Showing papers on "Scramjet published in 1985"

Use of silane-methane mixtures for scramjet ignition

Abstract: Calculations have been made to demonstrate the use of silane-methan e liquid mixtures as an ignition promoter and flame stabilizer in hydrogen- (or other) fueled scramjet applications. The calculations show that mixtures initially nonflammable undergo rapid methane evaporation to produce drops capable of spontaneous inflammation. This technique may be expanded to other spontaneously flammable materials. It was assumed that the liquid silane-methane mixture was sprayed into the combijstion chamber upstream of the main fuelflow positions. Drop size, flow velocity, temperature, and initial liquid mixture composition were varied.

Numerical simulation of flow through scramjet inlets using a three-dimensional Navier-Stokes code

Abstract: A three-dimensional Reynolds-averaged Navier-Stokes code has been used to numerically analyze flow through a two-strut, supersonic combustion ramjet (scramjet) inlet configuration. It solves the governing equations in full conservation form using either the fully explicit or explicit-implicit method due to MacCormack. An algebraic two-layer eddy viscosity model is used for turbulent flow calculations. The analysis allows inclusion of end effects which are caused by the aft placement of the cowl on the underside of the inlet. A special grid has been developed to accommodate the struts embedded in the inlet flow field. Detailed numerical results are presented here for the two-strut configuration, and a comparison is made with the available experimental results.

Heat transfer in the vicinity of a large-scale obstruction in a turbulent boundary layer

Abstract: 'Cookson, R. A., "An Analysis of Non-Constant Area Heat Addition Due to Combustion in a Supersonic Air-Steam," U.S. Air Force Office of Scientific Research Interim Scientific Report AFOSRTR-75-1483,Sept. 1975. Pinckney, S. Z., "Integral Performance Predictions for Langley Scramjet Engine Module," NASA TM-X-74038, Jan. 1978. Ferri, A., Libby, P., and Zakkay, V., "Theoretical and Experimental Investigation of Supersonic Combustion," Proceedings, 3rd[Congress ICAS, 1962, MacMillan Press, London, 1964, pp. 10891156. Hsia, H. T. S., "A Criteria for the Combustion Modes in Constant Area Combustors," ASME Paper 70-WA/AV-4, 1970. 5 Henry, J. R. and Anderson, G. Y., "Design Considerations for the Airframe Integrated Scramjet," 1st International Symposium on Air Breathing Engine, Marseille, France, June 1972; also NASA TMX-2895,1973. Bussing, T. R. A. and Murman, E. M., "A One-Dimensional Unsteady Model of Dual Mode Scramjet Operation," AIAA Paper 83-0422, Jan. 1983. Billig, F. S. and Dugger, G. L., "The Interaction of Shock Waves and Heat Addition in the Design of Supersonic Combustors," Twelfth Symposium (International) on Combustion, The Combustion Institute, Pittsburgh, Pa., 1969, p. 1125. Waltrup, P. J. and Billig, F. S., "Prediction of Precombustion Wall Pressure Distributions in Scramjet Engines," Journal of Spacecraft and Rockets, Vol. 10, Sept. -Oct. 1973, pp. 620-622. Waltrup, P. J., Billig, F. S., and Stockbridge, R. D., "A Procedure for Optimizing the Design of Scramjet Engines," Journal of Spacecraft and Rockets, Vol. 16, May-June 1979, pp. 163-172. Waltrup, P. J., Dugger, G. L., Billig, F. S., and Orth, R. C., "Direct Connect Tests of Hydrogen-Fueled Supersonic Combustors," Sixteenth Symposium (International) on Combustion, The Combustion Institute, 1976, p. 1619. 11 Anderson, G. Y., Eggers, J. M., Waltrup, P. J., and Orth, R. C., "Investigation of Step Fuel Injectors for an Integrated Modular Scramjet Engine," CPIA Publication 281, Vol. Ill, pp. 175-190. Waltrup, P. J., Billig, F. S., and Evans, M. C., "Critical Considerations in the Design of Supersonic Combustion Ramjet (scramjet) Engines," Journal of Spacecraft and Rockets, Vol. 18, July-Aug. 1981, pp. 350-356.

Langley Mach 4 scramjet test facility

Abstract: An engine test facility was constructed at the NASA Langley Research Center in support of a supersonic combustion ramjet (scramjet) technology development program. Hydrogen combustion in air with oxygen replenishment provides simulated air at Mach 4 flight velocity, pressure, and true total temperature for an altitude range from 57,000 to 86,000 feet. A facility nozzle with a 13 in square exit produces a Mach 3.5 free jet flow for engine propulsion tests. The facility is described and calibration results are presented which demonstrate the suitability of the test flow for conducting scramjet engine research.

Scramjet sidewall burning: Preliminary shock tunnel results

Abstract: Experiments performed with a two dimensional model scramjet with particular emphasis on the effect of fuel injection from a wall are reported. Air low with a nominal Mach number of 3.5 and varied enthalpies was produced. It was found that neither hydrogen injection angle nor combustor divergence angle had any appreciable effect on thrust values while increased combustor length appeared to increase thrust levels. Specific impulse was observed to peak when hydrogen was injected at an equivalence ratio of about 2. Lowering the Mach number of the injected hydrogen at low equivalence ratios, less than 4, appeared to benefit specific impulse while hydrogen Mach number had little effect at higher equivalence ratios. When a 1:1 mixture by volume of nitrogen and oxygen is used instead of air as a test gas, it is found that hydrogen combustion is enhanced but only at high enthalpies.

Supersonic combustion of a silane/methane mixture

Abstract: A test program was conducted to evaluate the ignition and combustion characteristics of a 20 volume percent silane/methane mixture using a model scramjet combustor with Mach 2 combustor entrance conditions. The test gas total temperature was varied from 1200 to 3900 R. The mixture autoignited at a total temperature of 1650 R. This autoignition temperature can be contrasted with 2330 R for hydrogen and 1350 R for a 20 percent silane/hydrogen mixture in similar hardware. Methane without the silane additive did not autoignite in this configuration at temperatures as high as 3900 R.

Shock tunnel measurements of heat transfer in a model scramjet

Abstract: The results of heat transfer measurements to the walls of a two dimensional scramjet combustion chamber in a shock tunnel are presented. Thin film heat transfer gauges on a ceramic glass substrate were used. The range of experimental conditions covered produced boundary layers ranging from laminar to transitional, as was independently checked by flow visualisation. Empirical flat plate correlations, corrected for local pressure disturbances were used to make a comparison with the experimental results. In the fully laminar regime the heating rates were found to give approximate agreement with the empirical estimates. In the non laminar tests the heating rate is found to be well below the fully turbulent levels. It is not known at present if this is a transition effect, or if the pressure corrected flat plate turbulent correlations do not apply to the configuration used.

Investigation of parabolic computational techniques for internal high-speed viscous flows

Abstract: A feasibility study was conducted to assess the applicability of an existing parabolic analysis (ADD-Axisymmetric Diffuser Duct), developed previously for subsonic viscous internal flows, to mixed supersonic/subsonic flows with heat addition simulating a SCRAMJET combustor A study was conducted with the ADD code modified to include additional convection effects in the normal momentum equation when supersonic expansion and compression waves were present It is concluded from the present study that for the class of problems where strong viscous/inviscid interactions are present a global iteration procedure is required

Numerical investigation of internal high-speed viscous flows using a parabolic technique

Abstract: A feasibility study has been conducted to assess the applicability of an existing parabolic analysis (ADD-Axisymmetric Diffuser Duct), developed previously for subsonic viscous internal flows, to mixed supersonic/subsonic flows with heat addition simulating a SCRAMJET combustor. A study was conducted with the ADD code modified to include additional convection effects in the normal momentum equation when supersonic expansion and compression waves are present. A set of test problems with weak shock and expansion waves have been analyzed with this modified ADD method and stable and accurate solutions were demonstrated provided the streamwise step size was maintained at levels larger than the boundary layer displacement thickness. Calculations made with further reductions in step size encountered departure solutions consistent with strong interaction theory. Calculations were also performed for a flow field with a flame front in which a specific heat release was imposed to simulate a SCRAMJET combustor. In this case the flame front generated relatively thick shear layers which aggravated the departure solution problem. Qualitatively correct results were obtained for these cases using a marching technique with the convective terms in the normal momentum equation suppressed. It is concluded from the present study that for the class of problems where strong viscous/inviscid interactions are present a global iteration procedure is required.