Transcript Lesson 14: Nuclear Propulsion Design

Nuclear Propulsion Design
Dr. Andrew Ketsdever
Design Process
• The design process is driven by mission
requirements (all propulsion systems)
– Payload mass
– Mission ∆V
– Operational environment
• Key differences for nuclear propulsion
system from the design of a liquid rocket
– Design of the reactor
– Design of the radiation shield
Design Process Steps
• Preliminary Design Decisions
– Evaluate propellants
• Thermochemistry
• Performance
– Size the overall system
• Propellant mass
• Total mass
– Determine the reactor power required
• From required thrust level, the mass flow rate can be
determined
• Drives the amount of heat (energy) needed to be delivered to
propellant
• Used to estimate the power required for the reactor
Design Process Steps
• Size the Reactor
– Determine appropriate
reactor type
• NERVA
• PBR
• CERMET
– Determine the core
dimensions
– Determine the
reactor mass
Design Process Steps
• Design the Other Systems
– Thrust (stagnation) chamber
• Efficient heat transfer
• No mixing or vaporizing of propellants required
– Propellant feed system
• Pressurization or turbopump fed
• Flow control
– Valves, regulators, feedlines
– Radiation shield
• Protect sensitive payloads
– Support structure
Design Process Steps
• Evaluate the Design
– Performance
– Mass
– Safety
– ITERATE
Thermochemistry Evaluation
• Much easier than liquid or solid
analysis since no combustion is
taking place
• The nuclear reactor is simply
used to heat a propellant gas
– Thermodynamics
– Gas dynamics
• Performance driven by
– Maximum achievable temperature
– Heat transfer efficiency
– Propellant selected
To
Isp ~
mw
Reactors
NERVA
PBR
CERMET
Power (MW)
1570
1945
2000
Max
Propellant
Temp(K)
Isp (s)
2361
3200
2507
825
971
930
Po (MPa)
3.102
6.893
4.136
Propellant
H2
H2
H2
Propellant Selection
• For high Isp, low molecular weight propellants
should be used
– Molecular hydrogen
– Methane
– Water
• Temperature variation of g can be found by
Cp
g

Cp 
mw
– where Cp = Cp (T) [Thermally perfect gas assumption]
Propellant Selection
• For H2:
1 
1165
0.75
1.5 
Cp 
 560.7T 
56.505 702.74T 
m w
T

T
T 
100
[T in Kelvin]
• Knowledge of the thermochemistry lends itself to
the estimation of c*, ve, and Isp.
Required Reactor Power
• The heat added to the propellant mass flow is
equal to the steady-state power generated by the
reactor
– Assuming no heat losses
• This power is used to increase the enthalpy of the
flow
Pcore  m hv   C p dT   mP
T1


T2
where hv is the enthalpy required to vaporize the liquid propellant and P is the
reactor specific power (W s/kg)
Reactor Power
• The specific reactor power is dependent on the
maximum propellant temperature
• For H2:
P  0.018061T  5.715417
– Assumes a linear approximation of empirical results
System Pressure Levels
• DPlosses include a major contribution from
the pressure drop through the reactor core
– NERVA: 11-38% of Po
– PBR: 5% of Po
– CERMET: 10-54% of Po
• Regenerative cooling (nozzle) losses of
20-30% are also possible depending on
the configuration
Size the Reactor
• NERVA-Type
– Short burn durations where loss of Uranium fuel due
to fission is not a major factor
Vcore
Pcore

PD
where PD is the power density of the reactor (W/m3)
Size the Reactor
• NERVA-Type
– To maintain the required power level over the entire
burn (long duration burns), more fuel must be added to
account for the Uranium lost by fission
Ecore  Pcoretburn
– The fission of a single Uranium-235 atom produces 200
MeV or 3.206x10-11 J of energy
N consumed
Ecore

11
3.206 x10 J
Size the Reactor
– The mass of Uranium consumed is
0.235 N cons
m consumed 
0.6023 x10 24
– The mass density of U-235 is 19,100 kg/m3
Vconsumed
mcons

19100kg
m3
Size the Reactor
– Combining the equations yields
Vcore

1
19
 Pcore 6.4 x10 tburn  
PD 

– As fuel is lost, the control rods are repositioned to
allow a constant reaction rate throughout the burn
– Critical mass is required to achieve a self-sustaining
fission chain reaction
keff ≥ 1
Size the Reactor
• NERVA-Type
– Diffusion theory adequately describes the
reactor physics
Size the Reactor
• NERVA-Type
Size the Reactor
• PBR (7, 19, 37 fuel element designs)
Size the Reactor
• CERMET
Size the Reactor
Reactor Mass
Reactor Type
Core Density (kg/m3)
NERVA
2300
PBR
1600
CERMET
8500
Preliminary design assuming a constant reactor core density.
Will give a mass estimate at least 90% accurate.
Estimates are for all components of the core.
Reactor Mass
Reactor Mass
Power Requirement
Less than 250 MW
Reactor Choice Based
Purely on Reactor Mass
PBR (7 element)
250 – 300 MW
CERMET
300-750 MW
PBR (19 element)
Greater than 750 MW
PBR (37 elements)
Reactor Shielding
• Nuclear reactor requires a radiation shield
to protect the payload from the damaging
effects of radiation
• REM = absorbed dose x quality factor
– Absorbed dose – rad
– 1 rad = amount of radiation caused by 100
ergs/gram of energy absorption
– Quality factor takes into account that different
forms of radiation cause different effects.
Radiation Shielding
Scenario
Radiation
Aircraft flight at 9 km
0.001 rem/hr
Chest X-Ray
0.01 rem
Living in Houston
0.25 rem/yr
Living in Denver
0.35 rem/yr
90-day ISS Mission
16 rem
NTP System
10 rem/yr
Single solar flare of
moderate energy
303 rem/yr
Radiation Shielding
Criteria
General Public
Astronauts
30-day limit
0.04 rem
Annual limit
0.50 rem
Career limit
NA
150 rem
(vomiting/hair
loss)
300 rem
(radiation
sickness)
600 rem
Radiation Shielding
• Many factors influence the geometry,
composition and mass of the radiation
shield
– Total allowed levels of radiation
– Length of the mission
– Type of radiation
– Size and nature of the reactor
– Spacecraft configuration
Radiation Shielding
• Each form of radiation is best attenuated
by different types of materials
– Combining different materials can make a
very effective radiation shield
– Beryllium (Be), Tungsten (W), Lithium Hydride
(LiH) sheild has been shown to be effective
– Used as a baseline shield for this course
•
•
•
•
18 cm Be, 5 cm W, 5 cm LiH
Be acts as a neutron reflector
W acts to attenuate gamma rays
LiH attenuates remaining neutron flux
Radiation Shielding
• Preliminary sizing
– Radius of shield is equal to radius of the core
– Use baseline radiation shield material
– Mass of 3500 kg/m2
• Radiation effects in materials
– Photoelectric Effect
• Material emits electrons due to energetic photon interactions (E>0.3
MeV)
– Compton Scattering
• Photon transfers energy to charged particles in material (E ~ 0.3 to
10 MeV)
– Pair Production
• Gamma ray completely absorbed and produces an electron-positron
pair (E>10 MeV)
– Bremstrahlung Radiation
• Slowing electron produces high energy X-Rays
Radiation Interaction With Matter
Nuclear Thermal Propulsion
Systems
• Nuclear Thermal Propulsion Systems have been proposed since the
late 1940’s.
• At the suggestion of Theodore von Kármán and following a request of
Gen. H. B. Thatcher, an Ad Hoc Committee of the Scientific Advisory
Board met in the Pentagon to consider the application of nuclear
energy to missile propulsion.
• In its report, the Committee "noted that there was an almost complete
hiatus in the study of the nuclear rocket from 1947 following a report
by North American Aviation, until a 1953 report by the Oak Ridge
National Laboratory.
• Because the technical problems appear so severe, and because
another 6 years of no progress in this area would seem to be
unfortunate," the Committee felt that a continuing study both analytical
and experimental, at a modest level of effort, should be carried on.
• NO NTP SYSTEMS HAVE BEEN FLOWN TO DATE
Hyperion - USA
•
•
•
•
Hyperion was considered in 1958 as a ca. 1970 Saturn follow-on. It used a
small jettisonable chemical booster stage that contained chemical engines
and the LOX oxidizer for the conventional engines. This booster stage
surrounded the nuclear core vehicle with its large liquid hydrogen tank. The
conventional stage would draw fuel from the main hydrogen tank until
burnout. Hyperion would have doubled the translunar trajectory
performance of the Saturn V and less than one third of the liftoff mass.
Manufacturer: Convair. LEO Payload: 145,000 kg. to: 485 km Orbit. at: 28.0
degrees. Payload: 82,000 kg. to a: parabolic escape trajectory. Apogee: 36
km. Liftoff Thrust: 1,090,000 kgf. Liftoff Thrust: 10,700.00 kN. Total Mass:
850,000 kg. Core Diameter: 8.54 m. Total Length: 85.40 m.
Stage Number: 0. 1 x Hyperion Booster Gross Mass: 394,625 kg. Empty
Mass: 18,144 kg. Thrust (vac): 1,400,000 kgf. Isp: 457 sec. Burn time: 70
sec. Isp(sl): 365 sec. Diameter: 8.54 m. Span: 13.00 m. Length: 12.00 m.
Propellants: Lox/LH2 No Engines: 4. Status: Study 1959.
Stage Number: 1. 1 x Hyperion Sustainer Gross Mass: 453,592 kg. Empty
Mass: 110,000 kg. Thrust (vac): 589,670 kgf. Isp: 800 sec. Burn time: 460
sec. Diameter: 8.54 m. Span: 8.54 m. Length: 51.00 m. Propellants:
Nuclear/LH2 No Engines: 2. Nerva 12 GW Status: Study 1959.
Orion - USA
• The final iteration of the Orion design was a
nuclear pulse propulsion module launched
into earth orbit by a Saturn V. The 100 tonne
unit would have had a diameter of 10 m to
match that of the booster. This would limit
specific impulse to 1800 to 2500 seconds,
still two to three times that of a nuclear
thermal system.
• A second launch would put a 100 tonne Mars
spacecraft with a crew of eight into orbit.
After rendezvous and checkout, the
combined 200 tonne spacecraft would set out
on a round trip to the Mars - total mission
duration as little as 125 days!.
• Manufacturer: General Atomic. Payload:
100,000 kg. to a: Mars and back trajectory.
Total Mass: 100,000 kg. Core Diameter:
10.00 m. Total Length: 50.00 m.
YaRD ICBM - USSR
• A 30 June 1958 resolution authorised development of this
astounding weapon, and the draft project was completed on 30
December 1959. Perhaps coming under the heading of
'inadvisable rocket science', test launches would have been
into an artificial reservoir in the target area to limit
contamination by having the reactor crash into water at the end
of its trajectory. Interestingly American spy Penkovskiy reported
development of this rocket in 1962, but the story was not
believed. Only in 1996 was the program revealed.
• Manufacturer: Korolev. Liftoff Thrust: 128,000 kgf. Total Mass:
84,400 kg. Core Diameter: 3.33 m. Total Length: 25.00 m.
• Stage Number: 1. 1 x YaRD ICBM OKB-670 Gross Mass:
96,000 kg. Empty Mass: 8,800 kg. Thrust (vac): 170,000 kgf.
Isp: 470 sec. Burn time: 235 sec. Isp(sl): 430 sec. Diameter:
3.33 m. Span: 3.33 m. Length: 23.00 m. Propellants:
Nuclear/Ammonia+Alcohol No Engines: 1. YaRD OKB-670
Status: Study 1959. Comments: Nuclear-propelled ICBM with
engines in development by Bondayuk. Four expansion nozzles
fed by single reactor. Payload 4,000 kg to 14,000 km. Empty
mass, vehicle length calculated.
N1 - USSR
•
•
•
•
•
For a Mars expedition, it was calculated that the AF engine would deliver 40% more
payload than a chemical stage. But Korolev’s study also effectively killed the program
by noting that his favoured solution, a nuclear electric ion engine, would deliver 70%
more payload than the Lox/LH2 stage.
Further investigation of nuclear thermal stages for the N1 does not seem to have been
pursued.
Manufacturer: Korolev. LEO Payload: 270,000 kg. to: 220 km Orbit. at: 51.6 degrees.
Payload: 24,600 kg. to a: lunar surface trajectory. Liftoff Thrust: 3,600,000 kgf. Liftoff
Thrust: 35,000.00 kN. Total Mass: 2,400,000 kg. Core Diameter: 17.00 m. Total Length:
180.00 m.
Stage Number: 1. 1 x N1 1962 - A Gross Mass: 1,384,000 kg. Empty Mass: 117,000 kg.
Thrust (vac): 4,020,000 kgf. Isp: 331 sec. Burn time: 103 sec. Isp(sl): 296 sec.
Diameter: 10.00 m. Span: 17.00 m. Length: 30.00 m. Propellants: Lox/Kerosene No
Engines: 24. NK-15 Status: Study 1962. Comments: Includes 14,000 kg for Stage 1-2
interstage and payload fairing. Compared to total fuelled mass excludes 15,000 kg
propellant expended in thrust build-up and boil-off prior to liftoff. Values as in draft
project as defended on 2-16 July 1962.
Stage Number: 2. 1 x N1 Nuclear A Gross Mass: 700,000 kg. Empty Mass: 250,000 kg.
Thrust (vac): 700,000 kgf. Isp: 900 sec. Burn time: 570 sec. Diameter: 12.00 m. Span:
12.00 m. Length: 90.00 m. Propellants: Nuclear/LH2 No Engines: 40. YaRD Type A
Status: Study 1963. Comments: N1 nuclear upper stage study, 1963. Figures calculated
based on given total stage thrust, specific impulse, engine mass.
RITA - USA
• Nuclear single-stage-to-orbit booster.
• Engineers at Douglas proposed this nuclear single-stage-to-orbit launch
vehicle in February 1961.
• The RITA-A (Nexus) launch vehicle would be equipped with a 91,000 kgf
nuclear engine with a specific impulse of 850 seconds. It could take 7300 kg
to low earth orbit. Use of a Saturn S-IB first stage would allow it to take
38,200 kg to low earth orbit or 13,600 kg to lunar orbit. First test flight would
come as early as 1965.
• The RITA-B follow-on vehicle would be equipped with 4 x 200 tonne thrust
nuclear engines with a specific impulse of 950 seconds. The RITA-B would
perform a manned Mars mission, with departure in late 1967. A single RITAB would be refuelled in low earth orbit by multiple launches of other RITAB's. It would then depart for a direct flight to the Martian surface and return.
The vehicle could use body lift to produce an L/D ratio of 0.7 for aerobraking
to the Martian surface and on return to earth. With a total delta-V capability
of 20.7 km/sec, RITA-B was capable of making the flight to Mars in 225
days.
RITA - USA
• RITA-C would be a shuttle version of the design, for transporting one
million pounds of payload to low earth orbit.
• Manufacturer: Douglas. LEO Payload: 454,500 kg. to: 325 km Orbit.
Liftoff Thrust: 6,074,691 kgf. Liftoff Thrust: 59,572.37 kN. Total Mass:
4,399,000 kg. Core Diameter: 21.30 m. Total Length: 60.00 m.
Stage Data - RITA C
Stage Number: 1. 1 x RITA C Gross Mass:
4,399,000 kg. Empty Mass: 880,000 kg.
Thrust (vac): 9,841,000 kgf. Isp: 810 sec.
Burn time: 285 sec. Isp(sl): 500 sec.
Diameter: 21.30 m. Span: 57.90 m. Length:
50.30 m. Propellants: Nuclear/LH2 No
Engines: 4. NERVA/Lox Mixed Cycle Status:
Study 1963. Comments: Same engine
chamber used to burn liquid oxygen and
hydrogen for boost phase, switching to pure
nuclear thermal engine for high-performance
final acceleration.
STCAEM - USA
•
•
STCAEM (Space Transfer Concepts and Analyses for Exploration Missions) was a
major NASA funded study produced by Boeing in 1991. It provided an exhaustive trade
analysis of mission profiles and trajectories for manned Mars missions using four
different propulsion technologies (cryogenic chemical with aerobraking, nuclear
thermal, nuclear electric, and solar electric). Within each study alternate mission
profiles using split/sprint missions, flyby rendezvous, and additional aerobraking were
examined. Only the baseline for the nuclear thermal mission is presented here.
The total space vehicle mass in low earth orbit was 673,475 kg with the mass
breakdown was as follows:
–
–
–
–
–
–
–
–
–
–
Habitat module, 34,939 kg, consisting of empty mass, 28,531 kg; 5,408 kg consumables and
1,000 kg of experimental equipment
MEV 73,118 kg
MTV spaceframe, NTR engine systems, and radiation shield: 12,086 kg
Trans-Mars injection propellant: 262,100 kg
Trans-Mars injection tanks: 39,973 kg
Mars orbit capture propellant: 138,800 kg
Mars orbit capture tanks: 24,296 kg
Trans-Earth injection propellant: 51,727 kg
Earth orbit capture propellant: 24,296 kg
EOC/TEI common tank: 23,962 kg