GOING TO MARS WITH
NUCLEAR THERMAL PROPULSION
Daniel Robert Kirk
Assistant Professor
Mechanical and Aerospace Engineering Department
Florida Institute of Technology
October 22, 2004
Department of Physics and Space Sciences Colloquium
1
GOING TO MARS WITH
NUCLEAR THERMAL ROCKET PROPULSION
• Comments from President Bush (January 2004)
– “Our third goal is to return to the Moon by 2020, as the launching point for
missions beyond.”
– “With the experience and knowledge gained on the Moon, we will then be
ready to take next steps of space exploration: Human missions to Mars and to
worlds beyond.”
• A human mission to Mars implies need to move large payloads as rapidly as
possible, in an efficient and cost-effective manner
• Renewed interest in break-through deep space science/exploration missions
– Project Prometheus and Jupiter Icy Moon Orbiter (JIMO)
2
OVERVIEW
• Rocket Overview
– Categorization of various types of Rockets
– Rocket Mission Selection Guide
– Rocket Performance Parameters
• Nuclear Thermal Propulsion
– Historical Overview
– Hot Hydrogen Properties
– Fluid Mechanic and Heat Transfer Modeling
– Simulation Results
• Future Work – What can we do at FIT?
– How to Simulate Nuclear Reactors for Space Applications
– New Experimental Facility
– Analytical and Computational Efforts
3
WHY ROCKET PROPULSION?
• Rockets provide means to:
– Insert payloads into space (satellites, experiments, defense applications, etc.)
– Space exploration (Atmospheric, solar system)
– Precise, continuous or pulsed, momentum change (station keeping)
– Weapons (wide range of missiles: cruise 1st stage boost, ICBM)
– Rapid change in momentum devices (retro-rockets, JATO, car air bags)
• Rockets vs. other propulsion devices
– Advantages: orbital insertion, deep space travel, etc.
– Disadvantages: carry all propellant, small payload fraction (STS ~ 0.01)
– Area of vigorous research and development
• Rocket propulsion is an exact, but not a fundamental subject
– No basic scientific laws of nature peculiar to propulsion
4
ROCKET CLASSIFICATION
• Rockets may be classified in many ways
– Depending on energy source (chemical, electrical, nuclear, etc.)
– Depending on gas acceleration mechanism / force on vehicle mechanism
– Basic function (booster, sustainer, station keeping)
– Type of vehicle (missile, aircraft, spaceship)
– Based on performance measures (T, T/W, Isp, h) and/or propellant types
– By number of stages
• Primary distinction between Chemical (thermal) and Electrical systems
– Only types of rockets in operation today
– However, a future human mission to Mars will likely utilize a NEW version
of an OLD concept: Nuclear Thermal Propulsion
5
CHEMICAL (THERMAL) ROCKETS: ENERGY LIMITED
6
Chemical
Energy
HOW A CHEMICAL ROCKET WORKS
F
Rocket Propulsion (class of jet propulsion)
Thermal
Energy
Kinetic
Energy
F  m eVe  Pe  Pa Ae
F  m eVe
that produces thrust by ejecting stored matter
• Propellants are combined in a combustion
chamber where chemically react to form high
T&P gases
• Gases accelerated and ejected at high velocity
through nozzle, imparting momentum to
engine
• Thrust force of rocket motor is reaction
experienced by structure
• Same phenomenon which pushes a garden
hose backward as water flows from nozzle
QUESTION:
Could a jet or rocket engine exert thrust while
discharging into a vacuum (with not
atmosphere to “push against”)?
7
ELECTRIC ROCKETS: POWER LIMITED
8
NUCLEAR PROPULSION: PROJECT ORION
•
•
1955 (classified paper) release atomic bombs behind a spacecraft
Bombs would explode, creating a hot plasma, which would them
push against a spacecraft pusher plate, propelling it forward
• Interstellar version: called for a 40-million-ton spacecraft to be
powered by the sequential release of ten million bombs, each
designed to explode roughly 60 m to vehicle's rear
• Nuclear test-ban treaty: explosion of nuclear devices illegal
“This is the first time in modern history that a major expansion of human
technology has been suppressed for political reasons."
9
ADVANCED PROPULSION TECHNOLOGIES
•
•
•
•
•
•
•
•
•
•
•
Solar sailing is a method of converting light energy from the
sun into a source of propulsion for spacecraft
Obtain propulsive power directly from Sun
No Engine → No Need to Carry Fuel
Photons are reflected off giant, mirror-like sails made of thin,
lightweight, reflective material
Continuous pressure exerted by photons provide thrust
Very high Isp
Open up new regions of solar system for exploration, with no
environmental impact on Earth
Leading candidate for missions that require spacecraft to hold
position in space, rather than orbit Earth or Sun
May also extend duration of other missions
Light Sails Do NOT harvest the solar wind for their
propulsion (solar wind < 0.1% due to that of light pressure)
Do not convert to electricity like solar cells
10
FUTURE OF SPACE PROPULSION (?)
11
MOMENTUM EXCHANGE TETHERS
• Momentum-eXchange/Electrodynamic-Reboost (MXER) tether is combination
of technologies designed to help propel satellites and spacecraft
• Long, strong cable rotating in an elliptical orbit around Earth
• Like a catapult, one end of tether catches payloads in LEO, accelerates them to
higher velocities, and then throws them into higher-energy orbits
• Momentum to payload restored using ED forces to push against Earth’s B field
• Solar power drives ionospheric current, tether reboost without using propellant
12
WHAT IS ENERGY LIMIT?
“The basic secret of space travel and extending human presence throughout
heliocentric space is energy, immense quantities of energy”
13
ANTIMATTER PROPULSION
= 1/10th gram
antimatter
• Propulsion by annihilation of matter and
antimatter is under investigation
• Mixture of matter/antimatter provides
highest energy density of any propellant
• Most efficient chemical reactions produce
about 1 x 107 J/kg, nuclear fission 8 x
1013 J/kg, and nuclear fusion 3 x 1014
J/kg, complete annihilation of matter and
antimatter, E = mc2, yields 9 x 1016 J/kg
• Matter-antimatter annihilation releases
about ten billion times more energy than
H2/LOX mixture that powers SSME and
300 times more than fusion reactions at
Sun's core
• Antimatter must be manufactured
• 1 gram of antimatter ~ $62.5 trillion
• Isp ~ 10,000,000 s
• Mars in 2 hours
14
EVERYTHING YOU NEED TO KNOW TO BE A ROCKET SCIENTIST
1. Thrust, [N]
–
–
–
“How much Force?”
T/W is key metric for launch vehicles
Less important with space exploration applications
2. Specific Impulse, [sec]
–
–
–
–
–
“How Efficient?”
High thrust (chemical) have low specific impulse
High specific impulse (electric) rockets usually have low thrust
Increasingly important for space exploration applications
Increases with increasing temperature and decreasing molecular weight
3. Ideal Rocket Equation, [m/s] (1 form of many)
 M initial 

V  Ve ln 
M

 final 
–
“When and How Fast?
Can a rocket travel to a speed faster than speed at which exhaust leaves rocket?
15
PERFORMANCE COMPARISON
16
OUTLINE
• Rocket Overview
– Categorization of various types of Rockets
– Rocket Mission Selection Guide
– Rocket Performance Parameters
• Nuclear Thermal Propulsion
– Historical Overview
– Hot Hydrogen Properties
– Fluid Mechanic and Heat Transfer Modeling
– Simulation Results
• Future Work – What can we do at FIT?
– How to Simulate Nuclear Reactors for Space Applications
– New Experimental Facility
– Analytical and Computational Efforts
17
WHY NUCLEAR THERMAL PROPULSION?
•
•
•
•
•
NTP improvement: 100-400 percent over best conventional rocket motors
Large gain in DV, Isp possible with NTP rockets
Operate for short time ~ 1-3 HRS to achieve desired DV
Highly reduced mission times (12-14 months vs. 2-3 years to Mars)
Combination of temperature and low molecular weight: Isp ~ 900 s (2 x SSME)
18
•
•
•
•
•
•
Rover/NERVA, GE-710, ANL (1955-1973)
Soviet Union (195?-1986)
SDI (1983-1988)
SEI (1989-1993)
INSPI (UF) /LUTCH (1993-1997)
INSPI (1992-Present)
RD-0410 Nuclear Thermal Engine
BACKGROUND: REVIEW OF PROGRAMS (1955-PRESENT)
19
ROVER/NERVA HISTORY
• Main objective of Rover/NERVA (Nuclear Engine for Rocket Vehicle
Application) was to develop a flight-rated thermodynamic nuclear rocket engine
– Initially program and engine designed for missile applications
– 1958: NASA use in advanced, long-term space missions
• Reactor Tests:
– Kiwi-A, Kiwi-B, Phoebus, Pewee, and the Nuclear Furnace, all conducted by
Los Alamos to prove concepts and test advanced ideas
• Rocket Engine Tests:
– Aerojet and Westinghouse tests: NRX-A2 (NERVA Reactor Experiment), A3,
EST (Engine System Test), A5, A6, and XE-Prime (Experimental Engine).
• Conducted at Nuclear Rocket Development Station at AEC's Nevada Test Site
• Late 1960's and early 1970's, Nixon Administration cut NASA and NERVA
funding cut dramatically and ultimately project ended in 1973
20
COMPARISON OF REACTORS TESTED IN
ROVER/NERVA PROGRAM
21
PICTURES FROM ROVER/NERVA TESTING
22
NTP BASIC OPERATION
• NTP rockets utilize fission energy to heat a reactor core to high temperatures
• Monopropellant H2 coolant/propellant flowing through core becomes
superheated and exits engine at very high exhaust velocities
23
CORE DETAILS AND MAJOR COMPONENTS
24
MORE CORE DETAILS: KIWI 4B
25
NTP BASIC OPERATION
26
NUCLEAR-FUEL MATERIALS
• Uranium – “The mother of all nuclear fuels”
– Uranium ‘found’ in 1727, ‘discovered’ as a unique, half-metal in 1789
– Concept of nuclear fission first introduced in 1939
– U3O8: U234, U235, U238
– Pu239 (may also be formed from U238)
– Th232 (→ U233) Note: BLUE: fissionable fuel, RED: source material
– Fuel is highly enriched (90-99% U235 present)
• Most important properties: Nuclear properties (cross sections, particle behavior,
burn up), physical, thermal, mechanical, chemical (hot H2) effects)
• About 10 billion nuclear fuel atoms undergoing fission / cm3-sec in reactor core
– May be varied to control Temperature very acutely
• Fission process is independent of propellant/coolant flow
600,000 pounds of chemical fuel = 1 pound of nuclear fuel
27
THERMAL GENERATION MECHANISM: NUCLEAR FISSION
Various coolants may be used:
H2O, Ar, He, liquid metals, H2
U  n  1  2  2.5n  radiation
Total Radiation Exposure Mission to Mars: NTP < Chemical Rocket
28
IF NTP SO GOOD, WHY HASN’T IT HAPPENED?
“Sounds too much like Buck Rogers!”, President Eisenhower (1958)
“The day is not far off when nuclear rockets will prove feasible for space flight.” (1965)
•
“Chicken and egg syndrome"
– “It takes longer to develop a NTP system than to develop a space mission. Project
managers cannot include NTP systems in mission planning until system has been
developed and tested.”
– “If only reactors could be developed, users would emerge to claim them.”
– “NTP ready for flight tests and yet no users have come forward in ensuing decades.”
•
“Cutbacks were made in response to a lack of public interest in human space flight, end of
space race, and growing use of low-cost unmanned, robotic space probes.”
"Post Vietnam Congresses appear more concerned with perceived excesses of science and
technology, hence their abolishment of [NTP] and space committees.”
“Cynical maneuvering, vicious attacks and double dealing that led to its closing after years
of toil to prove the successful development of then Project Rover/NERVA in 1973.”
“They pushed NASA hard because it was dominated by people who built there lives
around chemical rockets they didn't want to see [nukes] come in cause they feared it.”
•
•
•
29
IF IT’S NOT NEW… WHAT IS THERE TO DO?
• Fuel sets upper limit of NTP performance
–
–
–
–
“No fuel geometry or material ever totally solved NERVA fuel degradation problem.”
“Mass loss limits life by causing significant perturbation to core neutronics.”
“Crack development in fuel element coating was never completely eliminated.”
“Non-nuclear testing of coated elements revealed relationship between diffusion and
temperature. For every 205 K increase, mass loss increased by factor of ten.”
• Limited experimental data at temperature, temperature ratio, heat flux, L/D for H2
– “Correlations have not been verified experimentally at heat flux levels present in
coolant channels and accuracy and applicability of these equations is in question.”
– “Even though Re, Pr, L/D within stated range of accuracy for existing correlations,
Tw/Tbulk ratio exceeds range of database if heat flux is high enough.”
"One overriding lesson from NERVA program is fuel and core development should
not be tied simply to a series of engine tests which require expensive nuclear
operation. Definitive techniques for fuel evaluation in loops or in non-nuclear
heated devices should be developed early and used throughout program..."
30
RESEARCH CONTEXT
A well thought out and carefully designed NTP roadmap is needed
• NTP is well investigated technology, but development / improvement remains
– Heat transfer relations, geometries, materials, etc.
• Fuel development and evaluation essential component of NTP program
– Testing at max temperatures, heat fluxes, transients, duration, re-start, etc.
• Preliminary Research Programs are Beginning to Form
– Non-Nuclear development to gain knowledge base
– Design of experiment, data acquisition and analysis
• Partnering to facilitate development
– Confluence of NASA, industry (P&W), and academia
• Hot H2 NTP experiments at MSFC
– Support / design / build-up from academia
31
APPROACH
• Non-nuclear testing in hot H2 environment key to engine development
– T=300-3200 K
– Realistic mass flow rates (0.8-1.5 g/s per) and inlet pressures (500 psi)
– Modular test section: investigate NERVA, particle bed, pebble bed, etc.
– Materials characterization and assessment of performance/stability in hot H2
– Safety, instrumentation, diagnostics, etc.
“The Rover/NERVA engine is to be used as a “reference,” against which other
concepts… will be compared.” - Dr. Stanley K. Borowski
“Solid core has plenty of growth potential. Just because it's 1960's era technology
doesn't mean it's obsolete. Object of a new program should be to build on this.”
“If you had kept on working [NERVA] you would now have a 4th generation system.
It would have Isp's over 1000, power densities 3000 MW, and maybe 30 hours of
engine lifetime with 180 stop & starts.” - Dr. James Dewar, AEC
32
RANGE OF NTP INPUT DATA
Reactor Power (MW)
Thrust (N)
Maximum Propellant Temperature (K)
Specific Impulse (s)
Ideal Exhaust Veloctiy, (m/s)
Chamber Pressure (psi)
Chamber Pressure (Mpa)
Total Mass Flow Rate (kg/s)
Number of Fuel Element
Number of Cooling Holes / Fuel Element
Power / Fuel Element (MW)
Power / Cooling Hole (W)
Mass Flow / Cooling Hole (g/s)
Cooling Hole Diameter (cm)
Coolant / Propellant
Preliminary Simulation Reactor Material
PHOEBUS 2A
5000
1113
2528
820
8036
550
3.79
138.4
4789
19
1.0441
54950
1.52
0.28
H2
Graphite, G10
NERVA
1570
334
2361
825
8085
450
3.10
41.3
1584
19
0.9912
52166
1.37
0.28
H2
Graphite, G10
Summary: Baseline Case Values
Power / Fuel Element ~ 1MW / Element for each case
Flow Rates ~ 1.5 g/s H2
Chamber Pressure ~ 3.5 MPa (~ 500 PSI)
Maximum Propellant Temperature ~ 2500 K
CERMET
2000
445
2507
930
9114
600
4.14
48.8
1600
19
1.2500
65789
1.61
0.28
H2
Tungsten
33
REACTOR TEMPERATURE DISTRIBUTION MODEL
Model: Single H2 cooling passage within single element
Test sub-section to replicate various portions of cooling path
Match power input, H2 temperature and wall temperature at various x/D, r/D locations
Cooling Hole ID: 0.1-0.125 inches / Cooling Holes OD: 0.183 inches
L/D ~ 500 for NERVA elements
34
H2 COOLANT / PROPELLANT COMMENTS
• Range of interest
– T=300-3000 K, P=14.7-1000 psi
• Important to model H2 properties accurately
– Up to ~ 1500 K, pressure has little effect on Cp, g, m, k
– For T > 1500 K, must include pressure effect on thermal transport properties
• References: NASA SP [King],[Kubin and Presley],[Weber], [McCarty], [Patch]
• Dissociation  with  P at constant T
– P=1 ATM, T=3000 K, NH=15 % vs. P=40 ATM, T=3000 K, NH=2.6 %
– Isp improvement with dissociation, but no impact on cf
• Ionization: not relevant at these temperatures
– P=1 ATM, T=3000 K, NH+ ~ O(10-11) vs. P=40 ATM, T=3000 K, NH+ ~ O(10-12)
• Compressibility: small effect
– P=1 ATM, T=3000 K, Z=8.2 % vs. P=40 ATM, T=3000 K, Z=1.3 %
35
H2 DATABASE: IMPORTANT TO CAPTURE T & P
36
H2 COMMENTS: UNIQUE BEHAVIOR
P=1 ATM, T=2600 K
NH=3.9 %, NH+=1.3x10-13
P=40 ATM, T=2600 K
NH=0.6 %, NH+=8.8x10-15
Effects of dissociation and ionization on cp, k are dramatic
Higher pressures  dissociation suppressed
NTP “nominal” range of operation T < 3500 K and P ~ 20-40 ATM
37
H2 COMMENTS: VACUUM SPECIFIC IMPULSE
Phoebus 2A, Ispvac~918 s
Vacuum Isp equation corrected for dissociation
Isp based on channel exit temperature, not mixed-out temperature
Mixed-out temperature (model) ~ 100-300 K lower than exit temperature, 10% Isp 
38
H2 COMMENTS: DISSOCIATION AND ISP,VAC
As Pch , mass flow , Thrust (T/W) 
System optimization for required T/W vs. Isp, future work consideration
Max material (U,Zr,Nb)C temperature ~ 3300 K (1hr): Max Tmelt (TaC, HfC) ~ 4200 K
39
HYDRODYNAMIC CONSIDERATIONS/MODELING
• Laminar and Turbulent Regions, critical Reynolds numbers
– ReD ~ 2,300 onset of turbulence
– ReD ~ 10,000 for fully turbulent conditions
– ReD ~ 70,000 for Phoebus/NERVA
• Entrance and fully developed region
– No satisfactory general expression for entry length in turbulent flow
– Fully-developed turbulent flow for x/D > 20 (approx. independent of ReD)
• Pressure drop and inlet/exit boundary conditions
• Total pressure decrease due to constant volume heat addition (~7 %)
• Thermal choking: Only 1/3rd of total DTt,max/Tt capacity
• H2 attack on core / degradation
– Corrodes/erodes away channel wall and protective coatings, “Scouring” action
– Radial pressure drops (channel to channel) which shakes core modules
– Mass loss and cracking of elements
40
REACTOR POWER DISTRIBUTION
41
REACTOR TEMPERATURE DISTRIBUTION
42
TEMPERATURE DISTRIBUTION: COMMENTS
• Note that Tbulk maximum at L=100%
• Maximum inner and centerline wall temperatures at L ~ 80%
• For metals, Re, Ta, W, TCL and TID close ~ 50 K
– For actual NTP materials, TCL and TID exhibit larger DT ~ 100-500 K
– For actual NTP materials, TCL and TID not at same axial location
• Location of maximum Twall-Tbulk, Axial Twall, Axial Tbulk all located in midband region
• Mid-band region of max corrosion from NERVA reports:
– "Corrosion most pronounced in mid-range region, about a third of distance
from cold end”
– “Fuel operating temperatures lower here than fabrication temperatures, hence
thermal stresses higher than at hot end. Also, neutron flux highest in this
region..."
• Flow time ~ 6 ms, Velocities ~ 1000 m/s at exit, but M ~ 0.2
• 55 kW to single cooling channel for H2 simulation
43
HEAT TRANSFER COEFFICIENT: VARIOUS FORMS
Nu D 
hg D
k
 A Re 4D 5 Pr B RCF  ACF 
C
D
• Various heat transfer correlations may be applicable within operational range
– Differ by up to 20% (not to mention H2 data uncertainty)
– Correlations at such elevated conditions, “that do exist have not been verified
experimentally at the heat flux levels present in coolant channels and
accuracy and applicability of these equations is in question.”
Equation Set
Reference
Nusselt
Number
A1
Given and Anghaie, Eq.8, [Error!
Reference source not found.]
Nu D , iso 
Re D  1000 P r f 2
12
1.07  12.7 P r 2 3  1 f 2 
B2
Incropera and De Witt, Eq. 8.60,
[Error! Reference source not
found.]
NuD  0.027 Re 4D 5 Pr n
C3
Incropera and De Witt, Eq. 8.62, [Error!
Reference source not found.]
Nu D 
nheating Twall  Tbulk   0.4
 f 8Re D  Pr
12
1.07  12 .7 f 8 Pr 2 3  1
ncooling Twall  Tbulk   0.3
Fanning
Friction
Factor6
Radial
Correction7
1
f  0.0014  Re D0.32
8
T 
Nu D  Nu D ,iso  wall 
 Tbulk 

 T 
n   log 10  wall 
 Tbulk 

Axial
Correction
Range of
Validity
n
f  1.82 log10 Re D  1.64 
2
Non-Applicable
Incropera and De Witt, 8.61, [Error!
Reference source not found.]
14
 0 .3
  x   0 . 7  T  0 .7 
Nu D  x   Nu D 1     wa ll  
  D   Tb u lk  
0.5 < Pr < 2000
2300 < ReD < 5x106
 m 

Nu D  0.027 Re 4D 5 Pr1 3 
 m wall 
Nu D 
 f 8Re D  1000  Pr
12
1  12 .7  f 8  Pr 2 3  1
f  0 .079 ln Re D  1 .64 
None Provided
None Provided
None Provided
None Provided
0.5 < Pr < 2000
104 < ReD < 5x106
0.5 < Pr < 2000
2300 < ReD < 5x106
0.14
None Provided
0.7 < Pr < 160
ReD > 10,000
L/D > 10
D4
Incropera and De Witt, Eq. 8.63, [Error!
Reference source not found.]
0.7 < Pr < 16,700
ReD > 10,000
L/D > 10
E5
Bartz Heat Flux Formula
Nu D  0.026 Re 0D.8 Pr
2
 T
Nu D  0.026 Re 0D.8 Pr
T
 avg
w  0 .6




0. 8  0 .2 w
44
HEAT TRANSFER COEFFICIENTS
45
HEAT TRANSFER EXPERIMENTAL SCALING


0.026
0.8
0.2  T 
hg  0.2  ru  c p m
T 
D
 avg 
•
•
•
•
•
0.68
Convective coefficient scales with diameter as hg~1/D0.2
– Doubling tube diameter will decrease hg by 13%
– Smaller diameters lead to larger heat fluxes (from Reynolds dependence on Cf)
Heat flux almost linear with pressure, scales as hg~r0.8~p0.8
– Halving inlet pressure will reduce coefficient by 57%
Lighter gases lead to higher heat fluxes, hg~1/M0.6
– Ratio of molecular weights of Ar:H2 ~ 20, heat flux for Ar:H2 ~ 16%
Evaluation of viscosity term is also important both at wall and fluid temperatures
– Accounts for differences in gas temperature within boundary layer and bulk flow
– Exponent less than unity, acts as enhancement of heat transfer coefficient
Careful evaluation of cp, m, k
46
SAMPLE MODEL OUTPUTS:
FLOW VARIABLES VS. AXIAL LOCATION
47
OVERVIEW
• Rocket Overview
– Categorization of various types of Rockets
– Rocket Mission Selection Guide
– Rocket Performance Parameters
• Nuclear Thermal Propulsion
– Historical Overview
– Hot Hydrogen Properties
– Fluid Mechanic and Heat Transfer Modeling
– Simulation Results
• Future Work – What can we do at FIT?
– How to Simulate Nuclear Reactors for Space Applications
– New Experimental Facility
– Analytical and Computational Efforts
48
BUILD-UP OF EXPERIMENT
• Surrogate test gases to build-up experiment in less-complex, cost effective way
– H2 and hot H2 logistics and safety precautions
– Reduced power requirements
• Development with bench-top 12.5 kW induction system
– Verification of experimental set-up, diagnostics, heat transfer correlations
– Reduced cost elements (Ta) vs. materials ~ 100% dense to H2 (Re)
• Make use of surrogate test gases, such as He, N2, and Ar
– Investigate cooling channel using 12.5 kW power supply
– Using Ar, test entire elements (19 cooling channels) at PRL using 100 kW
• Using surrogate test gases, match:
– Non-dimensional and actual temperatures
– Heat fluxes
– Heat transfer coefficients
• Scale power input, mass flow, gas type, etc.
49
SURROGATE TEST GASES: He, N2, Ar
H2
He
N2
Ar
50
PRELIMINARY TEST MATRIX
•
Test Series 1: Cold Flow Tests Using He, N2 or Ar
– Objectives: Verify design, instrumentation, sealing, operation, etc., T~ 800 K
– Materials: Stainless Steel ($80/tube)
•
Test Series 2: Hot Flow Tests Using He, N2 or Ar
– Inductive heating of test specimen to T~3000 K
– Verify power/temperature distribution of test specimen
– Heat flux correlations
– Materials: Tantalum ($800/tube)
•
Test Series 3 at MSFC: Cold Flow H2
– H2 safety check-out, sealing, test emergency shut-down
•
Test Series 4 at MSFC: Hot Flow H2, Full Cooling Channel Simulation
– Inductive heating of test specimen
– Material assessment, H2 corrosion, impact on heat transfer correlations, etc.
– Materials: Rhenium ($8,000/tube), make use of actual non-enriched
elements/material
51
PRELIMINARY EXPERIMENTAL CONFIGURATION
•
•
•
Initial Test Chamber
– 77.6 inch (Full scale test article: L=55 inches), 8.25 inch OD Chamber
– 16 ports already in place, D=1.38 inches
• 12 located near ends, 4 located near center
• Induction in and out feeds, vacuum, pyrometer access, instrumentation, etc.
– Chamber modifications
• Vacuum ready, outer cooling jacket, ports to capture mid-band and peak
• 1 inch bellows fittings to relieve thermal expansion of material
– Re, Ta, W, expect 0.5-0.75 inches thermal expansion at max T
– Radiation loss modeling
• Loss estimate ~ 10-20 kW for 12.5 kW, need GRAFOIL insulator
Induction Heating
– Heating material with alternating EM field, 150 < f < 350 kHz vs. d penetration
– Coil design for sinusoidal power distribution ~ 1/r2
– Design for test coupon, tubular (prismatic) and particle bed reactor type
Test Duration
– H2 11 min/bottle, 4 hour H2: 23, He: 12, N2: 2, Ar: 2
52
SUPPORT ANALYSES IN PROGRESS
•
•
•
•
•
•
Reactor Power Profile Optimization
– “From a nuclear rocket design standpoint, a flat power profile may not be best
configuration and that an optimum power profile probably exists that gives that lowest
fuel temperatures for a given core and operating condition.”
Mixing Model
– Mixed out flow temperature for a given radial profile and number of elements
– Compound flow, vorticity generation, mixing time scale, void support structure
– Sample result: Tmix ~ 300 K lower than (Texit)max, 10% Isp 
Optimization of T/W vs. Isp for low pressure operation
NTP Materials: Behavior of UyZr1-yC1-x (Fuel) and ZrC1-x (Coating)
Plug Nozzle vs. Traditional Bell
– “Some of the things that have been rejected in the chemical engines, such as
expansion-deflection nozzles, spike nozzles, and plug nozzles, all become candidates
for reexamination to see what would be the optimum way to design a thrust
chamber/nozzle for hydrogen recombination.”
Potential for tailoring of flow path cross-sectional area
– Minimum area located at maximum heat transfer locations
– Minimize potential heat transfer hot-spots
53
SUMMARY
• NTP is well investigated technology, but development / improvement remains
– Heat transfer relations, geometries, materials, etc.
• Fuel development and evaluation essential component of NTP program
– Testing at max temperatures, heat fluxes, transients, duration, re-start, etc.
• Preliminary Research Programs are Beginning to Form
– Non-Nuclear development to gain knowledge base
– Design of experiment, data acquisition and analysis
– Various expertise essential (materials, diagnostics, hot H2, etc.)
• Partnering to facilitate development
– Confluence of NASA, industry (P&W), and academia (FIT, UF)
• Hot H2 NTP experiment at MSFC
– Support / design / build-up from academia
54
“BREAKTHROUGH” IDEA #1 ?
• Significant gains possible with high T + low P operation → H2 dissociation
– However, low P implies low mass flow → low thrust
– Dissociation driven by static temperature
– Heat transfer driven by total temperature
• Current channels, constant cross sectional area
• Introduce converging-diverging geometry within channel
– Choke mass flow to desired value upstream, retain high thrust
– Large Dp downstream, continuously heat, integrated nozzle/channel
– Recombination in final expansion portion → double benefit !
• Approach: 1-D finite differencing of full influence coefficients (Mach parameter)
– Variable cp and W
– Area ratio optimization, geometric confinement and friction
55
SCHEMATIC REPRESENTATION
Traditional: Constant Area
Only dissociation possible with high static T, Mach ~ 0.2
New: ‘Supersonic Core’
Retain high total temperatures for heat transfer
Static pressure drop for dissociation
Potential DIsp ~ 150 seconds
Integrated nozzle
56
“BREAKTHROUGH” IDEA #2 ?
• Chemical rocket propulsion system benefit from scaling
– T~A, W~V, T/W~1/L
• Does not appear NTP scalable due to critical mass
• Examine use of radioisotope as heat source
– Used on prior space missions, but for electrical
– Trade of half-life vs. specific power
• Candidates Po210, Pu238, Cm242, 244
– Examine scalability
– Deep space missions, Isp ~ 700-800 s (H2)
– Metal foil bonding technique (W, Re possible)
• White paper design in progress with LLNL
57
SUPPLEMENTAL SLIDES
58
TYPES OF ROCKETS
LAUNCHERS
SPACECRAFT
SPACE STATIONS
Atlas (USA)
Mercury (USA)
Skylab (USA)
Delta (USA)
Gemini (USA)
Salyut (USSR)
Titan (USA)
Apollo (USA)
Mir (Russia)
Pegasus (USA)
Shuttle Orbiter (USA)
ISS
Saturn (USA)
Vostok (USSR)
Space Shuttle (USA)
Soyuz (Russia)
A-Vehicle (Russia)
Proton (Russia)
Long March (China)
59
ROCKETS: ENERGY VS. POWER LIMITED
•
•
Chemical Rockets are Energy Limited
Unit of Energy: JOULE, Energy=F*Displacement =[kg m/s2]*[m]==[kg m2/s2]
– Quantity of energy (per unit mass of propellant) that can be released during combustion
is limited by fundamental chemical behavior of propellant
– Low Isp: high thrust, launch, high thrust escape at perigee
•
•
Electrical Rockets are Power Limited
Unit of Power: WATT (J/s), Power= F*V=[kg m/s2]*[m/s]==[kg m2/s3]
– Usually a separate energy source is used (nuclear or solar) and much higher propellant
energy is possible
– However, rate of conversion of nuclear or solar energy to electrical energy and thence
to propellant kinetic energy is limited by mass of conversion equipment required. Since
mass is large portion of total mass of vehicle, electrical rocket is essentially power
limited
60
CHEMICAL: LIQUID VS. SOLID ROCKETS
Liquid Rockets, Shuttle Main Engines
Fuels: Liquid hydrogen and liquid oxygen
Advantages
– High Thrust, throttle, shutdown
Disadvantages
– Highly complex (plumbing, cooling,
steerting, throttle, structures, etc.)
Liquid-Propellant
Rocket Engine 11D33
Solid Rockets, Shuttle SRB
Fuel: Aluminum and Nitrate
Advantages
– Simple, low cost, safe
Disadvantages
– Thrust cannot be controlled, no shut down
61
EXAMPLE: ATLAS / CENTAUR
•
•
•
Independently developed by USAF as first ICBM,
cold war mission to deter nuclear attack
Part of Project Mercury. Mission goal to put a
human into orbit, accomplished Feb. 20, 1962.
Used today to launch payloads into orbit
ATLAS CENTAUR FAMILY RECORD
• First launch: 8-May-1962
• Number launched: 97 to end-1995
• Launch sites: Cape Canaveral pads 36A/B;
Vandenberg AFB SLC-3E from 1998
• Vehicle success rate: 86.60% to end-1995
• Success rate, past 20 launches: 100% to end-1995
For more on Atlas / Centaur Rockets:
http://users.commkey.net/Braeunig/space/specs.htm
62
EXAMPLE: ATLAS IIAS
•
•
•
•
•
•
•
47 m tall, 3-4 m diameter, 234,000 kg
Lockheed-Martin 2-stage liquid propellant (LOXRP1) booster
$95-105M per launch
First stage booster section
– 2 Rocketdyne engines
– 1.84 MN thrust, Isp=263 seconds
– runs about 3 minutes
Second stage is “sustainer section”
– 1 Rocketdyne engine
– 269 KN thrust, Isp=220 seconds
– 5 minutes burn with booster
Strap-on solid rockets
– Four Thiokol Castor IVA SRMs
– 433 KN thrust, Isp=229 seconds
– 9 m tall, 1 m diameter, 12K kg
– Burn about 56 seconds
Uses a Centaur upper stage
– 2 Pratt & Whitney engines
– LOX-LH2
– 185 KN thrust, Isp=449 seconds
– 1st engine runs about 5 minutes
– 2nd engine runs about 1-2 minutes
63
EXAMPLE: DELTA
•
•
•
In use since 1960, Delta launched successfully over
250 times
Scientific satellites placed into orbit by a Delta
rocket include IUE, COBE, ROSAT, EUVE, WIND,
RXTE, Iridium, Navstar GPS
Manufactured for USAF and NASA by Boeing
DELTA FAMILY RECORD
• First launch: 13-May-1960
• Number launched: 230 to end-1995
• Launch sites: Cape Canaveral pads 17A/B;
Vandenberg AFB SLC-2W
• Vehicle success rate: 94.8%
• Success rate, past 25 launches: 100%
For more Delta Rockets:
http://users.commkey.net/Braeunig/space/specs.htm
64
EXAMPLE: TITAN
•
•
•
•
•
•
Titan is a family of expendable rockets.
Most Titan’s are derivatives of Titan II ICBM.
Titan III is stretched Titan II with optional solid
rocket boosters. Used to launch NASA scientific
probes such as the Voyagers.
Titan IV is stretched Titan III with non-optional
solid rocket boosters. Used to launch US Military
payloads, NASA's Galileo and Cassini probes to
Jupiter and Saturn.
Titan IV is a horrendously expensive launch
vehicle.
Currently, three Titan IVBs remain to be launched,
no more ordered. Current owners of Titan line
(Lockheed-Martin) decided to extend Atlas family
instead of Titans. By 2005 the Titans will likely be
extinct.
For more on Titan Rockets:
http://users.commkey.net/Braeunig/space/specs.htm
65
EXAMPLE: STS
•
•
•
•
Space Shuttle developed by NASA. NASA
coordinates and manages, oversees launch and
space flight requirements for civilian and
commercial use.
STS consists of four primary elements: orbiter
spacecraft, two Solid Rocket Boosters (SRB), an
external tank for three Shuttle main engines
Shuttle will transport cargo into near Earth orbit
100 to 217 nautical miles (115 to 250 statute
miles) above the Earth. Payload is carried in bay
15 feet in diameter, 60 ft long.
1st Launch: April 12, 1981, 7:00:03 a.m, EST.
QUESTION:
How many rockets systems on STS?
For more on STS:
http://users.commkey.net/Braeunig/space/specs.htm
66
EXAMPLE: A-VEHICLE (RUSSIA)
•
•
•
A-class Soviet launch vehicles are based on
Soviet SS-6 ICBM
Vehicles in this class are Vostok, Soyuz and
Molniya launchers
Three vehicles all use same core stage and four
strap-on boosters (liquid oxygen and kerosene
propellant)
QUESTION:
Why does this rocket have many primary
engines (20 in picture) instead of 1 or 2 primary
engines?
Note: Saturn V was powered by 5 F-1 engines.
Why not just use 1 big one?
For more on A-Vehicles:
http://users.commkey.net/Braeunig/space/specs.htm
67
EXAMPLE: PROTON (RUSSIA)
•
•
•
•
Proton medium-lift launch vehicle 1965
First Russian launcher not based on a ballistic
missile prototype.
Proton used in 3 and 4-stage versions, with 3stage version used for many of Mir support
missions. 4-stage Proton used primarily for
geostationary satellite missions.
First stage incorporates 6 strap-on boosters,
provides over 2 million pounds of thrust. 3-stage
Proton launch vehicle can place over 44,000
pounds into LEO, will be used for largest of ISS
components that are launched by RSA.
For more on Proton Rockets:
http://users.commkey.net/Braeunig/space/specs.htm
68
ROCKET CLASSIFICATION 2
•
•
•
Another way to classify rocket engines depends on propellant (gas) acceleration
mechanism or the force on the vehicle mechanism
– Thermal
• Gas pushes directly on walls by pressure forces
• Nozzle accelerates gas by pressure forces
• Most large rockets, chemical, nuclear, some electric (arcjet, resistojet)
– Electrostatic
• Ions accelerated by E field
• Electrostatic force (push) on electrodes (Ion Engines)
• Force (push) on magnetic coils through j (Hall Thrusters)
– Electromagnetic
• Gas accelerated by j x B body forces
• Force (push) on coils or conductors (magnetoplasmadynamic (MPD))
Distinction between Chemical and Electrical
– Energy vs. Power Limited
Other types
– Nuclear, Pulse Detonation, Air-Breathing (Hybrids), vehicle caries only fuel and takes
oxidizer from air, Photon (ejection), Solar Sail (radiation pressure via absorption)
69
ROCKETS IN USE TODAY
2 Primary Classes: Chemical and Electrical
•
Liquid Rockets (Chemical: Energy Limited)
– Gas feed or turbopump supplied
– Liquid propellants, mix and burn in combustion chamber
– Almost all launch vehicles for space are liquid rocket engines
– More thrust per pound (T/W) than solid rockets
•
Solid Rockets (Chemical: Energy Limited)
– Solid propellant inside pressure tube, no separate combustion chamber, entire rocket
burning on inside, fuel and oxidizer mixed together (fireworks).
– Several ways to burn. From end up (like a cigarette), or from center outward. Grain
may be circular or star-shaped.
– Once started cannot be shut off until they burn-out
– Solid rocket motors can be stored for months or years without leaking or degrading.
Missiles have to sit for years, then used quickly and without delay.
– Strap-on rockets of shuttle and other launch vehicles are solid rocket motors
China 800-1200 AD, War of 1812, ‘…rockets red glare’ National Anthem
70
ROCKETS IN USE TODAY
•
Ion (Electrical, Electrostatic: Power Limited)
– Electricity to accelerate a small amount of gas VERY fast, O(1000 km/s)
– Strip off electron, accelerate gas very fast, neutralize and eject. Typical gas is Xenon
(heavy, inert, non-radioactive gas)
– Electrical source: solar to high powered nuclear sources (such as radiographic
thermal generators (RTG)).
– Extremely high specific impulse, lowest thrust.
• Useless in atmosphere and as a launch vehicle. Highly useful in space.
– Used today as final thruster to higher orbit, adjusting orbits, station keeping
•
Arc Jet (Electrical, Electrothermal: Power Limited)
– Short low power thrusters (station keeping). Non-flammable propellant is heated by
electrical heat source (coil). Expanded and expelled at high speed.
– Propellant expelled is not combusted, just heated (typically changing phase from
liquid to gas) so that it is under pressure.
– Systems are low thrust, very reliable, may use electrical power from solar sails or
batteries
71
FUTURE LAUNCH VEHICLES?
Linear Aerospike Engine
Taurus
Minotaur
More Information:
http://www.spaceandtech.com/spacedata/rlvs/venturestar_sum.shtml
http://www.orbital.com/LaunchVehicle/SpaceLaunchVehicles/index.html
http://www.aerospaceweb.org/design/aerospike/main.shtml
Pegasus
72
EXAMPLE: AUTOMOBILE AIRBAG
Airbags have been
clocked at 300 MPH.
Most airbags deploy
at 200-300 mph.
Side airbags deploy
at 3 times speed of
frontal airbags
•
•
•
Airbag inflators are a spin-off of military
and rocket industries
Equivalent of solid rocket booster
Major suppliers of inflators is rocket fuel
manufacturer, Morton Thiokol (also make
space shuttle boosters).
Why Rocket-type? How does it work?
• To ignite, 12 volt input from airbag control
computer, heats a resistive wire element
initiating exothermic chemical reaction
which decomposes sodium azide (NaN3) in
a three step process. Chemical deflagration
includes potassium nitrate (KNO3) and
silicon dioxide (SiO2).
• Sodium azide (NaN3) and potassium nitrate
(KNO3) react very quickly to produce a
large pulse of hot nitrogen gas
73
EXAMPLE: ELECTRIC AND ION THRUSTERS
•
Designer: Rocketdyne.
•
Developed in: 1999.
Propellants: Electric/Xenon
Thrust (vac): 0.001 N
•
Isp: 3,500 s
•
Satellite orbit raising and station-keeping
applications.
Thrust created accelerating positive ions
through gridded electrodes, more than 3,000
tiny beams of thrust.
Ions ejected travel in an invisible stream at a
speed of 30 kilometers per second (62,900 miles
per hour), nearly 10 times that of its chemical
counterpart.
Ion thrusters operate at lower force levels,
attitude disturbances during thruster operation
are reduced, further simplifying the
stationkeeping task.
For more on Electric Propulsion:
http://hpcc.engin.umich.edu/CFD/research/NGPD/ElectricP
ropulsion/
http://www.marsacademy.com/propul/propul7.htm
http://richard.hofer.com/electric_propulsion.html
http://www.stanford.edu/group/pdl/EP/EP.html
No Combustion
74
EXAMPLE: ELECTRIC AND ION THRUSTERS
•
Designer: Rocketdyne.
•
Developed in: 1999.
Propellants: Electric/Xenon
Thrust (vac): 0.001 N
•
Isp: 3,500 s
•
Satellite orbit raising and station-keeping
applications.
Thrust created accelerating positive ions
through gridded electrodes, more than 3,000
tiny beams of thrust.
Ions ejected travel in an invisible stream at a
speed of 30 kilometers per second (62,900 miles
per hour), nearly 10 times that of its chemical
counterpart.
Ion thrusters operate at lower force levels,
attitude disturbances during thruster operation
are reduced, further simplifying the
stationkeeping task.
For more on Electric Propulsion:
http://hpcc.engin.umich.edu/CFD/research/NGPD/ElectricP
ropulsion/
http://www.marsacademy.com/propul/propul7.htm
http://richard.hofer.com/electric_propulsion.html
http://www.stanford.edu/group/pdl/EP/EP.html
No Combustion
75
EXAMPLE: NUCLEAR POWER
•
Project Prometheus will develop the means to efficiently increase power for spacecraft, thereby
fundamentally increasing our capability for Solar System exploration.
•
Space fission power can be used as the power source to provide large amounts of electricity for
electric propulsion systems (Nuclear Electric Propulsion)
•
The heat generated by the fission process can be used directly to create thrust (Nuclear Thermal
Propulsion)
Increased power for spacecraft means not only traveling farther or faster, but it also means exploring
more efficiently with enormously greater scientific return.
High levels of sustained power would permit a new era of Solar System missions designed for agility,
longevity, flexibility, and comprehensive scientific exploration
Today, only nuclear power can enable scientifically vital, but incredibly challenging missions
•
•
•
76
EXAMPLES OF NUCLEAR PROPULSION
No Combustion In These Devices
77
ROCKET SELECTION GUIDE
MISSION REQUIREMENT
ROCKET TYPE
1. Non-Space Missions
– Atmospheric / Ionospheric Sounding
– Tactical Missiles
– Medium-Long Range Missiles
Solid Propellant, 1-4 stages
Solid Propellant, 1-2 stages
Solid or Liquid Propellant, 2-3 stages (very high
acceleration)
2. Launch to Space
3. Impulsive DV in Space
• Time critical maneuvers
• Energy change from elliptic orbits, plane
change from elliptic orbits
• Non-fuel limited situations
4. Low Thrust DV in Space
• Mass-limited missions
• Non-time critical missions
• Small, continuous orbit corrections, near
circular orbits
Solid, liquid or combination, 2-4 stages (2-4g),
Possible: hybrid, 2-4 stages
Small solid propellant (apogee kick, etc.)
Bi-propellant (storable), liquids, monopropellant
(storable) liquids. Future: nuclear thermal
Solar-electric systems:
Arcjet (a bit faster, less Isp), Hall, Ion (slower,
higher Isp), PPT (precision maneuvers),
Nuclear-electric systems, direct solar-thermal
78
PERFORMANCE MEASURES: THRUST
•
•
•
•
•
Thrust, (T, F), Thrust to Weight Ratio, (T/W, F/W)
Thrust is the force that propels a rocket or spacecraft and is measured in pounds (lbf),
kilograms (kgf) or Newtons (kg m/s2)
– Result of pressure force which is exerted on wall of combustion chamber
– Existence of pressure force results in a momentum flux
Weight is measured in pounds (lbf), kilograms (kgf) or Newtons (kg m/s2)
T/W is a non-dimensional metric
Some Example Numbers
– Very Large: 20-100, Chemical Rockets
– Medium: 5-20, Nuclear
– Very Low: O(10-3), Solar, Electric Propulsion, Power Limited)
– Typical payload ration ~ 0.02 (mass of payload/mass of entire rocket)
– Engines ~ 2 x payload
– Combustion Temp ~ 2500-4500 K, Ve ~ 1500-4500 m/s
79
PERFORMANCE MEASURES: THRUST
•
For our simple rocket we had (Pe=Pa):
 eVe
F m
•
For a given exit momentum flux relative to rocket, thrust is independent of flight speed
of vehicle.
Could a rocket vehicle be propelled to a speed much higher than the speed at which the
jet leaves the rocket nozzle?
How about for an airplane?
80
PERFORMANCE MEASURES: SPECIFIC IMPULSE
•
•
Specific Impulse, Isp, (measured in seconds)
Specific impulse is the amount of thrust you get for the fuel weight flow rate
F
Isp 
m e g earth
Isp 
•
•
•
F
m e g earth
kg  m 

N

s 2 


 s 
kg  m
kg  m
2
 s  s
 s  s 2
m eVe
V

 e
m e g earth g earth
 
 
ge is measured on the earth’s surface, ge=9.8 m/s2
Some Example Numbers
– Chemical rocket range: 200-500 s (500 is just about the limit)
– Shuttle Main Engine: 455 s (T=1670 kN each), SRB: 250 s (T=14700 kN each)
– Nuclear Thermal: 800-1200 s
– Trade-off vs. mass for EP, 500-6000 s
– Nuclear Electric Rocket: 20,000 s (T/W=0.0001)
Specific impulse improves with LOW molecular weight, LOW specific heat ratio, and HIGH
temperature
81
Isp
ROCKET VS. TURBOJET ISP
Mach Number
82
NUCLEAR THERMAL ROCKET APPLCIATIONS
•
•
•
“This gigantic (nuclear) missile would dwarf the V-2. Even though a practical design
might reduce considerably the amount of propellant required, nuclear powered rockets
seem remote.” (1946)
Limitations of Nuclear Propulsion for Earth to Orbit. (2001 NASA Study)
– Only very best… reactors might be applicable for earth-to-orbit
However, in terms of high mass, space travel, NTP is among the best
– Proven concept
– Marriage of two well proven technologies
1. Liquid, chemical rocket development
2. Solid-fuel nuclear reactors
83
KIWI-A PRIME ATOMIC REACTOR
•
Kiwi-A Prime is one of a series of atomic reactors for studying the feasibility of nuclear rocket
propulsion, in Los Alamos, New Mexico. Developed by the Los Alamos Scientific Laboratory for the
U.S. Atomic Energy Commission, the reactor underwent a highly successful full-power run on July
8, 1960, at Nevada Test Site in Jackass Flats, Nevada. Kiwi was a project under the National Nuclear
Rocket development program, sponsored jointly by Atomic Energy Commission and NASA as part
of project Rover/NERVA (Nuclear Engine for Rocket Vehicle Application).
84
XECF
•
The first ground experimental nuclear rocket engine
(XE) assembly, in a "cold flow" configuration, is
shown being installed in Engine Test Stand No. 1 at
the Nuclear Rocket Development Station in Jackass
Flats, Nevada. Cold flow experiments are conducted
using an assembly identical to the design used in
power tests except that the cold assembly does not
contain any fissionable material nor produce a
nuclear reaction. Therefore, no fission power is
generated. Functionally, the XECF (Experimental
Engine Cold Flow) is similar to the breadboard
nuclear engine system (NERVA Reactor
Experiment/Engine System Test or NRX/EST) tested
in 1966, except that the experimental engine more
closely resembles flight configuration. In addition to
the nozzle-reactor assembly, the XCEF has two
major subassemblies: an "upper thrust module"
(attached to test stand) and a "lower thrust module"
containing propellant feed system components. This
arrangement is used to facilitate remote removal and
replacement of major subassemblies in the event of a
malfunction. The cold flow experiential engine
underwent a series of tests designed to verify that the
initial test stand was ready for "hot" engine testing,
as well as to investigate engine start up under
simulated altitude conditions, and to check operating
procedures not previously demonstrated. The XECF
engine was part of project Rover/NERVA.
85
JFK VISIT
•
President John F. Kennedy departs from the
Nuclear Rocket Development Station, after a
brief inspection visit on December 8, 1962. At
the President's left are: Dr. Glenn T. Seaborg,
Chairman of the U.S. Atomic Energy
Commission; Senator Howard Cannon, (DNV); Harold B. Finger, Manager of the Space
Nuclear Propulsion Office; and Dr. Alvin C.
Graves, Director of test activities for the Los
Alamos Scientific
86
PERFORMANCE COMPARISON
87
TYPES OF ROCKETS
LAUNCHERS
SPACECRAFT
SPACE STATIONS
Atlas (USA)
Mercury (USA)
Skylab (USA)
Delta (USA)
Gemini (USA)
Salyut (USSR)
Titan (USA)
Apollo (USA)
Mir (Russia)
Pegasus (USA)
Shuttle Orbiter (USA)
ISS
Saturn (USA)
Vostok (USSR)
Space Shuttle (USA)
Soyuz (Russia)
A-Vehicle (Russia)
Proton (Russia)
Long March (China)
88
MOMENTUM EXCHANGE TETHERS
89
H2 COMMENTS: VARIATION IN DATA SETS
NASA SP Reports
Kubin and Presley
McCarty
Patch
Incropera and De Witt
Curve Fits
Hill and Peterson
CHEMKIN
90
•
•
•
•
H2 COMMENTS: H2 ATTACK ON CORE
H2 rapid increase in temperature (300  3000 K) and velocity (100  2000 m/s)
Under such conditions GH2 takes on aggressive characteristics and attacks core
– Chemically
• Corrodes/erodes away channel wall and protective coatings, “Scouring” action
– “Small hard pebble swirling around inside of a soft channel matrix”
– Greater flow rate, more scouring, enhanced by higher temperatures
• Penetrates into fuel-matrix structure and weakens core
– Mechanically:
• Radial pressure drops (channel to channel) which shakes core modules
Resistance to core attack depends on core type and specific design of protective coating
– TiC, ZrC, and NbC are potential coatings which are H2 resistant
Experiment should be able to study these affects over a range of core types (starting with
simple tubular/prismatic structure), materials and coatings
91
RADIATION DOSSAGE (rem)
rem=absorbed radiation dose x quality factor
92
NUCLEAR SHIELDING / REFLECTOR MATERIALS
•
•
•
Goal is to reflect neutrons back into system, attenuate radiation
The principal absorber is the core itself
– 85% go into fission fragments and are recovered as heat
– 5% go into birth of new neutrons
– 10% goes into the ejection of b and g rays, most of which can be recovered in form of
auxiliary heating and preheating of propellant
– Loose about 3-5% of the fission energy through escaping radiation
Common reflector materials
– Beryllium, graphite, Zirconium Carbide, Tungsten, Titanium, Aluminum
93
NUCLEAR FUEL COMPOUNDS
• Physical properties are key: most important metric is melting temperature
• Uranium has poor melting point (1405 K), but very high compounding stability
• Binary Compounds: Uranium + 1 other material
– Intermetallic
• Aluminum, beryllium, bismuth, copper, molybdenum, nickel, titanium
– Ceramic (5 > UBe13)
• UC, UC2, US, UN, UO2
• Ternary Compounds: Uranium + 2 other materials (Tmelt ~ 3560 K)
– U-Ta-C
– U-Nb-C
– U-Zr-C
94
FISSION CONTROL
95
NEW NASA MESSAGES
“A well thought and carefully designed NTP roadmap is needed” – Prof. Anghaie
•
NTP is well investigated technology, but development / improvement remains
– Heat transfer relations, geometries, materials, etc.
•
Fuel development and evaluation essential component of NTP program
– Testing at maximum temperatures, heat fluxes, transients, duration, re-start, etc.
•
Preliminary Research Programs are Beginning to Form
– Non-Nuclear development to gain knowledge base
– Design of experiment, data acquisition and analysis
– Handling H2 levels required for simulation at engine conditions
– Various expertise essential (materials, diagnostics, hot H2, etc.)
•
Partnering to facilitate development
– Confluence of NASA, industry (P&W), and academia
•
Hot H2 NTP experiments
– Support / design / build-up from academia
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RADIATION
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Radiation is of two forms (and emanate on two time scales)
– Beta Rays:
• Mass and charge of an electron
• Do not escape from core
– Gamma Rays:
• Non-charged particles without mass
• Tend to escape from core
Both types of radiation have prompt and delayed components
– Prompt radiations emanate instantly with fission
– Delayed radiations emanate over varying periods of time
REM=absorbed radiation dose x quality factor
Examples:
– Natural radioactive material in bones: 0.034 rem/year
– Chest x-ray: 0.01 rem
– 90-day space station mission: 16 rem
– Properly shielded nuclear reactor: 10 rem/year
Total Radiation Exposure Mission to Mars: NTP < Chemical Rocket
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REACTOR COOLANTS
• Over 100 types of reactor coolants can be used
– Ordinary gases, Water, organic liquids, liquid metals, molten metals, liquefied
salts, fluidized dusts, etc.
• For NTP: Reactor coolant becomes propellant
– Space-based applications, hydrogen is best for Isp
– Space-based reactors looking to operate just below melting point of materials
– In deep space-based system radioactive exhaust jettisoned from rocket
• H2 stored in liquid form and then converted to gas
• H2 is one of best moderating materials for slowing down neutrons and can also
serve as pre-core moderator/reflector and shield
• H2 does not participate in fission reaction nor does it have any direct contact with
fission fragments
Remember: Only one propellant needed – system complexity is greatly reduced
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APPROACH
• Non-nuclear testing in hot H2 environment key to engine development
– T=300-3200 K
– Realistic mass flow rates (0.8-1.5 g/s per cooling channel)
– Realistic inlet pressures (500 psi)
– Modular test section: investigate NERVA, particle bed, pebble bed, etc.
– Materials characterization and assessment of performance/stability in hot H2
– Safety, instrumentation, diagnostics, etc.
• Technological Archeology
– Many texts, reports, data sets, workshop reviews, etc.
– What other hardware and test apparatus available?
• ANL Nuclear Rocket Program, H2 test loops
• LUTCH (Russia), hot H2 test facility
• PET (Prototypical Element Tester) Grumman / Sverdrup ($3.5 M)
• Who should be involved?
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ANALYSIS APPROACH FOR EXPERIMENTAL DESIGN
“The Rover/NERVA engine is to be used as a “reference,” against which other concepts… will
be compared.” - Dr. Stanley K. Borowski, Nuclear Thermal Rocket Workshop (1990)
“Solid core has plenty of growth potential. Just because it's 1960's era technology doesn't
mean it's obsolete. Object of a new program should be to build on this and get it flying.”
“If you had kept on working [NERVA] you would now have a 4th generation system. It would
have Isp's over 1000, power densities 3000 MW, and maybe 30 hours of engine lifetime
with 180 stop & starts.” - Dr. James Dewar, AEC
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“Reference” Cases Considered:
– PHOEBUS-2A
• 5 GW Reactor: “Most powerful nuclear rocket reactor ever built”
– NERVA
• UC2/graphite, composite carbide/graphite, (U, Zr)C, UO2/T-111, UO2/W-Re,
UN/W-Re
– CERMET Type
• W matrix material
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H2 SAFETY AND WORST CASE SCENARIOS
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Test Cell Safety
– Worst case scenario: All H2 leaks into room and ignition source (overhead light is
turned on)?
• Room size: 50’ x 50’ x 30’
• H2 will diffuse to ceiling of room
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Chamber / Test Apparatus
– Worst case scenario 1: Optical window breaks completely (1 inch hole), at the same
time, tube shatters and H2 is filling into the chamber at 1.5 g/s
• Chamber initially at vacuum and ambient air begins to fill
• Detonation limit with air ~ 6 psi, so have time for emergency shut-off
• Flammability limit ~ 0.1 psi, so might have time for emergency shut-off
• At equilibrium, H2/Air ratio ~ 2 % at 1 ATM, non-flammable
– Worst case scenario 2: Optical window small crack, such that ideal ratio of H2/Air
enters chamber
• 20 < H2/Air < 60 and above 6 psi, detonation region
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ANALYSES IN PROGRESS
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Transient Analysis: Experiment Start-Up
– Ensure that coolant gas is flowing
prior to heating
– Emergency Shut-Off System: Cuts
Heating  Cuts H2  Purge
Flow Perturbation
– Step change 50% reduction in mass
flow for Dt 0.5 ms
Simulate: Cracked element, blockage, back-flow response time, turbo-pump transients, etc.
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INDUCTIVE HEATING
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Induction heating is a method of heating conductive material by subjecting it to an
alternating electromagnetic field
– MSFC: Power 12.5 kW (currently in lab) 100 kW also available, 150 < f < 350 kHz
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Basic Operation
– Magnetic field induces eddy currents in piece, and electrical resistance of piece to
flow of current causes piece to heat up
– Depth at which current flows is dependant on frequency of magnetic field
• Higher frequencies: shallow depth, lower frequencies: deeper depth
– Depth of heating is function of electrical frequency applied, heating time and power
density applied to component being heated
– Coil ID ~ 1.5 inches
• Depends on internal configuration of power supply
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ANALYSES IN PROGRESS
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Inductive Coil Design
– Working with Dr. Dahake (Ameritherm) and Dr. Rudnev (Inductoheat Group)
– Example: Heating a Titanium Rod to 1200 K, 13 turn coil over 5 inches
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NTP Materials: Behavior of UyZr1-yC1-x (Fuel) and ZrC1-x (Coating)
– Considered as fuel and coating in several designs of NTP reactors (Tmelt ~ 3693 K)
– Vaporization rate and melting point prediction in presence of flowing hot H2
– If flowing H2 changes C content
– Change in Tmelt with presence of U (~400 K per 10%)
– Major influences: component vapor pressure, diffusion constants of C and U,
equilibrium pressure of various hydrocarbon reaction products, degree to which
equilibrium with various gases is achieved
– Question: Based on mass loss rate and degradation, is there an optimum mass flow
rate, i.e., is it better to have more small channels or fewer larger ones?
• Question can be answered from a flow physics point of view also
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Optimization of T/W vs. Isp for low pressure operation
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EXPERIMENTAL DESIGN CONSIDERATIONS
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Test Duration
– Preliminary tests
• H2 ~ 11 min/bottle
– For a 4 hour duration test at 1.5 g/s using 2400 psi (11 inch x 61 inch) bottles (BOC):
• H2: 23, He: 12, N2: 2, Ar: 2 bottles required
• Measurement of mass loss, volume change, grain growth, etc.
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Preliminary Materials Selection / Specimen Quotes
– Tantalum:
• Metal Technology, Inc.
– $747/piece, OD: 0.165 inch, ID: 0.125 inch, 6 ft. long
– Seamless, as drawn condition, un-annealed, ASTM B521
• The Rembar Company, Inc.
– $810/piece, OD: 0.187 inch, ID: 0.202 inch, 4 ft. long
– Seamless, annealed
– Rhenium
• ULTRAMET
– $ 6,172/piece, OD: 0.155 inch, ID: 0.125 inch, 4.5 ft. long
– CVD deposited on graphite mandrel, ID roughness 0.0005-0.001 inch
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FUTURE WORK: ANALYSES
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Low Pressure Operation
– Radial Out-Flow Core (Maximize Flow)
– “One of the advantages of low pressure is that the heat flux is greatly reduced. It is
about a factor of 50:1 less than the high pressure NERVA engine.”
Plug Nozzle
– “A nozzle concept already exists which, philosophically in retrospect, appears as
though it were predestined for nuclear rockets.” Crouch, 252
– “Some of the things that have been rejected in the chemical engines, such as
expansion-deflection nozzles, spike nozzles, and plug nozzles, all become candidates
for reexamination to see what would be the optimum way to design a thrust
chamber/nozzle for hydrogen recombination.” (Ramsthaler, 127-150).
Potential for tailoring of flow path cross-sectional area
– Minimum area located at maximum heat transfer locations
– Minimize potential heat transfer hot-spots
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ANALYSES IN PROGRESS
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Reactor Power Profile Optimization
– “From a nuclear rocket design standpoint, a flat power profile may not be best
configuration and that an optimum power profile probably exists that gives that lowest
fuel temperatures for a given core and operating condition.”
– Constrained to half-cosine function
• Maximum propellant exit temperature exists when zero at each end point
– Constrained to half-cosine function, but allow for skew
– Exponential power distribution
• Determine exponential shape factor for maximization of coolant exit temperature
for a given reactor length
• Given a maximum allowable reactor length, determine shape profile to maximize
coolant exit temperature
Mixing Model
– Mixed out flow temperature for a given radial profile and number of elements
– Compound flow, baroclinic vorticity generation, mixing time scale
– Include “void flow” for support structures
– Sample result: Tmix ~ 300 K lower than (Texit)max, 10% Isp 
Optimization of T/W vs. Isp for low pressure operation
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FUTURE WORK: EXPERIMENTAL DEVELOPMENT
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How to vent H2 exhaust gas
– 43 lb/hr of H2 without burn-stack run at MSFC
• (43 lb/hr)*(1 kg/2.2 lb)*(1 hr/3600 s)=0.00543 kg/s=5.43 g/s
• Proposed NTP experiment H2 mass flow ~ 1-2 g/s
Burst calculation for tube samples at pressure and T
– Ductility of sample probably increases with T
– Hydrostatic validation test (calculation suggests OK with SF of 5)
Upgrade thermal H2 model to contain NTP carbide fuel mass loss mechanisms
– Carbon loss by H2 corrosion (T~1500 K)
– Corrosion inhibited by hydrocarbon in gas stream (T~2800 K)
– Mass loss by vaporization (T~3000 K)
Model axially segmented core sections, with differing power distributions
– May mitigate mid-band issues resulting from fabrication technique
How does Acs change with time due to H2 erosion, what is impact on hg
Surface roughness for pressure loss management, friction flow factor
“Investigate laminar flow instability at NERVA maximum and minimum core inlet
temperature during cooldown.”, P.116 [low pressure applications]
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