,550 K (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 (NRX-A6 - single burn) > 3.5 hrs (NRX-XE: 28 burns / accumulated burn time) • Engine thrust-to-weight ~3 for 75 klbf NERVA • “Open Air” testing at Nevada Test Site ----------------------------- The NERVA Experimental Engine (XE) demonstrated 28 start-up / shut-down cycles during tests in 1969. * NERVA: Nuclear Engine for Rocket Vehicle Applications Rover/NERVA* era (1955-1972) • 20 Rocket/reactors designed, built & tested at cost of ~1.4 B$ • Engine sizes tested 25, 50, 75 and 250 klbf • H2 exit temperatures achieved 2,350-2,550 K (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 (NRX-A6 - single burn) > 3.5 hrs (NRX-XE: 28 burns / accumulated burn time) • Engine thrust-to-weight ~3 for 75 klbf NERVA • “Open Air” testing at Nevada Test Site ----------------------------- * NERVA: Nuclear Engine for Rocket Vehicle Applications NRX-A6 (1972): Hydrogen exit temperature = 2556 K 1 hr. Rover/NERVA Development Overview Geometry – Extruded Hexagonal . Length: 51 in. Flat-to-Flat: 0.75 in. 19 Coolant Channels . Fabrication: Extrusion and Sinter Fuel Compound . UO2 . UC2 Emerging Fuels . (U,Zr)C fuel web Highest level of development status for any fuel with . (U,Zr)C solid solution significant remaining challenges Ca rblde/Graohl tc Composite UC 2 Parll<;les/Graphltc M~trhc PyC Coated uc2 Snhcrcs/Gr aphtte Matrl~ Carbide Solld Solution NbC coatings • ZrC coatings ... ZrC + Mo coatings Two major failure mechanisms were discovered through engine testing Corrosion Mechanical Failure NRX-A2-A3 ,.o TYPICAL w 0.8 - WANL 1- ..a: z 0 iii 0.6 - 0 a: a: 8 w 0.4 - > .: NRX-A&XE ...., TYPICAL Y-12 w a: 0.2 - A4 AS 1965 1966 1967 1968 1969 COMPARISON OF A-4 &AS CRACK PATTERNS @ STATION 20 200X Composite NbC ZrC Graphite (U,Zr)C-C CTE Graphite Matrix: Post Exposure 7.0 - 7.2 7.6 - 7.7 3.0 6.0 - 6.7 (µm/m·K) 15 Historic U.S. Cermet Fuel Development (1957-1968) ANL Nuclear Thermal Propulsion Ceramic Metallic (cermet) Fuels W-UO2 (net-shape HIP, prismatic) (1962 - 1966) 0-Type Assembly E-Type Anembly Anem11tyefm~ 4tnu111 can comp,nKJts =[ Tllb8sc:oq ntJstd 0nt opku' Tllfles upill'J In klhol·f U ~r tuUrt(yClll 11a t-1upresswrc:,c1e (Pins su~stCjl,l ent)J Inched aul al tubl:sj Nose Piece Reflector GE·NMPO f i t . ).10 - "-"l•f e.,.... .. , RMCIO< !,,ti ooml~. ·--... ,~ t.-i.,.PUl1-258, "'4-ll,2SA. I = Cermets fail from the inside out due to high temperature instability of ceramic fuel particles Before 60 vol% Before UO2 cermets cermets 2 2 dUO - dUO - W MoW After Hot Hydrogen Exposure Pre-Test Post-Test After Hot Hydrogen Exposure -Oxide and nitride ceramic fuels susceptible to H2 corrosion -necessitates the need for fuel cladding Successfully developed cermets prevented free oxygen loss and thermal stresses Mass Loss Mitigation Parameters i i i, 30 - W-alloy Claddings _,_~ f.. 20 Ji - Gd2O3 and ThO2 stabilizers "'ii, :II ... ( l1ungsten tlii!l lalllil'T(. 0. 003 iJl!. - High fuel density to reduce free U, O / 011 1 ·r fas, edgl fillln unelllll f migration 2 • ~ 8 10 12 - Eliminate interparticle connectivity TciJl tes.ttlme; hr 1...,..... • , l6DOlJ:102Gl7JOOl'XIO 2100 XO> l900 - Spherical particles to allow for favorable lO - o.,.._P • SIO (e•)~- - -r-- (R.t 1) interfaces at grain boundaries i--.. .,- 1 - UO → UN fuel particles 2 >-..._..,......,.1P • 3l.2 (e•)~ -r-- - ~ ·-== I= (11.f.1) ! 1.,- ' -..... 1 H,,..._ p • 'lT7 1...,) -10.CW:/ l00 j ~ ..... ~ i"-.. .. -. !§ -:::- I I ~ .,­. --::::r: .. 16 .. •• u •• Rec.,.C..l t....,.-. JO'/ •I( f l9. S-C..,.,l a• ef ~ IN uefflcle1111 fer 11itra,._, .. ,i,,_,.., -4 llll'ffN .__.. INC •CUI lllflf1 '" 18 Solid-solution carbides have high melting temperature and exhibit high temperature stability 5 X IQ--4 .------.----,----~---,------r----.--, 3095 K Ill IQ-5 I\( E ~ "'.. --~- 2980K -·E -w 1- <t 3450 a:: 2880 K (f) NbC 00--c'=---~--c>3-=-o-- 4>--co~___,..50~----,,G-=-o-~1'-=o--sc'-co~___,_9-=-o-_,101§>2,c (f) {3505°) COMPOSITION ZrC {M OLE PERCENT) (3495 °C ) 0 _J (f) f/) \ 40QQ ..-------r--...-----r--,---,------,----,----,--,-----, <t ,o-6 •~ .... --=.- 2770K :? PURE HYDROGEN FLOW • 730 scfh u Ua.os Zro.95 C1.01 0 3500 .....£ .!: 0 0. Cl C 30001-----+-- --1--- ~...i--~~-At-----t Zl ,o-70 '-----2L___ ...;4L-----J6L----Je ___ ..1.,o---~,z~_, i TOTAL PERCENT MASS LOSS (TPML) ZrC rich solid solutions show excellent 20 40 60 BO 100 resistance to hydrogen corrosion mol'/,UC History: USSR RD-0410 Carbide Fuel Development The NTP fuel form development efforts of the former Soviet Union was comparable to United States efforts if not exceeded that of the United States. Geometry – ‘Twisted Ribbon’ . Length: ~100 mm . Diameter: 2 mm . Attempt to maximize heat transfer while maintaining fuel integrity Fuel Compounds . Fuel composition was focused on maximizing the operating temperature of the fuel . Carbide • (U,Zr)C • (U,Zr,Nb)C • (U,Zr,Ta)C .
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