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- 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 . Carbonitride • (U,Zr)C,N Reported ternary-carbide fuel performance of operation at 3100 K for up to 1 hr. Carbide fuels also experience the most failure in the mid-band region where power densities are high and ductility low Highest Rate of Brittle Failure Failure Ductile Failure 3000 1.0 T, K 2500 2000 0.6 q 0.4 1000 0.2 T 500 100 200 300 400 500 L , mm Fueled Region 21 History: US Carbide Fuel Development NERVA/Rover Carbide Fuels (1955 – 1972) Hot Hydrogen Hot Hydrogen Prototypic Prototypic Static Testing Thermal Cycling Irradiation Testing Engine Testing ✓ ✓ ✓ Aluminum Tube Carbide (U,Zr)C fuels were tested in NF-1 and survived exposure to over 2700 K for 109 minutes under flowing hydrogen and irradiation. Fuel element architecture that the Rover/NERVA program used for all their fuel compositions Overview of NTP Fuel Timeline USSR Solid-Solution Carbide Fuels (1955 – 1993) Solid Solution Carbides Solid Solution Carbonitrides Carbonitride Cermets 1950 1960 1970 1980 1990 2000 2010 2020 Graphite Matrix Cermet, Composite Coated Carbide, Solid Solution Carbides Particle Emerging Fuels NERVA/Rover SNTP AES NTP, GCD NTP (1955 – 1972) (1987 – 1993) (2010 – 2019) MoMo-U0-UO2 hothot-rolled-rolled & sinteredintered W--U0UO22,, WW-UN-UN sintered Cermet Development ANL, NASA LeRC, GE 710 (1962 – 1968) Nuclear Materials Group – NASA MSFC • Marvin Barnes • Omar Mireles, Ph.D • Omar Rodriguez, Ph.D • Jhonathan Rosales, Ph.D • Brian Taylor • Martin Volz, Ph.D • Ryan Wilkerson, Ph.D Thank you!