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## Lesson 14: Nuclear Propulsion Design

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**Lesson 14: 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**Propellant Selection**• For high Isp, low molecular weight propellants should be used • Molecular hydrogen • Methane • Water • Temperature variation of g can be found by • where Cp = Cp (T) [Thermally perfect gas assumption]**Propellant Selection**• For H2: • Knowledge of the thermochemistry lends itself to the estimation of c*, ve, and Isp. [T in Kelvin]**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 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: • 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 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 • The fission of a single Uranium-235 atom produces 200 MeV or 3.206x10-11 J of energy**Size the Reactor**• The mass of Uranium consumed is • The mass density of U-235 is 19,100 kg/m3**Size the Reactor**• Combining the equations yields • 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**Reactor Mass**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 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**• 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**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 BoosterGross 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/LH2No Engines: 4. Status: Study 1959. • Stage Number: 1. 1 x Hyperion SustainerGross 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/LH2No Engines: 2. Nerva 12 GWStatus: 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-670Gross 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+AlcoholNo Engines: 1. YaRD OKB-670Status: 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 - AGross 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/KeroseneNo Engines: 24. NK-15Status: 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 AGross 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/LH2No Engines: 40. YaRD Type AStatus: 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 RITA-B would be refuelled in low earth orbit by multiple launches of other RITA-B'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 CGross 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/LH2No Engines: 4. NERVA/Lox Mixed CycleStatus: 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