SPACE NUCLEAR REACTORS

Sümer ŞahinFaculty of EngineeringNear East University

Turkish Republic of Northern Cyprus

sumer.sahin@NEU.edu.trhttps://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.

The structure of a thermoelement

Schematic of a silicon germanium alloy (Si0:8Ge0:2) thermoelectric unicouple

TEM in the generator mode

Basic principles of thermionic energy converter

Emitter: Mo, W,

Re

Collector: Nb

Tem=1800–2000 oK

Tcol = 900–1100 oK

SNAP-27 70 Watts

~4 Kg Pu-238

600oC - 275oC PbTe

unicouples

SNAP System for Nuclear Auxiliary Power

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

General purpose heat source radioisotope thermoelectric generator (GPHS-RTG)

MULTI-MISSION RADIOISOTOPE THERMOELECTRIC GENERATOR (MMRTG)

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

Incore

thermionic

element

(inside the colling

channel-German)

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

Nuclear Thermal Rocket (Topping Cycle)

• 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

Thermionic element for a fast space reactor (USA)

Hollow vermicelli carbide fuel configuration

Tri-carbide Foam Fuel Assembly for NTR

• 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)

Temperature field in the reflector by 100 % B4C content

in the strips during power phase

Temperature field in the reflector by 100 % B4C content

in the strips during thrust phase (F=5000N)

Temperature field in the reflector by 20 % B4C content in

the strips during thrust phase (F=5000N)

Temperature field in the reflector by 10 % B4C content in

the strips during thrust phase (F=5000N)

• 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

Nuclear Thermal Rocket (Topping Cycle)

Nuclear Electric Propulsion

• 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

Brayton Power System

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

Geometrical model of the target assembly

Nuclear radiation shielding for supermagnetic coils

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 --

Plataforma Solar de Almería (PSA)

Central Receiver Facility: 7-MW(Th) CESA-I FACILITY

Linear focusing facility: DISS

Canales parabólicos

(Receptor central

(Fresnel reflectors)

(Stirling)

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.

YOU ARE ALL CORDIALLY INVITED

TO ATTEND

ICENES2015, 17th INTERNATIONAL

CONFERENCE ON EMERGING

NUCLEAR ENERGY SYSTEMS

(April 2017)

HEFEI, CHINA