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SPACE NUCLEAR REACTORS
Sümer ŞahinFaculty of EngineeringNear East University
Turkish Republic of Northern Cyprus
[email protected]://neu.edu.tr/wp-
content/uploads/2016/08/cveng2.pdf
• Fuel energy densities ~ 107 that of chemical systemsEnables or significantly enhances:• Space Power and Propulsion
– Power and propulsion independent of proximity to sun or solar illumination• Constant power level available for thrusting and
braking– Go where you want, when you want
• Expanded launch windows• Enhanced maneuverability• Faster trip times / reduced human radiation dose
– Enables planetary global access– Enables Lunar overnight stays
Why Nuclear Power in Space?
Nuclear fusion fuels:
2H1 (D); 3H1 (T); 3He2Natural fuels: D (isotopic fraction in natural water: 150
ppm) and 3He2 (isotopic fraction in natural helium: 1.38
ppm). Abundant 3He2 on the Moon (109 kg), in the
Jupiter (1022 kg), Saturn (1022 kg), Uranus (1020 kg) and
Neptune (1020 kg) atmosphere. Fusion energy
availability for 100’s of millions years!!!
Tritium is an artificial radioactive element!!!:3H1
3He2 + 0ß-1 (T½ = 12.323 a)
“T” production:6Li3 + 1n0
3H1 (T) + 4He2 + 4.784 MeV7Li3 + 1n0
3H1 (T) + 4He2 + 1n0` + 2:467 MeV
• Radioisotope Power Generator (RTG)
– 100’s We – a few kWe
• Nuclear Thermal Propulsion (NTP)
– Burn Time: 1-10 hours
– Thrust: 25-125 kN
– Specific impulse: 800-1300 s (Hydrogen, solid fuel)
– Max Temperature: 3000±300 K
• Nuclear Electric Propulsion (NEP)
– Lifetime: 1-10 Years
– Power level & Thrust: kW-MMW, N -MkN
– Specific impulse: 1,500-10,000 s
– Max Core Temperature: 2,000±500 K
Nuclear Power in Space
Energy conversion system for space craftsmust be of statical nature! They must not have moving part in order to avoid
kinematic instabilities of the space craftsas well as to eliminate the need for maintenance!
For fission reactors, most prospective onesare:
• Thermoelectric reactors.• Thermionic reactors.
Basic principles of thermionic energy converter
Emitter: Mo, W,
Re
Collector: Nb
Tem=1800–2000 oK
Tcol = 900–1100 oK
Radioisotope Power Systems
5.6 MeV
Pu-238
U-234
(He-4)
• Portion of heat energy (~ 6%) converted to
electricity via passive or dynamic processes
• Thermoelectric (existing)
• Stirling (under development)
• Brayton (future candidate)
• Waste heat can be used for thermal control
Heat Source Assembly
(18 GPHS Modules)
Radioisotope Thermoelectric Generator
(RTG)
Thermoelectric
Converter
Radiator Assembly
Pu-238
Thermal
Source
Power
Conversion
Radiators
Electrical
Power
Waste
Heat
Low
Temp
TOPAZ space reactor power systemLength: 4.7 m; Major diameter: 1.3 m; NaK-78 coolant ~970 K 79 TFEs with five thermionic cells each, later single cell used
YENISEI reactor employed a single monolithic block of ZrH moderator,37 single-cell TFEs. 34 were dedicated to the electrical load and three providing high-current, low-voltage DC power to the electromagnetic
induction pump.
A layout of the Heat pipe cooled reactor with segmented
thermoelectric module converters space reactor power
system for a nominal power of 110 kWe
Basic structure of a nuclear thermionic element with auxiliary emitter
Typical auxiliary emitter, suggested by Rasor Associates
Cooling channels of the thermionic fuel elements are imbedded in ZrH1.7 moderator with variable (increasing) mesh
width for power flattening
• Performance is measured in terms of specific impulse and thrust to weight ratio
– Isp ~ (T/M).5 ; lowest mass propellant at highest temperature
– High F/W ratio means very high specific power
• Design uniqueness
– Hydrogen propellant
– Ultrahigh temperature
• Fuel temperature ~ 3000 K
• Chief challenges
– Fuel materials & design
– Structural materials
– Design methodology and testing
Nuclear Thermal Rockets
Functional view
of the hybrid
reactor
Nuclear thermal
thrust
F = ~5000 N
Specific impulse
~ 670 sec-1
Hydrogen exit
temperature
~1900 oK
• The foam fuel can operate at extremely high temperatures usingrefractory materials with low neutron absorption cross-section andextended surface area. The foam fuel material consists of atricarbide, UZrNbC, made of enriched uranium that is vaporinfiltrated into an open-cell foam matrix of NbC and ZrC. It is theequivalent of a solid eutectic solution.
• The base matrix is created by starting with a polyurethane foamwith suitable porosity and pore density. The polymer is thenpyrolyzed above 2400 oC to produce a carbonaceous foamskeleton. The carbon foam is then heated, and carefully selectedgas mixtures are flowed through the foam enabling the chemicalvapor to actually infiltrate (CVI) the graphite webbing and react toform UNbZrC.
• The coating method consists of decomposition of a gaseousprecursor (usually a metal chloride or fluoride), flowed over orthrough a heated substrate, and subsequent condensation from thevapor state to form a solid deposit
• Natural B4C neutron absorbers
(Radial reflector = thickness 16 cm, drum diameter = 13.5 cm, strip width = 5 mm).
10B isotope (20 % in natural boron) producessignificant nuclear heat via neutron irradiation in the reflector
• 10B + n 7Li + 4He + Q (2.79 MeV)
• 100 % natural B4C; keff,max = 10.7 %; Tdrum = 1020 oK
• 20 % natural B4C; keff,max = 8.4 %; Tdrum = 660 oK
• 10 % natural B4C; keff,max = 7.7 %; Tdrum = 520 oK
• Performance is measured in terms of specific impulse and thrust to weight ratio– Isp ~ (T/M).5 ; lowest mass propellant at highest temperature
– High F/W ratio means very high specific power
• Design uniqueness– Hydrogen propellant
– Ultrahigh temperature
• Fuel temperature ~ 3000 K
• Chief challenges– Fuel materials & design
– Structural materials
– Design methodology and testing
Nuclear Thermal Rockets
• Compact system capable of providing spacecraft propulsion and electrical power for deep space robotic missions or near-Earth cargo and piloted Mars missions.
• Primary subsystems include: reactor system, power conversion unit(s), power management and distribution unit, heat rejection system, and electric thrusters.
• Characterized by extended operation and minimum propellant mass.
Nuclear Electric Propulsion
Thrusters
Reactor SystemPCU
ShieldReactor
Main Radiator
PCU
PCU
HX
Nuclear Electric Propulsion
VISTA – A Vehicle for Interplanetary Space Transport
Application Powered by Inertial Confinement Fusion
Artificial gravity at the peripheral zones of the VISTA space craft as a function of vehicle rotation
Application of solar energy
• High temperature conversion with solarcollectors, (parabolic Fresnel mirrors).
• spacecraft application with thermo-electricor thermionic converters requires hardtechnology!!!!!)
Diameter of the mirror (m) 20,93
Mass of the thermionic system (kg) 37,3
Mass of the mirror (kg) 92,88
Total mass with an excess of 50 % 180
Mass-to-power ratio (kg/kW(el)) 3,6
Conversion efficiency (%) 12
Basic technical values of a 50 kW(el) solar energy thermionic generator on earth orbit
Planets Distance
from the
sun
[AU]
Diameter
of the
mirror
(m)
Total
mass
(kg)
Mass-to-
power ratio
(kg/kW(el))
Concentration
factor
Mercury 0,39 8,16 62 1,34 185
Venus 0,72 15,07 113 2,26 632
Earth 1,00 20,93 180 3,6 1219
Mars 1,52 31,81 363 7,26 2816
Basic technical values of a 50 kW(el) solar energy thermionic generator on the orbit of
different planets
Emitter temperature 1500 K
Collector temperature 900K
Electrode spacing 1,3 mm
Specific power at the emitter surface 10 Wel/cm2
Power per converter 1200 Wel
Gross conversion efficiency 30 %
Part of the waste heat through emitter back face 82 %
Part of the waste heat through radiator 18 %
Mirror concentration ratio 380
Electrical power output 50 kW 10 GW
Diameter of the mirror 12,8m 5,7 km
Surface of the mirror 130 m2 25,5 km2
Total mass (TI-system + mirror + 50% for the support) 100 kg 46800 t
Mass-to-power ratio (kg/kWel) 2 4,68
Main technical data of solar energy generators with advanced thermionic converters in Earth orbit
Velocity requirments for divers
applications
Velocity
(m/sec)
Typical location
High Earth Orbit 4000 55’000 km
Lunar Orbit 4250 25’000 km
Solar Orbit 4450 0,85 AU
Lunar Surface soft landing 6050 Backside surface
Solar System Escape 8750 --
Injection into the sun 24000 --
Availability of solar energy in space issignificantly higher compared toterrestrial solar energy!
Neither seasonal nor daily shortage!!!
Application of statical conversiontechniques is most suitable.
CONCLUSIONS
• Great potential of solar thermalelectricity via nuclear and solar energyconversion for both thermoelectric aswell as thermionic with modularconverters for low to very high powerneeds warrants research efforts andinvestments on that line forintermediate future and space industry.
FINAL CONCLUSIONS
• Nuclear power is a viable and available option forMKw to MMw space power applications
• Current space missions are powered by RTG andnon-nuclear power systems with technologiesdeveloped 30 – 50 years ago
• If there is a compelling need, technology for spacenuclear power and propulsion is in significanlymore advanced stage than presently applied!!!
• International collaboration of governments andprivate enterprises are key to realization of futurespace flights in industrial level.