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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Further analytical study of hybrid rocket combustion

Analytical studies of the transient and steady-state combustion processes in a hybrid rocket system are discussed. The particular system chosen consists of a gaseous oxidizer flowing within a tube of solid fuel, resulting in a heterogeneous combustion. Finite rate chemical kinetics with appropriate reaction mechanisms were incorporated in the model. A temperature dependent Arrhenius type fuel surface regression rate equation was chosen for the current study. The governing mathematical equations employed for the reacting gas phase and for the solid phase are the general, two-dimensional, time-dependent conservation equations in a cylindrical coordinate system. Keeping the simplifying assumptions to a minimum, these basic equations were programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and, the Crank-Nicolson scheme for the solid phase.

Hung, W. S. Y.↗
NASA NTRS1994

Study Of Fuels For Hybrid Rockets

Report describes experimental study of combustion and rates of regression of selected fuels for hybrid rocket engines. Part of continuing effort to develop fuels with greater rates of regression and lesser dependence on shapes of fuel grains and to maximize potential specific impulse at low cost.

Strand, Leon D.↗
NASA NTRS1996

CFD Analysis of the 24-inch JIRAD Hybrid Rocket Motor

A series of multispecies, multiphase computational fluid dynamics (CFD) analyses of the 24-inch diameter joint government industry industrial research and development (JIRAD) hybrid rocket motor is described. The 24-inch JIRAD hybrid motor operates by injection of liquid oxygen (LOX) into a vaporization plenum chamber upstream of ports in the hydroxyl-terminated polybutadiene (HTPB) solid fuel. The injector spray pattern had a strong influence on combustion stability of the JIRAD motor so a CFD study was initiated to define the injector end flow field under different oxidizer spray patterns and operating conditions. By using CFD to gain a clear picture of the flow field and temperature distribution within the JIRAD motor, it is hoped that the fundamental mechanisms of hybrid combustion instability may be identified and then suppressed by simple alterations to the oxidizer injection parameters such as injection angle and velocity. The simulations in this study were carried out using the General Algorithm for Analysis of Combustion SYstems (GALACSY) multiphase combustion codes. GALACSY consists of a comprehensive set of droplet dynamic submodels (atomization, evaporation, etc.) and a computationally efficient hydrocarbon chemistry package built around a robust Navier-Stokes solver optimized for low Mach number flows. Lagrangian tracking of dispersed particles describes a closely coupled spray phase. The CFD cases described in this paper represent various levels of simplification of the problem. They include: (A) gaseous oxygen with combusting fuel vapor blowing off the walls at various oxidizer injection angles and velocities, (B) gaseous oxygen with combusting fuel vapor blowing off the walls, and (C) liquid oxygen with combusting fuel vapor blowing off the walls. The study used an axisymmetric model and the results indicate that the injector design significantly effects the flow field in the injector end of the motor. Markedly different recirculation patterns are observed in the vaporization chamber as the oxygen velocity and/or spray pattern is varied. The ability of these recirculation patterns to stabilize the diffusion flame above the surface of the solid fuel gives a plausible explanation for the experimentally determined combustion stability characteristics of the JIRAD motor, and suggests how combustion stability can be assured by modifications to the injector design.

Liang, Pak-Yan↗
NASA NTRSDate not supplied

Aluminum borohydride ammines as fuels for hybrid rocket motors: theoretical performance and particle synthesis

The aluminum borohydride ammines, Al(BH4)3·XNH3, are a family of complexes between aluminum borohydride and ammonia for which up to six moles of ammonia can be complexed per mole of aluminum borohydride. Two of these complexes, the diammine (X = 2) and hexammine (X = 6) are of particular interest for use in hybrid rocket motors as particle additives and stand alone fuels. This study compares these compounds against other metal hydrides for their potential to increase the regression rate of solid fuel grains. The theoretical performance of the diammine as a stand alone fuel paired with non-toxic oxidizers is also investigated. The low temperature and good performance of these combinations make them of interest for low cost orbital maneuvering thrusters. Finally, results on the synthesis of Al(BH4)3·6NH3 and its thermal characterization are presented.

hybrid rockets↗
NASA NTRSDate not supplied

Aluminum Borohydride Ammines as Fuels for Hybrid Rocket Motors: Theoretical Performance and Particle Synthesis

The aluminum borohydride ammines, Al(BH4)3·XNH3, are a family of complexes between aluminum borohydride and ammonia for which up to six moles of ammonia can be complexed per mole of aluminum borohydride. Two of these complexes, the diammine (X = 2) and hexammine (X = 6) are of particular interest for use in hybrid rocket motors as additives and stand alone fuels. This study compares these compounds against other metal hydride for their potential as additives to increase the regression rate of solid fuel grains. The theoretical performance of the diammine as a stand alone fuel paired with non-toxic oxidizers is also investigated. The low temperature and good performance of these combinations make them of interest for low cost orbital maneuvering thrusters. Finally, the synthesis process for Al(BH4)3·6NH3 is discussed and TGA/DSC results are presented.

hybrid rockets↗
NASA NTRS1972

Numerical solutions of reactive fluid flows during postignition transients in hybrid rocket systems.

A computational method has been developed for the study of the post-ignition transients in hybrid rocket systems. The particular system chosen consisted of a gaseous oxidizer flowing within a tube of solid fuel, resulting in heterogeneous combustion. With the appropriate assumptions, two-dimensional, time-dependent conservation equations were derived for the reacting gas phase, and for the solid phase, in a cylindrical coordinate system. These were then programmed for numerical computation, using two implicit finite-difference schemes, the Lax-Wendroff scheme for the gas phase, and the Crank-Nicolson scheme for the solid phase. Appropriate initial and boundary conditions were represented, including heat and mass conservation at the interface between gas and solid. Initially, no attempt was made to relate the recession rate at the surface to the surface temperature, or to include heat transfer by radiation. A simple case was selected for preliminary calculations, with aluminum and oxygen as fuel and oxidizer, and aluminum oxide as the product.

Hung, W. S. Y.↗
NASA NTRS1995

Computational fluid dynamic analysis of hybrid rocket combustor flowfields

Computational fluid dynamic analyses of the Navier-Stokes equations coupled with solid-phase pyrolysis, gas-phase combustion, turbulence and radiation are performed to study hybrid rocket combustor flowfields. The computational study is closely co-ordinated with a companion experimental program using a planar slab burner configuration with HTPB as fuel and gaseous oxygen. Computational predictions agree reasonably well with measurement data of fuel regression rates and surface temperatures. Additionally, most of the parametric trends predicted by the model are in general agreement with experimental trends. The computational model is applied to extend the results from the lab-scale to a full-scale axisymmetric configuration. The numerical predictions indicate that the full-scale configuration burns at a slower rate than the lab-scale combustor under identical specific flow rate conditions. The results demonstrate that detailed CFD analyses can play a useful role in the design of hybrid combustors.

Venkateswaran, S.↗
NASA NTRS1996

Fundamental Phenomena on Fuel Decomposition and Boundary-Layer Combustion Precesses with Applications to Hybrid Rocket Motors: Experimental Investigation - Part 1

This final report summarizes the major findings on the subject of 'Fundamental Phenomena on Fuel Decomposition and Boundary-Layer Combustion Processes with Applications to Hybrid Rocket Motors', performed from 1 April 1994 to 30 June 1996. Both experimental results from Task 1 and theoretical/numerical results from Task 2 are reported here in two parts. Part 1 covers the experimental work performed and describes the test facility setup, data reduction techniques employed, and results of the test firings, including effects of operating conditions and fuel additives on solid fuel regression rate and thermal profiles of the condensed phase. Part 2 concerns the theoretical/numerical work. It covers physical modeling of the combustion processes including gas/surface coupling, and radiation effect on regression rate. The numerical solution of the flowfield structure and condensed phase regression behavior are presented. Experimental data from the test firings were used for numerical model validation.

Kuo, Kenneth K.↗
NASA NTRS1994

Characterization of Hybrid Rocket Internal Heat Flux and HTPB Fuel Pyrolysis

None given. From statement of problem:The broad nature of this work has been to extend our understanding of the nature of the conventional or classical hybrid rocket combustion process (fuel decomposition, heat and mass transfer, and chemical reaction), and the critical engineering parameters that define this process.

rocket↗
NASA NTRS1999

Initial Test Firing Results for Solid CO/GOX Cryogenic Hybrid Rocket Engine for Mars ISRU Propulsion Applications

ORBITEC is developing methods for producing, testing, and utilizing Mars-based ISRU fuel/oxidizer combinations to support low cost, planetary surface and flight propulsion and power systems. When humans explore Mars we will need to use in situ resources that are available, such as: energy (solar); gases or liquids for life support, ground transportation, and flight to and from other surface locations and Earth; and materials for shielding and building habitats and infrastructure. Probably the easiest use of Martian resources to reduce the cost of human exploration activities is the use of the carbon and oxygen readily available from the CO2 in the Mars atmosphere. ORBITEC has conducted preliminary R&D that will eventually allow us to reliably use these resources. ORBITEC is focusing on the innovative use of solid CO as a fuel. A new advanced cryogenic hybrid rocket propulsion system is suggested that will offer advantages over LCO/LOX propulsion, making it the best option for a Mars sample return vehicle and other flight vehicles. This technology could also greatly support logistics and base operations by providing a reliable and simple way to store solar or nuclear generated energy in the form of chemical energy that can be used for ground transportation (rovers/land vehicles) and planetary surface power generators. This paper describes the overall concept and the test results of the first ever solid carbon monoxide/oxygen rocket engine firing.

Rice, Eric E.↗
NASA NTRS1999

Simulation of Non-Acoustic Combustion Instability in a Hybrid Rocket Motor

A transient model of a hybrid motor was formulated to study the cause and elimination of non-acoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne and Martin Marietta at NASA/Marshall Space Flight Center (NASA/MSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5 Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5 Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in the elimination of combustion instability with the installation of an orifice immediately upstream of the injector. The formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom, and Claflin. The previous model simulated an unstable IR&D hybrid motor test performed by Thiokol. There was very good agreement between the model and the test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, the tests performed at NASA/MSFC under the HPTLVB program were actually simulated. In the current model, the hybrid motor consisting of the liquid oxygen (LOX) injector, the multi-port solid fuel grain and the nozzle was simulated. Also, simulated in the model was the LOX feed system consisting of the tank, venturi, valve and feed lines. All components of the hybrid motor and LOX feed system are treated by a lumped-parameter approach. Agreement between the results of the transient model and the actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes. The first cause was a LOX feed system of insufficient stiffness. The second cause was a LOX injector with an impedance or pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, Marvin↗
NASA NTRS1999

Simulation of Non-Acoustic Combustion Instability in a Hybrid Rocket Motor

A transient model of a hybrid motor was formulated to study the cause and elimination of non-acoustic combustion instability. The transient model was used to simulate four key tests out of a series of seventeen hybrid motor tests conducted by Thiokol, Rocketdyne and Martin Marietta at NASA/Marshall Space Flight Center (NASAIMSFC). These tests were performed under the Hybrid Propulsion Technology for Launch Vehicle Boosters (HPTLVB) program. The first test resulted in stable combustion. The second test resulted in large-amplitude, 6.5 Hz chamber pressure oscillations that gradually damped away by the end of the test. The third test resulted in large-amplitude, 7.5 Hz chamber pressure oscillations that were sustained throughout the test. The seventh test resulted in the elimination of combustion instability with the installation of an orifice immediately upstream of the injector. The formulation and implementation of the model are the scope of this presentation. The current model is an independent continuation of modeling presented previously by joint Thiokol-Rocketdyne collaborators Boardman, Hawkins, Wassom, and Claflin. The previous model simulated an unstable IR&D hybrid motor test performed by Thiokol. There was very good agreement between the model and the test data. Like the previous model, the current model was developed using Matrix-x simulation software. However, the tests performed at NASA/MSFC under the HPTLVB program were actually simulated. In the current model, the hybrid motor consisting of the liquid oxygen (LOX) injector, the multi-port solid fuel grain and the nozzle was simulated. Also, simulated in the model was the LOX feed system consisting of the tank, venturi, valve and feed lines. All components of the hybrid motor and LOX feed system are treated by a lumped-parameter approach. Agreement between the results of the transient model and the actual test data was very good. This agreement between simulated and actual test data indicated that the combustion instability in the hybrid motor was due to two causes. The first cause was a LOX feed system of insufficient stiffness. The second cause was a LOX injector with an impedance or pressure drop that was too low to provide damping against the feed system oscillations. Also, it was discovered that testing with a new grain of solid fuel sustained the combustion instability. However, testing with a used grain of solid fuel caused the combustion instability to gradually decay.

Rocker, Marvin↗
DOE OSTI2025

Direct numerical simulations for hybrid rocket boundary layers: Performance modeling and scaling

This paper presents a comprehensive performance and scaling analysis of direct numerical simulations for reacting boundary layers, focusing on slab burner configurations. Using a PETSc-based finite volume CFD framework, the study evaluates the scalability and computational cost of flow, chemistry, and radiation evaluations across 2D and 3D simulations. Polymethyl methacrylate (PMMA) is the fuel with pure O 2 as the oxidizer, modeled using a detailed chemical kinetics mechanism with 113 species and 660 reactions. A ray-tracing-based radiation solver, designed for distributed memory applications, is implemented to model radiation heat transfer. Parallel scalability is analyzed for the coupled flow, chemistry, and radiation heat transfer processes. Weak and strong scaling studies are conducted on up to 15,000 computational ranks, revealing robust performance when flow cells exceed 200 per rank. Chemistry evaluations dominate the computational cost in large 3D simulations, accounting for approximately 40% of the total runtime, while flow processes contribute around 35%, and radiation solver contributions remain below 10% due to reduced evaluation frequencies. GPU accelerated chemistry evaluation, implemented with Zero-RK, demonstrates significant promise, achieving up to a 4x speedup for workloads exceeding 30,000 cells per GPU. However, diminishing returns are observed for smaller workloads due to CPU-GPU communication overhead. This study identifies key challenges, including memory bottlenecks and the effects of domain partitioning on flow scalability, while highlighting the potential of GPU-accelerated chemistry to reduce computational costs. In conclusion, these findings provide realizable run configurations for 2D, 3D, and GPU-accelerated cases, offering insights for optimizing reactive flow solvers.

CFD Scalability↗
NASA NTRS1993

Hybrid rocket combustion study

The objectives of this study of 'pure' or 'classic' hybrids are to (1) extend our understanding of the boundary layer combustion process and the critical engineering parameters that define this process, (2) develop an up-to-date hybrid fuel combustion model, and (3) apply the model to correlate the regression rate and scaling properties of potential fuel candidates. Tests were carried out with a hybrid slab window motor, using several diagnostic techniques, over a range of motor pressure and oxidizer mass flux conditions. The results basically confirmed turbulent boundary layer heat and mass transfer as the rate limiting process for hybrid fuel decomposition and combustion. The measured fuel regression rates showed good agreement with the analytical model predictions. The results of model scaling calculations to Shuttle SRM size conditions are presented.

Strand, L. D.↗
NASA NTRS1995

Improved hybrid rocket fuel

McDonnell Douglas Aerospace, as part of its Independent R&D, has initiated development of a clean burning, high performance hybrid fuel for consideration as an alternative to the solid rocket thrust augmentation currently utilized by American space launch systems including Atlas, Delta, Pegasus, Space Shuttle, and Titan. It could also be used in single stage to orbit or as the only propulsion system in a new launch vehicle. Compared to solid propellants based on aluminum and ammonium perchlorate, this fuel is more environmentally benign in that it totally eliminates hydrogen chloride and aluminum oxide by products, producing only water, hydrogen, nitrogen, carbon oxides, and trace amounts of nitrogen oxides. Compared to other hybrid fuel formulations under development, this fuel is cheaper, denser, and faster burning. The specific impulse of this fuel is comparable to other hybrid fuels and is between that of solids and liquids. The fuel also requires less oxygen than similar hybrid fuels to produce maximum specific impulse, thus reducing oxygen delivery system requirements.

Dean, David L.↗
NASA NTRS1992

Hybrid rocket fuel combustion and regression rate study

The objectives of this study are to develop hybrid fuels (1) with higher regression rates and reduced dependence on fuel grain geometry and (2) that maximize potential specific impulse using low-cost materials. A hybrid slab window motor system was developed to screen candidate fuels - their combustion behavior and regression rate. Combustion behavior diagnostics consisted of video and high speed motion pictures coverage. The mean fuel regression rates were determined by before and after measurements of the fuel slabs. The fuel for this initial investigation consisted of hydroxyl-terminated polybutadiene binder with coal and aluminum fillers. At low oxidizer flux levels (and corresponding fuel regression rates) the filled-binder fuels burn in a layered fashion, forming an aluminum containing binder/coal surface melt that, in turn, forms into filigrees or flakes that are stripped off by the crossflow. This melt process appears to diminish with increasing oxidizer flux level. Heat transfer by radiation is a significant contributor, producing the desired increase in magnitude and reduction in flow dependency (power law exponent) of the fuel regression rate.

Strand, L. D.↗
NASA NTRS2017

Peregrine Hybrid Rocket Ground Test Video

The attached edited and compressed video shows the final ground test of the Peregrine hybrid rocketflight-weight motor that took place at NASA Ames on March 15, 2017.

Zilliac, Gregory↗