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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 397 records · Page 22

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NASA NTRS1988

The ram accelerator - A chemically driven mass launcher

The ram accelerator, a chemically propelled mass driver, is presented as a viable new approach for directly launching acceleration-insensitive payloads into low earth orbit. The propulsion principle is similar to that of a conventional air-breathing ramjet. The cargo vehicle resembles the center-body of a ramjet and travels through a tube filled with a pre-mixed fuel and oxidizer mixture. The launch tube acts as the outer cowling of the ramjet and the combustion process travels with the vehicle. Two drive modes of the ram accelerator propulsion system are described, which when used in sequence are capable of accelerating the vehicle to as high as 10 km/sec. The requirements are examined for placing a 2000 kg vehicle into a 500 km orbit with a minimum of on-board rocket propellant for circularization maneuvers. It is shown that aerodynamic heating during atmospheric transit results in very little ablation of the nose. An indirect orbital insertion scenario is selected, utilizing a three step maneuver consisting of two burns and aerobraking. An on-board propulsion system using storable liquid propellants is chosen in order to minimize propellant mass requirements, and the use of a parking orbit below the desired final orbit is suggested as a means to increase the flexibility of the mass launch concept. A vehicle design using composite materials is proposed that will best meet the structural requirements, and a preliminary launch tube design is presented.

Kaloupis, P.↗

NASA NTRS1989

Evolution Of Cenospheres

Microscopy reveals changes in structure during combustion. Experiments performed to trace evolution of structures of droplets of burning fuel oil. Many droplets burn incompletely, forming cenospheres. Knowledge of structure at various stages of combustion process contributes to efforts to design equipment to burn cenospheres.

Kwack, Eug Y.↗

NASA NTRS1988

Modeling of hydrogen-air diffusion flame

The present research objective is to determine the effects of contaminants on extinction limits of simple, well defined, counterflow Hydrogen 2-air diffusion flames, with combustion at 1 atmosphere. Results of extinction studies and other flame characterizations, with appropriate mechanistic modeling (presently underway), will be used to rationalize the observed effects of contamination over a reasonably wide range of diffusion flame conditions. The knowledge gained should help efforts to anticipate the effects of contaminants on combustion processes in Hydrogen 2-fueled scramjets.

Isaac, Kakkattukuzhy↗

NASA NTRS1989

The solid surface combustion space shuttle experiment hardware description and ground-based test results

The Lewis Research Center is developing a series of microgravity combustion experiments for the Space Shuttle. The Solid Surface Combustion Experiment (SSCE) is the first to be completed. SSCE will study flame spreading over thermally thin fuels (ashless filter paper) under microgravity conditions. The flight hardware consists of a combustion chamber containing the sample and a computer which takes the data and controls the experiment. Experimental data will include gas-phase and solid-phase temperature measurements and motion pictures of the combustion process. Flame spread rates will be determined from the motion pictures.

Vento, D. M.↗

NASA NTRS1989

A numerical study of chemically reacting flow in nozzles

The space station uses small rocket motors, called thrusters, for orientation control. Because of the lack of viable design tools for small rockets, the initial thruster design was basically a very small version of a large rocket motor. Thrust measurements of the initial design were lower than predicted. To improve predictions it was decided to develop a verison of the RPLUS2D reacting flow code for thruster calculations. RPLUS2D employs an implicit finite volume, lower-upper symmetric successive overrelaxation (LU-SSOR) scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a coupled and very efficient manner. The combustion processes are modeled by a 9-species, 18 step finite-rate chemistry model, and the turbulence is simulated by a Baldwin-Lomax algebraic model. The code is extended to handle multiple subsonic inlet conditions where the total mass flow is governed by conditions calculated at the thruster-throat. Results are shown for a thruster design where the overall mixture ratio is hydrogen rich. A calculation of a large area ratio divergent nozzle is also presented.

Vanoverbeke, Thomas J.↗

NASA NTRS1989

The solid surface combustion Space Shuttle experiment hardware description and ground-based test results

Vento, D. M.↗

NASA NTRS1989

A numerical study of chemically reacting flow in nozzles

The Space Station uses small rocket motors, called thrusters, for orientation control. Because of the lack of viable design tools for small rockets, the initial thruster design was basically a very small version of a large rocket motor. Thrust measurements of the initial design were lower than predicted. To improve predictions it was decided to develop a version of the RPLUS2D reacting flow code for thruster calculations. RPLUS2D employs an implicit finite volume, lower-upper symmetric successive overrelaxation (LU-SSOR) scheme for solving the complete two-dimensional Navier-Stokes equations and species transport equations in a coupled and very efficient manner. The combustion processes are modeled by a 9-species, 18 step finite-rate chemistry model, and the turbulence is simulated by a Baldwin-Lomax algebraic model. The code is extended to handle multiple subsonic inlet conditions where the total mass flow is governed by conditions calculated at the thruster-throat. Results are shown for a thruster design where the overall mixture ratio is hydrogen rich. A calculation of a large area ratio divergent nozzle is also presented.

Vanoverbeke, Thomas J.↗

NASA NTRS1989

Conceptual design for the space station Freedom modular combustion facility

A definition study and conceptual design for a combustion science facility that will be located in the Space Station Freedom's baseline U.S. Laboratory module is being performed. This modular, user-friendly facility, called the Modular Combustion Facility, will be available for use by industry, academic, and government research communities in the mid-1990's. The Facility will support research experiments dealing with the study of combustion and its byproducts. Because of the lack of gravity-induced convection, research into the mechanisms of combustion in the absence of gravity will help to provide a better understanding of the fundamentals of the combustion process. The background, current status, and future activities of the effort are covered.

Source record↗

NASA NTRS1990

Computational experience with a three-dimensional rotary engine combustion model

A new computer code was developed to analyze the chemically reactive flow and spray combustion processes occurring inside a stratified-charge rotary engine. Mathematical and numerical details of the new code were recently described by the present authors. The results are presented of limited, initial computational trials as a first step in a long-term assessment/validation process. The engine configuration studied was chosen to approximate existing rotary engine flow visualization and hot firing test rigs. Typical results include: (1) pressure and temperature histories, (2) torque generated by the nonuniform pressure distribution within the chamber, (3) energy release rates, and (4) various flow-related phenomena. These are discussed and compared with other predictions reported in the literature. The adequacy or need for improvement in the spray/combustion models and the need for incorporating an appropriate turbulence model are also discussed.

Raju, M. S.↗

NASA NTRS1990

A computational investigation of flow losses in a supersonic combustor

This computational investigation provides a detailed comparative study of both non-reacting and reacting supersonic combustor flow-fields and attendant flow losses. Three different injection configurations in the same combustor are examined; a swept-sided ramp with base hydrogen injection, a straight ramp with base hydrogen injection and a thirty-degree downstream directed wall jet. Detailed comparisons are made with available reacting experimental data for the swept and the unswept ramps. A seven reaction, seven species reaction model as well as a global (one reaction) model are used for all three reacting cases. Relative performance (as measured by thrust) of all cases are described. Details of the flow, mixing, and combustion processes are discussed.

Riggins, David W.↗

NASA NTRS1990

Calculations of gaseous H2/O2 thruster

Calculations were made for the gaseous H2/O2 thruster for the Space Station by using the RPLUS code which employs an implicit finite volume, LUSSOR scheme to solve the Navier-Stokes equations and the species equations. The combustion processes of hydrogen and oxygen are modeled by a 9 species and 18 step reaction mechanism and the turbulence is simulated by the Baldwin-Lomax turbulence model for the thruster wall boundary layer and the modified Prandtl's mixing length model for the reacting shear layer. Results are presented for different mixture ratios and fuel film cooling percents and compared with the experimental data. The calculated performance predictions for the thruster agree well with the experimental data and the results demonstrate that the RPLUS code can be used for design and analysis of thrusters and rockets.

Kim, S. C.↗

NASA NTRS1991

Computer code for single-point thermodynamic analysis of hydrogen/oxygen expander-cycle rocket engines

This analysis and this computer code apply to full, split, and dual expander cycles. Heat regeneration from the turbine exhaust to the pump exhaust is allowed. The combustion process is modeled as one of chemical equilibrium in an infinite-area or a finite-area combustor. Gas composition in the nozzle may be either equilibrium or frozen during expansion. This report, which serves as a users guide for the computer code, describes the system, the analysis methodology, and the program input and output. Sample calculations are included to show effects of key variables such as nozzle area ratio and oxidizer-to-fuel mass ratio.

Glassman, Arthur J.↗

NASA NTRS1990

A three-dimensional approach for analysis of sidewall injector mixing and combustion

A three-dimensional (3-D) analytical approach to calculate the mixing and combustion efficiency along engine combustors that use sidewall fuel injectors is described. The analysis consists of three parts: (1) an empirical correlation for fuel penetration, (2) application of integral conservation laws following jet turning and pressure equilibration, and (3) an analytical solution for the simplified three-dimensional partial differential equations to describe the downstream mixing and combustion process. Numerical results are compared to some empirical and experimental work, and favorable agreements are shown. The importance of including anisotropic turbulence and three-dimensional effects are illustrated by comparison with data and other less comprehensive analyses. The analytical solution is suitable for rapid estimates of engine performance and parametric studies of combustor designs containing transverse or angled injector ports arranged in rows and/or columns.

Quan, Victor↗

NASA NTRS1991

CFD analysis of a hydrogen fueled ramjet engine at Mach 3.44

The full Navier-Stokes solver, RPLUS, has been used to compute the flow field of a ramjet engine in the low supersonic operating regime. Computations have been made for the case where hydrogen is transversely injected and ignited using nine species and eighteen reactions to model the combustion process. The flow fields for two methods of hydrogen injection have been computed and compared to the available experimental data. The purpose of these computations was to assess the mixing effectiveness of the injected hydrogen with the primary supersonic stream. Results indicate significant ignition delay resulting from the poor mixing dynamics in the combustor section.

Duncan, Beverly↗

NASA NTRS1991

Numerical study of high-area-ratio H2/O2 rocket nozzles

Calculations were made for high-area-ratio H2/O2 rocket nozzles by using the RPLUS code which employs an implicit finite volume, lower-upper symmetric successive over-relaxation scheme to solve the Navier-Stokes equations and the species equations in a coupled manner. The combustion processes of hydrogen and oxygen are modeled by an 8-species and 18-step reaction mechanism, and the turbulence is simulated by the Baldwin-Lomax turbulence model with a pressure gradient correction on the van Driest's damping constant in the inner-layer model. The calculated results for two different rocket nozzles are presented as contours, profiles, and performance values. The predicted specific impulse and thrust agree well with the experimental data and those from the industry standard code. It is demonstrated that the RPLUS code can be used for analysis of rocket nozzles.

Kim, S. C.↗

NASA NTRS1991

Application of a Reynolds Stress turbulence model to a supersonic hydrogen-air diffusion flame

A second-order differential Reynolds Stress turbulence model has been applied to the Favre-averaged Navier-Stokes equations for the study of supersonic flows undergoing hydrogen-air chemical reactions. An assumed Beta Probability Density Function is applied to account for the chemical source terms in the conservation equations. An algebraic Reynolds Flux model is used for the fluctuating density-velocity as well as the species mass fraction-velocity correlations. The variances of temperature and species fluctuations are also modelled using an algebraic flux technique. A seven-species, seven-reaction finite rate chemistry mechanism is used to simulate the combustion processes. The resulting formulation is validated by comparison with experimental data on reacting supersonic axisymmetric jets. Results obtained for specific conditions indicate that the effect of chemical reaction on the turbulence is significant.

Chandrasekhar, R.↗

NASA NTRS1991

Navier-Stokes simulation of the supersonic combustion flowfield in a ram accelerator

A computational study of the ram accelerator, a ramjet-in-tube device for accelerating projectiles to ultrahigh velocities, is presented. The analysis is performed using a fully implicit TVD scheme that efficiently solves the Reynolds-averaged Navier-Stokes equations and the species continuity equations associated with a finite rate combustion model. The present results indicate that viscous effects are of primary importance in all the cases studied, shock-induced combustion always started in the boundary layer. The effects of Mach number, mixture composition, pressure and tubulence are investigated for various configurations. Two types of combustion processes, one stable and the other unstable, were observed depending on the inflow conditions. The possibility of stabilizing the detonation wave by means of a backward facing step is also investigated. Two numerical techniques were tested: vector extrapolation, to accelerate convergence, and a diagonal formulation that eliminates the expense of inverting large block matrices which arise in chemically reacting flows.

Yungster, Shaye↗

NASA NTRS1991

Experiments on hypersonic ramjet propulsion cycles using a ram accelerator

Work on hypersonic propulsion research using a ram accelerator is presented. Several different ram accelerator propulsive cycles have been experimentally demonstrated over the Mach number range of 3 to 8.5. The subsonic, thermally choked combustion mode has accelerated projectiles to near the Chapman-Jouguet (C-J) detonation velocity within many different propellant mixtures. In the transdetonative velocity regime (85 to 115 percent of C-J speed), projectiles have established a propulsive cycle which allows them to transition smoothly from subdetonative to superdetonative velocities. Luminosity data indicate that the combustion process moves forward onto the projectile body as it approaches the C-J speed. In the superdetonative velocity range, the projectiles accelerate while always traveling faster than the C-J velocity. Ram accelerator projectiles operating continuously through these velocity regimes generate distinctive hypersonic phenomena which can be studied very effectively in the laboratory. These results would be very useful for validating sophisticated CFD computer codes and in collecting engineering data for potential airbreathing hypersonic propulsive systems.

Chew, G.↗