Simulated LOX-augmented nuclear thermal rocket (LANTR) testing

Bulman, Melvin; Neill, Todd

doi:10.2514/6.2000-3897

Cited by 18 publications

(6 citation statements)

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“…These experiments found a 40% increase in thrust due to the oxygen afterburning, with significant combustion occurring beyond the nozzle exit [14]. Follow-on experiments found similar results for larger nozzles with higher expansion ratios [15,16] …”

Section: Oxygen Afterburningsupporting

confidence: 63%

“…Another novel concept of recent interest uses the nuclear reactor to produce thrust and electric power for the spacecraft. Both concepts have been studied extensively; the former has been developed extensively by the Aerojet Corporation [14]- [18], but due the weight penalty and combustion losses incurred in such a system, it is not examined here. The latter, known as "bimodal" operation, is currently of great interest to NASA and is the focus of the remainder of this dissertation.…”

Section: Operationmentioning

confidence: 99%

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Integrated Propulsion and Power Modeling for Bimodal Nuclear Thermal Rockets

Clough¹,

Starkey²,

Lewis³

et al. 2007

43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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Design point results show that a 316 MWt reactor produces a thrust and specific impulse of 66.6 kN and 917 s, respectively. The same reactor can be run at 73.8 kWt to produce the necessary 16.7 kW electric power with a Brayton cycle generator. This demonstrates the feasibility of BNTR operation with a NERVAderived reactor but also indicates that the reactor control system must be able to operate with precision across a wide power range, and that the transient analysis of reactor decay heat merits future investigation.Results also identify a significant reactor pressure-drop limitation during propulsion and power-generation operation that is caused by poor tie tube thermal conductivity. This leads to the conclusion that, while BNTR operation is possible with a NERVA-derived reactor, doing so requires careful consideration of the Brayton cycle design point and fuel element survivability.

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Section: Oxygen Afterburningsupporting

confidence: 63%

Section: Operationmentioning

confidence: 99%

Integrated Propulsion and Power Modeling for Bimodal Nuclear Thermal Rockets

Clough¹,

Starkey²,

Lewis³

et al. 2007

43rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference &Amp;amp; Exhibit

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“…The LOx provides an augmentation factor of up to 4:1 (3:1 in this example, from Ref. [18][19][20] and a mission average Isp of 821 seconds, but reactor and structural mass lower payload fractions to negative margin with the selected 120 klb GLOW and take-off T/W of 1.3. Scale favors rocketry, so a positive margin solution may exist but this study was restricted to a known reactor architecture.…”

Section: A Earth To Orbitmentioning

confidence: 99%

The Nuclear Thermal Turbo Rocket - A Conceptual High-Performance Earth-to-Orbit Propulsion System

Bucknell¹

2015

51st AIAA/SAE/ASEE Joint Propulsion Conference

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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 ambientsourced 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. Nomenclature m = mass flow η c = compressor isentropic efficiency c p = constant pressure specific heat γ = ratio of specific heats T = absolute temperature P = absolute pressure M = Mach number MW th = Megawatt, thermal M f = Final Mass PR = Pressure Ratio (Outlet/Inlet) ΔV = change in velocity Φ = equivalence ratio k = reactivity coefficient Ψ = absolute temperature ratio β = oblique shock angle  = flow deflection angle

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“…Here oxygen alone is injected in the diverging part of the nozzle and thrust is augmented due to supersonic combustion with nuclear preheated hydrogen. Later Bulman et al (2000) conducted experiments with the injection of gaseous oxygen in the down-stream of the throat of a NTR nozzle to study the thrust enhancement. Numerical simulations also have been performed for a hot fire test of the concept and successfully predicted the performance.…”

Section: Review Of Literaturementioning

confidence: 99%

Numerical Studies on Thrust Augmentation in High Area Ratio Rocket Nozzles by Secondary Injection

Shyji¹,

Deepu

Kumar³

et al. 2017

JAFM

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Single stage to orbit propulsion devices are being developed as part of low cost access to space endeavors. Sea level operation of high area ratio rocket nozzle used in rocket engines leads to an overexpanded flow condition resulting in high side loads. Secondary injection of propellants in high area ratio nozzle is an attractive option to overcome the inefficiency of operation of such nozzles in sea level conditions in addition to the augmentation of thrust. A numerical study on thrust augmentation in high area ratio nozzle by secondary injection of propellants is presented here. The turbulent compressible reacting flow in rocket nozzle with auxiliary injection is simulated using conservation equations for chemical species based on finite rate chemistry model and compressible Navier-Stokes equations with AUSM+-up upwind scheme based unstuctured finite volume solver. An optimized eight step, six species reduced H2-O2finite chemistry reaction model is used to model the supersonic combustion. The indigenously developed solver has an efficient rescaling algorithm to alleviate the effect of stiffness in conventional explicit algorithm for simultaneous solution of reacting flow. The code is validated using the wall pressure and hydrogen concentration values reported for the similar high area ratio rocket nozzle. Accurate prediction of nozzle performance is possible with present turbulent reacting flow simulation as it take care of all losses in nozzle flow. Extensive computations have been performed for the performance estimation of high area ratio rocket nozzle for various prospective auxiliary injection options.

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Simulated LOX-augmented nuclear thermal rocket (LANTR) testing

Cited by 18 publications

References 0 publications

Integrated Propulsion and Power Modeling for Bimodal Nuclear Thermal Rockets

Integrated Propulsion and Power Modeling for Bimodal Nuclear Thermal Rockets

The Nuclear Thermal Turbo Rocket - A Conceptual High-Performance Earth-to-Orbit Propulsion System

Numerical Studies on Thrust Augmentation in High Area Ratio Rocket Nozzles by Secondary Injection

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