Transcript Document

A Comparison of Nuclear Thermal to
Nuclear Electric Propulsion for
Interplanetary Missions
Mike Osenar
Mentor: LtCol Lawrence
Overview
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

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Introduction
Objective
Establish parameters
NTR Design
NEP Design
Discussion and Conclusion
Introduction
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NASA is developing Nuclear Electric
Propulsion (NEP) systems for Project
Prometheus, a series of interplanetary
missions
What happened to Nuclear Thermal Rocket
(NTR) systems? Should NASA only invest in
NEP systems?
Objectives
Prove the feasibility of different nuclear
propulsion systems for interplanetary
missions which fit in a single launch vehicle
 Compare NTR and NEP system designs for
given missions
Method: take a set of inputs, use a series of
calculations and SPAD process along with
reasonable design assumptions to design a
spacecraft to reach a given ΔV
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Establish Parameters
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Establish ΔV’s and flight times for both NEP
and NTR systems to Jupiter and Pluto
Determine launch vehicle payload
restrictions
Obtain design points – inert mass fractions
based on thruster specific impulses
Establish Parameters
NTR ΔV
(km/sec)
NEP ΔV
(km/sec)
NTR TOF
(years)
NEP TOF
(years)
Jupiter
3.83
7.66
4.13
4.13
Pluto
6.70
13.40
19.00
19.00
•NEP ΔV’s and flight times based on AIAA 2002-4729 –
low thrust gravity assist trajectories
•NTR data derived from NEP data
Establish Parameters
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Relationship between NEP ΔV/TOF and NTR
ΔV/TOF
Table shows that NTR has same TOF for 50% of the
ΔV
NTR numbers based on AIAA 1992-3778
Mission
ΔV (km/s)
TOF (yrs)
Pluto NEP
13.4
19
Pluto NTR
6.52
16
Pluto NTR
12.9
10
Establish Parameters
Ariane 5 Payload
Specifications
Mass to orbit
(kg)
18000
Height (m)
12.5
Diameter (m)
4.5
Establish Parameters
Dumbkopff Chart - Jupiter NTR 1000 kg
Dumbkopff Chart - Jupiter NEP 1000 kg
0.1
0.2
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0
0.3
0.4
0.5
0.6
0.7
0.8
Initial Mass (kg)
18000
Initial Mass (kg)
18000
0
0.9
0
500
1000
1500
0.9
0
1000
2000
Isp (sec)
3000
4000
5000
Isp (sec)
Dumbkopff Chart - Pluto NTR 500 kg
Dumbkopff Chart - Pluto NEP 500 kg
0.1
0.2
18000
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0
0.3
0.4
0.5
0.6
0.7
0.8
0.9
0
500
1000
Isp (sec)
1500
Initial Mass (kg)
Initial Mass (kg)
18000
0
0.9
0
1000
2000
3000
Isp (sec)
4000
5000
Establish Parameters
Design points established from Dumbkopff charts
Design Isp (sec)
ΔV (km/sec)
f-inert
Jupiter NTR
1000
3.83
0.65
Jupiter NEP (Ion)
3500
7.66
0.80
Jupiter NEP (Hall)
1500
7.66
0.60
Pluto NTR
1000
6.70
0.50
Pluto NEP (Ion)
3500
13.40
0.65
Pluto NEP (Hall)
1500
13.40
0.32
NTR Design
Size system so that it meets 3 specifications
1. Under max payload mass
2. Fits in payload fairing
3. Reaches required ΔV
NTR Design
Inputs from Dumbkopff: finert, ΔV
Assumptions
Po = 7 MPa
Isp = 1000 s – hydrogen
Tc = 3200 K
T/W = .3 – experimented, balance between
high thrust short burn time and low reactor
mass (low power)
NTR Design
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Equations for basic parameters
m prop
  V  
 I g  
m pay  e  sp 0   11  f inert 







1  f inert e

FF
W
mi
 V 
 I sp g 0 



F
m 
I sp g 0
 0.018061T  5.715417
Pm
NTR Design
Subsystem Sizing (note: volume constraint height)
Payload
1000 kg to Jupiter, 500 to Pluto
based on densities of actual space mission
sized as 2 m tall cylinder
mtank
Tank
biggest part – hydrogen has low density
pbVtot

g 0ta.nk
NTR Design
Turbo Pump Feed System
Nuclear Reactor
5
4
3
Rcore  2.655(10) 12 Pcore
 8.946(10) 9 Pcore
 1.1703(10) 5 Pcore
2
 7.427(10) 3 Pcore
 2.2955Pcore  313.34
2
H core  4.027(10) 5 Pcore
 0.1427Pcore  17.9883
Radiation Shield
standard SPAD design – 18 cm Be, 5 cm W,
5 cm LiH2
NTR Design
Nozzle
Columbium, designed to be ideally expanded
in space (ε=100)
Miscellaneous
Avionics
Reactor containment vessel
Attitude thrusters
Structural mass
NTR Design
Achievable ΔV verified
with Rocket Equation
 mi
V  I sp g 0 ln
m
 f




Payload
Propellant
Tank
Pump
Shield
Vehicle height determined
by stacking parts
according to Figure
Reactor
Nozzle
NTR Design
Final Results of NTR Design
ΔV
(km/s)
Jupiter
NTR
Pluto
NTR
f-inert
Initial
Mass (kg)
Height
(m)
Power
(MWe)
TOF
(years)
4.191
0.6094
9100.41
7.23
281.23
4.13
8.103
0.4182
14853.83
12.29
281.23
19.00
NEP Design
Size system so that it meets 2 specifications
1. Under max payload mass
2. Reaches required ΔV
No size requirement – analysis showed that
NEP systems would violate mass
constraints before volume – no low-density
hydrogen propellant
NEP Design
Power Source
 Nuclear Reactors (P>6 kWe)
–
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Critical reactors designed as small as 6 kWe
Radioisotope Thermoelectric Generators
(RTG) (P<6 kWe)
Solar?
NEP Design
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Solar Power proportional to inverse square of
distance from sun
to receive power equal to 1 m2 solar panel in
earth orbit, would need 27 m2 panel at
Jupiter and 1562 m2 panel at Pluto
does not factor in degradation – significant
for long lifetimes
engineering, GNC concerns with huge solar
array
mass too much
NEP Design
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Thrusters based on actual designed thrusters
from SPAD
Baselines used: T6, XIPS-25, RIT-XT
Design allowed thrusters to be clustered in
groups of up to 3 – proven to work, increases
force and power appropriately
NEP Design
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Use NTR equations for propellant mass,
thrust, mass flow and power
NEP equations:
Ve  I sp g 0
1 m V 2
e
P 2

NEP Design
Subsystem Design
 Power system
 Propellant tank
 Thruster mass
 Power conditioning mass
 Other mass (structural, feed systems,
avionics, etc.)
NEP Design
NEP Design Results
ΔV
(km/s)
Jupiter
(Kaufman)
f-inert
Initial Mass
(kg)
TOF
(years)
Power
(kWe)
# of
thrusters
15.860
0.5266
4068.58
4.13
10.258
2
Jupiter (MESC)
14.051
0.5685
3673.06
4.13
8.425
2
Jupiter (RIT)
15.433
0.5622
3768.34
4.13
9.555
2
Jupiter (Hall)
12.242
0.3351
6645.87
4.18
6.180
3
Pluto
(Kaufman)
42.725
0.2656
9495.62
18.79
10.258
2
Pluto (MESC)
41.420
0.2849
8079.27
19.40
8.425
2
Pluto (RIT)
44.626
0.2826
8352.61
19.19
9.555
2
Pluto (Hall)
13.771
0.3433
6719
19.02
1.471
1
Discussion and Conclusion
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Overall, ΔV’s were low – real science mission
would need higher ΔV to capture orbit of
planet, maneuver
Accurate data on EP trajectories was desired
over ΔV’s for realistic missions
Discussion and Conclusion
NTR Design
 Almost failed Pluto design – tank volume
 High thrust, impulsive burn more reliable –
operates for short time
 Much less efficient then NEP
 Other applications? launch vehicle, human
Mars exploration
Discussion and Conclusion
NEP Design
 Low thrust, long trip times
 Lifetime analysis – electric thrusters tested to
3.5 years – less than Jupiter TOF
 Space Nuclear reactors require extensive
testing
Discussion and Conclusion
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Testing – extensive testing needed for either
system – facilities, money needed to test for
operational lifetime
Safety – perennial concern with nuclear
systems, real hazards to be considered
Radiological hazard – higher with NEP (low
power but long burn time), must be
addressed for either system
Discussion and Conclusion
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NASA probably right to go with NEP for
interplanetary missions
Much stands between now and operational
nuclear propulsion system
Much to be gained from nuclear propulsion
technology
Discussion and Conclusion
Questions?