A new propulsion concept called the Nuclear Thermal Turbo Rocket (NTTR) is proposed for Earth to Orbit applications. The NTTR utilizes a nuclear fission reactor to thermally heat hydrogen propellant into a rocket plenum. The rocket nozzles are located at the tips of a variable pitch thrust fan connected to the plenum by passages in the fan blades, and each nozzle is a linear aerospike on the trailing edge of the blade. The thrust fan is located in a duct such that the heated hydrogen propellant is combusted with ambient sourced oxygen to augment the rocket thrust. The fan is of variable pitch to provide maximum thrust for varying inlet velocity. The duct has a variable geometry inlet, able to provide appropriate mass flow and compression to the combustor throughout the trajectory, and a variable geometry outlet to provide appropriate nozzle area for maximum thrust. The rocket nozzles act as propellant injectors during the airbreathing portion and pure rockets during low atmospheric density portions, with the NTTR utilizing a single gas path from launch to orbital velocity. The propulsion concept is of high performance and is able to transport more than 50% mass fraction in a Single Stage to Orbit (SSTO) via an air-breathing rocket trajectory with intended complete reusability. Payload fractions of up to 19% are predicted (inert mass includes reactor radiation shielding) due to a mission average Specific Impulse (Isp) of 1,662 seconds.
Examining the shortcomings of prior concepts and systematically integrating major features of existing technical work is the innovation pathway. What is desired is a propulsion system capable of SSTO for highest reusability simultaneously with highest mass fraction delivered for lowest operational costs so as to minimize access to orbit costs. Nuclear thermal rockets already offer the highest Isp of launch-capable pure rocket propulsion systems, whereas the Supercharged Ejector Scramjet (SESJ, Ref. 8) and afterburning supersonic Rocket Fan are the highest launch to hypersonic Isp chemical combined cycle systems proposed so it is logical to attempt to integrate those cycles. The Rocket Fan study predates a later study where the performance of Supersonic Through-Flow Fans (STFF) was expanded and detailed by adding variable pitch fan capability with enhanced low speed thrust and reduced mass as compared to Airturbo Ramjet solutions for Mach 5 cruise applications. RBCC solutions typically use the supercharging fan only for subsonic or low supersonic augmentation, whereas STFFs can operate with fan face Mach numbers from 0 to around 4.5. Air-augmented nuclear rockets are only mentioned once in the literature – the Nuclear Air-Enhanced Rocket (NEAR) which is a ducted rocket that performed as an ejector rocket and ramjet.
The new proposed propulsion concept is called the Nuclear Thermal Turbo Rocket (NTTR), which is a supercharged air-augmented nuclear thermal rocket architecture. It operates in rocket fan, ramjet, scramjet and pure rocket modes.
Previous Nuclear Thermal Rocket designs and proposals
103 page pdf, advanced propulsion study for the US Air Force made in 2004 prepared by Eric Davis of Warp Drive Metrics. Eric W Davis is an advisor to the Lifeboat Foundation. On pages 48-57, the recent nuclear thermal rocket variants are described. They are estimated to require 5 years of technological development and could have launch costs of $85-150/kg for a single stage to orbit vehicle.
The ETO performance capability of Nuclear DC-X (Paul March, 2001
March, P. (2001), “LANTR VTOL-SSTO Reusable Heavy Cargo Lifter Launch
Vehicle,” Briefing to the Advanced Deep Space Transport Group-Propulsion and Power
Subgroup, and Private Communications, Lockheed-Martin Co., Houston, TX):
* VTOL-SSTO Heavy Cargo Lifter
* Nuclear DC-X propulsion system: 5,000 MWt class LANTR engine
* Utilize Air Force Timber Wind PeBR (see below for discussion of Timber Wind) or Russian Zrhydride heterogeneous reactor design with ternary-carbide fuels, operating at power densities ≈ 20 – 40 MWt/liter with reactor temperature of 3,000 K
* LANTR segmented aerospike exhaust nozzle with variable thrust control on each engine for attitude and flight trajectory control (no gimballing): 5-throttled LANTR engines per vehicle [LANTR is LOX-Augmented Nuclear Thermal Rocket]
* Canard stabilator flight control surfaces
* Landing struts (5) perform multiple functions: provide vehicle support, aerodynamic control, heat rejection, and landing shock absorption
* X-33 type Metallic Reentry Thermal Protection System on the bottom of the vehicle, plus carboncarbon leading edges on all landing struts/stabilators
* The LANTR engines are tilted inboard to place neutron shadow-shield between ground observers and the engines after lift-off – rely on the Conda-effect for flow turning on the aerospike exhaust nozzle
* Neutron shields: graphite-Al walled tanks filled with H2O loaded with 10B
* LANTR engine Oxidizer/Fuel = 4:1
* LANTR engine run time = 200 seconds, total boost time = 500 seconds
* Nuclear DC-X is VTOL from any prepared concrete pad
* 40% GTOW can be carried to LEO (at 400 km altitude and 51° inclination) from a 45° latitude launch
* Dry vehicle mass fraction is 30%, thus giving a payload fraction of 10%, or a payload mass of 10^5 kg to orbit on each flight
* Launch cost estimate: $150/kg of payload, if commercially developed and operated
* Launch cost estimate: $85/kg of payload, if developed and operated by the U.S. government
LANTR : LOX-Augmented Nuclear Thermal Rocket benefits
Summary of LANTR performance improvements over conventional NTR’s (ISNPS, 2003):
LANTR couples a reverse scramjet LOX-afterburner nozzle to a conventional LH2-cooled NTR to
achieve the following benefits:
• LANTR engines are smaller, cheaper NTR’s with “big engine” performance
• Smaller, cheaper facilities for contained ground testing
• Variable thrust and Isp capability from constant power NTR
• Shortened burn times and extended engine life
• Reduced LH2 propellant tank size, mass, and boil-off
• Reduced stage size allowing smaller launch vehicles
• Increased operational range – ability to utilize extraterrestrial sources of O2 and H2
Variants: Pebble Bed, LANTR and Fission Fragment
1. Pellet Bed Reactor (PeBR) NTR (Nuclear Thermal Rocket)
a) Performance in pure NTR mode:
* Isp ≈ 1,000 seconds
* Thrust = 1,112 kN/engine
* Thrust/Weight > 12
* vex (exhaust velocity) = 9.8 km/sec
b) Performance in LANTR mode:
* Isp ≈ 600 seconds
* Thrust = 3,336 kN/engine
* Thrust/Weight > 38
* vex = 5.9 km/sec
2. Thin-Film Fission Fragment Heated NTR
• A high-performance NTR formulated by C. Rubbia
• Reactor core consists of thin-walled porous propellant flow passages coated with a thin layer of Americium-242m
• Propellant is injected radially into the flow passages and heated directly by fission fragments from the Am-242m liner
• This approach allow for much higher bulk temperatures in the propellant than in a
conventional NTR while keeping the propellant in contact with the walls (within the
material temperature limits)
• Theoretical Isp = 2,000 – 4,000 seconds
• Thrust is comparable to a conventional NTR
• Fission Fragment LANTR mode performance is comparable to the PeBR LANTR mode
The PeBR NTR in item 1 above has nuclear fuel that is in the form of a particulate bed (fluidized-bed, dust-bed, or rotating-bed) through which the propellant is pumped (El-Genk et al., 1990; Ludewig, 1990; Horman et al., 1991; ISNPS, 2003). This permits NTR operation at a higher temperature than solid-core NTR’s by reducing the fuel strength requirements. This results in the increased engine performance noted above. The core of the reactor is rotated about its longitudinal axis at approximately 3,000 rpm so that the fuel bed is centrifuged against the inner surface of a cylindrical wall through which H2 gas is injected.
This rotating bed reactor has the advantage that the radioactive particle core can be dumped at the end of an operational cycle and recharged prior to a subsequent burn, thus eliminating the need for decay heat removal, minimizing shielding requirements, and simplifying maintenance and refurbishment operations.
Thin-film fission fragment propulsion involves allowing the energetic fragments produced in the nuclear fission process to directly escape the reactor. Thus, the fission fragments, moving at several percent of the speed of light, can be directly used as the propellant (Chapline, 1988; Wright, 1990; Ronen et al., 2000a, b). However, March (2001) prefers to use Carlo Rubbia’s modification of this concept in which the fragments are used to directly heat a conventional NTR propellant (H2) for propulsion, as described in item 2 above (Rubbia, 1999, 2000; Ronen et al., 2000a, b). In order for the fragments to escape from the nuclear fuel and reactor, a low-mass density critical reactor must be constructed. In order to design such a reactor, highly fissionable nuclear fuels such as Americium (Am) or Curium (Cm) must be used. These fairly rare fuels are produced from reprocessed spent nuclear fuel (via the extraction of Pu-241 and Am-241), which is a very expensive multistep process. However, small amounts of Am- 242m are already available. Ronen et al. (2000a, b) demonstrate that Am-242m can maintain sustained nuclear fission as an extremely thin metallic film, less than 1/1000th of a millimeter thick. Am-242m requires only 1% of the mass of U-235 or Pu-239 to reach its critical state. It should be noted that obtaining fission fragments is not possible with U-235 and Pu-239 nuclear fuels because they both require large fuel rods, which absorb their fission products. The fission fragment propulsion concept is near-term technology, however it requires the development of new technology and technology risk reduction.
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