EXTROVERT
Space Propulsion 14
Electric Propulsion Continued
EXTROVERT
Space Propulsion 14
Nuclear Rocket Engines
EXTROVERT
Space Propulsion 14
Nuclear Rocket Engines
•Nuclear Thermal Rockets : Propellant gets heated by conduction/
convection from fuel.
•Nuclear Electric Propulsion: Electric power generated by
heat engine or thermo-electric effects is used to drive electric
Propulsion system.
•Specific Impulse more than twice that of chemical - AND
•Large-thrust possible because of high thrust/weight ratio.
Shorter mission times:
Proponents point out that for long-duration space missions, this leads to
LOWER total radiation exposure per mission. (Less time exposed to the Big
Nuclear Furnaces in the Sky)
EXTROVERT
Space Propulsion 14
Nuclear Thermal Rockets (NTR)
Originally researched by the U.S. Air Force and NASA from the late 1940s to
the 1960s, the use of nuclear energy to power a rocket engine has many
advantages:
•High thrust (as much as chemical rockets – temperature and pressure limits
are similar)
•High Specific Impulse
•Multiple re-starts
Disadvantages:
•Heavy (reactor and shielding )
•Radiation concerns
•Testing / development / political issues
“Thermal” reactor: neutrons slowed down below 1eV, moderator of
light elements used.
“Fast”: broad spectrum of neutron energy up to 15eV. No moderator.
EXTROVERT
Space Propulsion 14
NTR Concept
Neutron Reflector: typically Beryllium shield
Propellant: H2 (high Ue), CH4 (better storage density) etc.
EXTROVERT
Space Propulsion 14
Types of Reactors
1.
NERVA (Nuclear Engine for Rocket Vehicle Applications )
Graphite fuel rods with uranium-carbide fuel particles – coated to protect
from hydrogen.
Coolant passes through channels in the rods.
Most fully developed, but low T/W
EXTROVERT
Space Propulsion 14
NERVA
EXTROVERT
Particle Bed Reactor (PBR)
Space Propulsion 14
Particles of uranium-carbide fuel (coated) are packed between two porous
cylinders. Hydrogen (or helium) is directly used to cool them.
PBRs have a higher fuel density and thus higher T/W. The configuration also allows a
higher temperature in the working fluid before the fuel melts – Thus higher specific
impulse.
inisjp.tokai.jaeri.go.jp/
ACT95E/11/1104.htm
EXTROVERT
Space Propulsion 14
“The technical risks of the PBR include:
+ Challenges in fabricating the high temperature fuel particles that
are the key to this technology -- efforts to date have failed to
conclusively demonstrate that fuel particles can withstand the rigors
of the reactor operating environment;
+ The low thermal capacity of the reactor core increases the risk of
thermal damage to the core in off-normal conditions, or during
reactor cool-down;”
lifesci3.arc.nasa.gov/.../thomas/
Adv.prop/advprop.html
EXTROVERT
Space Propulsion 14
CERMET-core
Source: www.ms.ornl.gov/.../ SCG/programs/OTT_HVPSM.html
•“Cermets: composite materials made by combining a ceramic and a metallic alloy.”
•“Cermets are made by 1) milling together powders of the desired materials, and 2) forming a shape,
and then 3) sintering (heating) those materials together to form dense parts. The sintering process
causes the component materials to bond together to form a new material with structural characteristics
that are better suited for the intended use than either of the starting materials.”
•CERMET fuel: Hexagonal fuel elements. Uranium oxide particles imbedded in
tungsten / tungsten-rhenium matrix; uranium oxide in molybdenum.
• Advantage: very long life ( > 40 hours).
• Uses fast fissioning reactor that does not depend as much on a “moderator”
material to slow down the neutrons in the fission reaction.
•A newer system that still requires development
•Has potential for multiple restarts (very robust fuel elements)
EXTROVERT
Space Propulsion 14
“ 6 - Cermet Reactor ( http://lifesci3.arc.nasa.gov/.../thomas/
Adv.prop/advprop.html )
General Electric (GE) is the leading proponent for the Cermet
Reactor, which has been evaluated for SDI(16) and SEI
applications.(17) .. advantages :
+ .. improved thermal conductivity compared to metal oxide fuel
elements.
+ .. extensive fuel test engineering heritage exists from the ANP and
710 programs.
+ … high retention of fission products in the fuel matrix. ..
experimentally demonstrated in the 710 Program, in which most test
fuel elements demonstrated fission gas fraction release of less than
10-9, while some fuel elements released fission fragment fractions in
the range of 10-5 to 10-4. This should produce less stringent
containment and confinement restrictions on Ground Test Facilities.
+ .. cermet fuel may offer improved swelling behavior, but this
remains uncertain.
EXTROVERT
Space Propulsion 14
Several issues remain open:
- .. lower fuel density relative to metal fuel elements, a potential
disadvantage of cermet fuels is large core size, and thus greater core
and shielding mass.
- In order to improve weldability, Rhenium is a potential Cermet
cladding material. “
EXTROVERT
Space Propulsion 14
Reactor Comparison: Table 8.2, Humble
Nerva
Particle-Bed
CERMET
Power (MW)
1570
1945
2000
Thrust (N)
334,061
333,617
445,267
Propellant
H2
H2
H2
Fuel element
Solid rod
Porous particle
bed
Solid rod
Max temp (K)
2361
3200
2507
Isp (s)
825
971
930
Chamber
pressure (Mpa)
3.102
6.893
4.136
Nozzle expansion
ratio
100
125
120
Engine mass(Kg)
10138
1705
9091
Total shield
mass(Kg)
1590
1590
1590
EXTROVERT
Space Propulsion 14
Candidate Working Fluids
Working Fluids: coolant, exhaust gases
As the propellant, we want something that can be easily heated and results in a
high Isp.
Low molecular weight is preferred, but propellants with higher molecular weight
may be used depending on storage volume constraints.
Hydrogen: MW = 2.016
Methane:
16.043
Carbon dioxide 44.01
Water:
18.015
Note: the above offer the possibility of replenishing propellant from extraterrestrial sites.
EXTROVERT
Space Propulsion 14
For a given temperature limit, we can determine the specific heats
from equations or tables and find the other quantities
such as :
cp
 , c *,CF ,Isp
cp
cp
 


cv c p  R c  Runiv
p
Mol .Wt
Runiv  8314J /(kmol  K )
EXTROVERT
Space Propulsion 14
NTR Performance Example
Assume a PBR NTR. Hydrogen is heated to 3000K at a total pressure of 60 atm
The engine operates in vacuum with e = 200, At = 100 cm2, hCF = 0.97, hC* = 1.0.
Find vacuum specific impulse, thrust and mass flow rate of hydrogen.
At 3000K,
1 
 T 
cp 
56.05

702.74

100 
2.016 



0.75 
kJ
 1165


18.048

kgK
  T 
100 
EXTROVERT
Space Propulsion 14
This gives
  1.289
c* 
c *  RTc
 1
  1
 2
 



1


 5287m / s
EXTROVERT
Space Propulsion 14
Pe
 0.000172
P0
At = 200,  =1.289,
cFideal
cFideal
 1
 2 1
 2
 

   1
 1.874
 2 
   1


 1
 Pe  
1
 
 Pc 
 Pe 
 c  
 Pc 
EXTROVERT
Space Propulsion 14
cF  cF cFideal  1.8178
Ispv
c * cF

 979.7 sec
g0
Tvac  110,520N
mp  11.5kg / s
Not counting any turbine/pump losses (I.e., closed cycle)
EXTROVERT
Space Propulsion 14
Reactor Power
The reactor provides the power necessary to heat the propellant to the required
temperature.
Pcore
hv
Psp
60
T2




 mp hv   c pdT   mpPsp


T1
Heat of vaporization
Specific power required. Depends on propellant initial / final
temperatures
H2
is 48 MW/(kg/s)
For H2 at 3000K, P
sp
CH4
Psp
CO2
0
2000
T
3500
Pcore  552MW
EXTROVERT
Space Propulsion 14
Reactor Mass
From Fig. 8.21 in Humble, (Reactor Mass vs. Reactor Power for different
NTR technologies ) the mass of the reactor is estimated as 500 kg.
EXTROVERT
Space Propulsion 14
Reactor Shielding
We also typically provide a “shadow shield” between the reactor and the
payload for in-space applications. A typical shield consists of layers of
Lithium hydride (LiH2) which is a neutron absorber, Tungsten to shield
against gamma rays, Beryllium as a neutron reflector. A typical shield may
be 25cm thick and have a mass-density of 3500 kg/m2 of surface area
Total mass = Reactor + shield + nozzle + turbopumps + core containment vessel
Typical thrust-to-weight ratio of an NTR is from 3 to 10.
EXTROVERT
Preliminary Design Decisions
Space Propulsion 14
NERVA type core limited to ~ 2360 K temperature
CERMET: 2500K
PBR: 3200K
Determine gas properties
Determine Nozzle expansion ratio and Isp.
Sizing the System
Inert-Mass Fraction (when using hydrogen propellant): 0.5 to 0.7
Lack of database of nuclear engines prevents good estimate of inert mass
fraction: iterate.
From Isp and required thrust, find propellant mass flow rate.
Determine required reactor power to heat propellant to required temperature
at the required flow rate.
EXTROVERT
Space Propulsion 14
Determine system pressure levels: empirical correlations
NERVA-1
Enhanced Particle
NERVA
Bed
CERMET
Chamber
Pressure
3.1 MPa
6.9MPa
6.2MPa
4.1MPa
Core
pressure
drop: %
of
chamber
pressure
38%
10.6%
5%
53.7%
% of
reactor
head
pressure
27.5%
9.6%
4.8%
34.9%
EXTROVERT
Space Propulsion 14
Prometheus Nuclear Engine: JIMO mission
Jupiter Icy Moons Orbiter: Nuclear electric primary propulsion
Two basic types of technology under consideration:
(1) radioisotope-based systems
(2) nuclear fission-based systems.
EXTROVERT
Space Propulsion 14
RadioIsotope Thermoelectric Generators
Heat from plutonium dioxide and solid-state thermocouples to
convert directly to electricity. Cold outer space is the cold junction.
General Purpose Heat Generator: 250 watts
Multimission Radioisotope Thermoelectric Generator: 100watts, 14+ years
http://spacescience.nasa.gov/missions/MMRTG.pdf
EXTROVERT
Space Propulsion 14
Stirling Radioisotope Generator
http://spacescience.nasa.gov/missions/Stirling.pdf
Closed cycle Stirling cycle free-piston machine. Heat from GPHG with 600g
Plutonium dioxide at 650C. Heat rejected from other end at 80C. Closed-cycle
Engine converts heat to reciprocating motion – linear alternator produces
62-65 watts AC – converts to 55w DC.