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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 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Laboratory test methods for combustion stability properties of solid propellants

An overview is presented of experimental methods for determining the combustion-stability properties of solid propellants. The methods are generally based on either the temporal response to an initial disturbance or on external methods for generating the required oscillations. The size distribution of condensed-phase combustion products are characterized by means of the experimental approaches. The 'T-burner' approach is shown to assist in the derivation of pressure-coupled driving contributions and particle damping in solid-propellant rocket motors. Other techniques examined include the rotating-valve apparatus, the impedance tube, the modulated throat-acoustic damping burner, and the magnetic flowmeter. The paper shows that experimental methods do not exist for measuring the interactions between acoustic velocity oscillations and burning propellant.

Strand, L. D.↗
NASA NTRS1987

Development of a generic combustion stability code for liquid propellant rocket engines

The mathematical framework for a combustion stability analysis code is outlined. The goal for the code is to be general enough in problem treatment so that its validity and accuracy extend over a wide range of problem applications and that it lends the convenience for any future model improvement if necessary. An approach for modeling the combustion dynamics is devised to meet both requirements. An open-loop numerical procedure is also developed to mechanistically model various combustion processes for determining the stability parameters.

Fang, J. J.↗
NASA NTRS1998

JANNAF 35th Combustion Subcommittee Meeting

Volume 1, the first of two volumes is a compilation of 63 unclassified/unlimited distribution technical papers presented at the 35th meeting of the Joint Army-Navy-NASA-Air Force (JANNAF) Combustion Subcommittee (CS) held jointly with the 17th Propulsion Systems Hazards Subcommittee (PSHS) and Airbreathing Propulsion Subcommittee (APS). The meeting was held on 7-11 December 1998 at Raytheon Systems Company and the Marriott Hotel, Tucson, AZ. Topics covered include solid gun propellant processing, ignition and combustion, charge concepts, barrel erosion and flash, gun interior ballistics, kinetics and molecular modeling, ETC gun modeling, simulation and diagnostics, and liquid gun propellant combustion; solid rocket motor propellant combustion, combustion instability fundamentals, motor instability, and measurement techniques; and liquid and hybrid rocket combustion.

Fry, Ronald S.↗
NASA NTRS1991

Ignition and combustion characteristics of metallized propellants

Over the past six months, experimental investigations were continued and theoretical work on the secondary atomization process was begun. Final shakedown of the sizing/velocity measuring system was completed and the aluminum combustion detection system was modified and tested. Atomizer operation was improved to allow steady state operation over long periods of time for several slurries. To validate the theoretical modeling, work involving carbon slurry atomization and combustion was begun and qualitative observations were made. Simultaneous measurements of aluminum slurry droplet size distributions and detection of burning aluminum particles were performed at several axial locations above the burner. The principle theoretical effort was the application of a rigid shell formation model to aluminum slurries and an investigation of the effects of various parameters on the shell formation process. This shell formation model was extended to include the process leading up to droplet disruption, and previously developed analytical models were applied to yield theoretical aluminum agglomerate ignition and combustion times. The several theoretical times were compared with the experimental results.

Mueller, D. C.↗
NASA NTRS1992

Ignition and combustion characteristics of metallized propellants

During this reporting period, theoretical work on the secondary atomization process was continued and the experimental apparatus was improved. A one-dimensional model of a rocket combustor, incorporating multiple droplet size classes, slurry combustion, secondary atomization, radiation heat transfer, and two-phase slip between slurry droplets and the gas flow was derived and a computer code was written to implement this model. The STANJAN chemical equilibrium solver was coupled with this code to yield gas temperature, density, and composition as functions of axial location. Preliminary results indicate that the model is performing correctly, given current model assumptions. Radiation heat transfer in the combustion chamber is treated as an optically-thick participating media problem requiring a solution of the radiative transfer equation. A cylindrical P sub 1 approximation was employed to yield an analytical expression for chamber-wall heat flux at each axial location. The code exercised to determine the effects of secondary atomization intensity, defined as the number of secondary drops produced per initial drop, on chamber burnout distance and final Al2O3 agglomerate diameter. These results indicate that only weak secondary atomization is required to significantly reduce these two parameters. Stronger atomization intensities were found to yield decreasing marginal benefits. The experimental apparatus was improved to reduce building vibration effects on the optical system alignment. This was accomplished by mounting the burner and the transmitting/receiving optics on a single frame supported by vibration-isolation legs. Calibration and shakedown tests indicate that vibration problems were eliminated and that the system is performing correctly.

Mueller, D. C.↗
NASA NTRS1977

Combustion response modeling for composite solid propellants

A computerized mathematical model of the combustion response function of composite solid propellants was developed with particular attention to the contributions of the solid phase heterogeneity. The one-dimensional model treats the solid phase as alternating layers of ammonium perchlorate and binder, with an exothermic melt layer at the surface. Solution of the Fourier heat equation in the solid provides temperature and heat flux distributions with space and time. The problem is solved by conserving the heat flux at the surface from that produced by a suitable model of the gas phase. An approximation of the BDP flame model is utilized to represent the gas phase. By the use of several reasonable assumptions, it is found that a significant portion of the problem can be solved in closed form. A method is presented by which the model can be applied to tetramodal particle size distributions. A computerized steady-state version of the model was completed, which served to validate the various approximations and lay a foundation for the combustion response modeling. The combustion response modeling was completed in a form which does not require an iterative solution, and some preliminary results were acquired.

Source record↗
NASA NTRS1994

Ignition and combustion characteristics of metallized propellants, phase 2

Secondary atomization and ignition characteristics of aluminum/hydrocarbon gel propellants were investigated. Models of gel droplet shell formation were applied to aluminum/liquid hydrocarbon propellants to examine the effects of solid loading and ultimate particle size on the minimum droplet diameter permitting secondary atomization. A one-dimensional model of a gel-fueled rocket combustion chamber was developed. A model for radiant heat transfer from hot aluminum oxide particles to the chamber walls is included. A two-dimensional, two-phase nozzle code was used to estimate nozzle two-phase losses and overall engine performance.

Mueller, D. C.↗
NASA NTRS1994

Ignition and combustion characteristics of metallized propellants, phase 2

Experimental and analytical investigations focusing on aluminum/hydrocarbon gel droplet secondary atomization and its effects on gel-fueled rocket engine performance are being conducted. A single laser sheet sizing/velocimetry diagnostic technique, which should eliminate sizing bias in the data collection process, has been designed and constructed to overcome limitations of the two-color forward-scatter technique used in previous work. Calibration of this system is in progress and the data acquisition/validation code is being written. Narrow-band measurements of radiant emission, discussed in previous reports, will be used to determine if aluminum ignition has occurred in a gel droplet. A one-dimensional model of a gel-fueled rocket combustion chamber, described in earlier reports, has been exercised in conjunction with a two-dimensional, two-phase nozzle code to predict the performance of an aluminum/hydrocarbon fueled engine. Estimated secondary atomization effects on propellant burnout distance, condensed particle radiation losses to the chamber walls, and nozzle two phase flow losses are also investigated. Calculations indicate that only modest secondary atomization is required to significantly reduce propellant burnout distances, aluminum oxide residual size, and radiation heat losses. Radiation losses equal to approximately 2-13 percent of the energy released during combustion were estimated, depending on secondary atomization intensity. A two-dimensional, two-phase nozzle code was employed to estimate radiation and nozzle two phase flow effects on overall engine performance. Radiation losses yielded a one percent decrease in engine Isp. Results also indicate that secondary atomization may have less effect on two-phase losses than it does on propellant burnout distance and no effect if oxide particle coagulation and shear induced droplet breakup govern oxide particle size. Engine Isp was found to decrease from 337.4 to 293.7 seconds as gel aluminum mass loading was varied from 0-70 wt percent. Engine Isp efficiencies, accounting for radiation and two phase flow effects, on the order of 0.946 were calculated for a 60 wt percent gel, assuming a fragmentation ratio of five.

Mueller, D. C.↗
NASA NTRS1953

An Investigation of High-frequency Combustion Oscillations in Liquid-propellant Rocket Engines

An experimental investigation of high-frequency combustion oscillations (screaming) was conducted with a 100-pound-thrust acid-hydrocarbon rocket engine and a 500-pound-thrust oxygen-fuel rocket engine. The oscillation frequencies could be correlated as a linear function of the parameter C/L, where C is the experimentally measured characteristic velocity and L is the combustion-chamber length. The tendency of the engines to scream increased as chamber length was increased. With engine configurations that normally had a low efficiency, screaming resulted in increased performance; at the same time, a five to tenfold increase in heat-transfer rate occurred. It was possible, however, to achieve good performance without screaming.

COMBUSTION - EFFECT OF ENGINE OPERATING CONDITIONS↗
NASA NTRS1995

Stabilization Of Combustion Of Sprayed Fuel

Several modifications of nozzle spraying liquid propellant into combustion chamber proposed to stabilize combustion. Proposed changes in design of nozzle alter flow field in combustion chamber, according to fluid-mechanical principles, in such way as to suppress oscillations. Various alternative nozzle configurations include bluff or toroidal body to generate turbulence. Other features helping to suppress oscillations include down-stream recirculation zones, baffles, and damping cavities. Similar modifications help to suppress oscillations in industrial combustion chambers and in commercial and domestic oil-burning furnaces.

Voecks, Gerald E.↗