GOING TO MARS WITH NUCLEAR THERMAL PROPULSION

1
GOING TO MARS WITHNUCLEAR 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

2
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)

3
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

4
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

5
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

6
CHEMICAL (THERMAL) ROCKETS ENERGY LIMITED
7
HOW A CHEMICAL ROCKET WORKS
F
Chemical Energy

• Rocket Propulsion (class of jet propulsion) that
  produces thrust by ejecting stored matter
• Propellants are combined in a combustion chamber
  where chemically react to form high TP 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)?

Thermal Energy
Kinetic Energy
8
ELECTRIC ROCKETS POWER LIMITED
9
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."

10
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 lt 0.1 due to that
  of light pressure)
• Do not convert to electricity like solar cells

11
FUTURE OF SPACE PROPULSION (?)
12
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 Earths B field
• Solar power drives ionospheric current, tether
  reboost without using propellant

13
WHAT IS ENERGY LIMIT?

• The basic secret of space travel and extending
  human presence throughout heliocentric space is
  energy, immense quantities of energy

14
ANTIMATTER PROPULSION

• 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

1/10th gram antimatter
15
EVERYTHING YOU NEED TO KNOW TO BE A ROCKET
SCIENTIST

• Thrust, N
• How much Force?
• T/W is key metric for launch vehicles
• Less important with space exploration
  applications
• 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
• Ideal Rocket Equation, m/s (1 form of many)
• When and How Fast?
• Can a rocket travel to a speed faster than speed
  at which exhaust leaves rocket?

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PERFORMANCE COMPARISON
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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

18
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)

19
BACKGROUND REVIEW OF PROGRAMS (1955-PRESENT)

• 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)

20
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

21
COMPARISON OF REACTORS TESTED IN ROVER/NERVA
PROGRAM
22
PICTURES FROM ROVER/NERVA TESTING
23
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

24
CORE DETAILS AND MAJOR COMPONENTS
25
MORE CORE DETAILS KIWI 4B
26
NTP BASIC OPERATION
27
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

28
THERMAL GENERATION MECHANISM NUCLEAR FISSION
Various coolants may be used H2O, Ar, He, liquid
metals, H2
Total Radiation Exposure Mission to Mars NTP lt
Chemical Rocket
29
IF NTP SO GOOD, WHY HASNT 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.

30
IF ITS 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..."

31
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 (PW), and academia
• Hot H2 NTP experiments at MSFC
• Support / design / build-up from academia

32
APPROACH

• Non-nuclear testing in hot H2 environment key to
  engine development
• T300-3200 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
33
RANGE OF NTP INPUT DATA
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
34
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
35
H2 COOLANT / PROPELLANT COMMENTS

• Range of interest
• T300-3000 K, P14.7-1000 psi
• Important to model H2 properties accurately
• Up to 1500 K, pressure has little effect on Cp,
  g, m, k
• For T gt 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
• P1 ATM, T3000 K, NH15 vs. P40 ATM, T3000
  K, NH2.6
• Isp improvement with dissociation, but no impact
  on cf
• Ionization not relevant at these temperatures
• P1 ATM, T3000 K, NH O(10-11) vs. P40 ATM,
  T3000 K, NH O(10-12)
• Compressibility small effect
• P1 ATM, T3000 K, Z8.2 vs. P40 ATM, T3000
  K, Z1.3

36
H2 DATABASE IMPORTANT TO CAPTURE T P
37
H2 COMMENTS UNIQUE BEHAVIOR
P1 ATM, T2600 K NH3.9 , NH1.3x10-13
P40 ATM, T2600 K NH0.6 , NH8.8x10-15
Effects of dissociation and ionization on cp, k
are dramatic Higher pressures ? dissociation
suppressed NTP nominal range of operation T lt
3500 K and P 20-40 ATM
38
H2 COMMENTS VACUUM SPECIFIC IMPULSE
Phoebus 2A, Ispvac918 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 ?
39
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
40
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 gt 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

41
REACTOR POWER DISTRIBUTION
42
REACTOR TEMPERATURE DISTRIBUTION
43
TEMPERATURE DISTRIBUTION COMMENTS

• Note that Tbulk maximum at L100
• 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 mid-band 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

44
HEAT TRANSFER COEFFICIENT VARIOUS FORMS

• 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.

45
HEAT TRANSFER COEFFICIENTS
46
HEAT TRANSFER EXPERIMENTAL SCALING

• Convective coefficient scales with diameter as
  hg1/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
  hgr0.8p0.8
• Halving inlet pressure will reduce coefficient by
  57
• Lighter gases lead to higher heat fluxes,
  hg1/M0.6
• Ratio of molecular weights of ArH2 20, heat
  flux for ArH2 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

47
SAMPLE MODEL OUTPUTSFLOW VARIABLES VS. AXIAL
LOCATION
48
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

49
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.

50
SURROGATE TEST GASES He, N2, Ar
H2
He
N2
Ar
51
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 T3000 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

52
PRELIMINARY EXPERIMENTAL CONFIGURATION

• Initial Test Chamber
• 77.6 inch (Full scale test article L55 inches),
  8.25 inch OD Chamber
• 16 ports already in place, D1.38 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 lt
  f lt 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

53
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

54
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 (PW), and academia
  (FIT, UF)
• Hot H2 NTP experiment at MSFC
• Support / design / build-up from academia

55
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

56
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
57
BREAKTHROUGH IDEA 2 ?

• Chemical rocket propulsion system benefit from
  scaling
• TA, WV, T/W1/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

58
SUPPLEMENTAL SLIDES
59
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)
60
ROCKETS ENERGY VS. POWER LIMITED

• Chemical Rockets are Energy Limited
• Unit of Energy JOULE, EnergyFDisplacement kg
  m/s2mkg 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 FVkg
  m/s2m/skg 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

61
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.)
• Solid Rockets, Shuttle SRB
• Fuel Aluminum and Nitrate
• Advantages
• Simple, low cost, safe
• Disadvantages
• Thrust cannot be controlled, no shut down

Liquid-Propellant Rocket Engine 11D33
62
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

63
EXAMPLE ATLAS IIAS

• 47 m tall, 3-4 m diameter, 234,000 kg
• Lockheed-Martin 2-stage liquid propellant
  (LOX-RP1) booster
• 95-105M per launch
• First stage booster section
• 2 Rocketdyne engines
• 1.84 MN thrust, Isp263 seconds
• runs about 3 minutes
• Second stage is sustainer section
• 1 Rocketdyne engine
• 269 KN thrust, Isp220 seconds
• 5 minutes burn with booster
• Strap-on solid rockets
• Four Thiokol Castor IVA SRMs
• 433 KN thrust, Isp229 seconds
• 9 m tall, 1 m diameter, 12K kg
• Burn about 56 seconds
• Uses a Centaur upper stage
• 2 Pratt Whitney engines
• LOX-LH2

64
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

65
EXAMPLE TITAN

• Titan is a family of expendable rockets.
• Most Titans 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

66
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, 70003 a.m, EST.
• QUESTION
• How many rockets systems on STS?
• For more on STS
• http//users.commkey.net/Braeunig/space/specs.htm

67
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

68
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
  3-stage 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

69
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)

70
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

71
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

72
FUTURE LAUNCH VEHICLES?
Linear Aerospike Engine
Taurus
Minotaur
More Information http//www.spaceandtech.com/spac
edata/rlvs/venturestar_sum.shtml http//www.orbita
l.com/LaunchVehicle/SpaceLaunchVehicles/index.html
http//www.aerospaceweb.org/design/aerospike/main
.shtml
Pegasus
73
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

74
EXAMPLE ELECTRIC AND ION THRUSTERS

• 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/Elec
  tricPropulsion/
• http//www.marsacademy.com/propul/propul7.htm
• http//richard.hofer.com/electric_propulsion.html
• http//www.stanford.edu/group/pdl/EP/EP.html

Designer Rocketdyne. Developed in 1999.
Propellants Electric/Xenon Thrust (vac) 0.001
N Isp 3,500 s
No Combustion
75
EXAMPLE ELECTRIC AND ION THRUSTERS

• 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/Elec
  tricPropulsion/
• http//www.marsacademy.com/propul/propul7.htm
• http//richard.hofer.com/electric_propulsion.html
• http//www.stanford.edu/group/pdl/EP/EP.html

Designer Rocketdyne. Developed in 1999.
Propellants Electric/Xenon Thrust (vac) 0.001
N Isp 3,500 s
No Combustion
76
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

77
EXAMPLES OF NUCLEAR PROPULSION
No Combustion In These Devices
78
ROCKET SELECTION GUIDE

• MISSION REQUIREMENT
• Non-Space Missions
• Atmospheric / Ionospheric Sounding
• Tactical Missiles
• Medium-Long Range Missiles
• Launch to Space
• Impulsive DV in Space
• Time critical maneuvers
• Energy change from elliptic orbits, plane change
  from elliptic orbits
• Non-fuel limited situations
• Low Thrust DV in Space
• Mass-limited missions
• Non-time critical missions
• Small, continuous orbit corrections, near
  circular orbits

• ROCKET TYPE
• Solid Propellant, 1-4 stages
• Solid Propellant, 1-2 stages
• Solid or Liquid Propellant, 2-3 stages (very high
  acceleration)
• 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

79
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

80
PERFORMANCE MEASURES THRUST

• For our simple rocket we had (PePa)
• 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?

81
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
• ge is measured on the earths surface, ge9.8
  m/s2
• Some Example Numbers
• Chemical rocket range 200-500 s (500 is just
  about the limit)
• Shuttle Main Engine 455 s (T1670 kN each), SRB
  250 s (T14700 kN each)
• Nuclear Thermal 800-1200 s
• Trade-off vs. mass for EP, 500-6000 s
• Nuclear Electric Rocket 20,000 s (T/W0.0001)
• Specific impulse improves with LOW molecular
  weight, LOW specific heat ratio, and HIGH
  temperature

82
ROCKET VS. TURBOJET ISP
Isp
Mach Number
83
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
• Liquid, chemical rocket development
• Solid-fuel nuclear reactors

84
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).

85
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.

86
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, (D-NV) Harold B. Finger,
  Manager of the Space Nuclear Propulsion Office
  and Dr. Alvin C. Graves, Director of test
  activities for the Los Alamos Scientific

87
PERFORMANCE COMPARISON
88
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)
89
MOMENTUM EXCHANGE TETHERS
90
H2 COMMENTS VARIATION IN DATA SETS
NASA SP Reports Kubin and Presley McCarty Patch
Curve Fits Hill and Peterson CHEMKIN
Incropera and De Witt
91
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

92
RADIATION DOSSAGE (rem)
remabsorbed radiation dose x quality factor
93
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

94
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 gt UBe13)
• UC, UC2, US, UN, UO2
• Ternary Compounds Uranium 2 other materials
  (Tmelt 3560 K)
• U-Ta-C
• U-Nb-C
• U-Zr-C

95
FISSION CONTROL
96
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 (PW), and academia
• Hot H2 NTP experiments
• Support / design / build-up from academia

97
RADIATION

• 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
• REMabsorbed 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 lt
Chemical Rocket
98
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

99
APPROACH

• Non-nuclear testing in hot H2 environment key to
  engine development
• T300-3200 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?

100
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 workin