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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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Hypersonic flow of air past circular cylinder with nonequilibrium oxygen dissociation, including dissociation of free stream
Gas turbine combustor liner film cooling for slot geometries in presence of high free stream turbulence
A study of the effect of free-stream thermal-energy release from shock-induced exothermic reactions on boundary-layer development and transition is presented. The flow model is that of a boundary layer developing behind a moving shock wave in two-dimensional unsteady flow over a shock-tube wall. Matched sets of combustible hydrogen-oxygen-nitrogen mixtures and inert hydrogen-nitrogen mixtures were used to obtain transition data over a range of transition Reynolds numbers from 1,100,000 to 21,300,000. The heat-energy is shown to significantly stabilize the boundary layer without changing its development character. A method for application of this data to flat-plate steady flows is included.
Revised solutions of the laminar-boundary-layer equations for cases which involved cooling at the wall combined with large pressure gradients in the main stream produced specific-weight-flow profiles which locally exceeded free-stream values. Heat-transfer and friction coefficients, boundary-layer thicknesses, and velocity, temperature, and specific-weight-flow distributions resulting from the revised solutions are presented for Euler numbers of 0.5 and 1, stream-to-wall temperature ratios of 2 and 4, and cooling-air flow rates through porous walls designated by flow parameters of 0, -0.5, and -1.
Hypersonic inviscid flow past a circular cylinder with nonequilibrium oxygen dissociation and dissociation of the free stream
Use of photoionization in measuring velocity profile of free stream flow in expansion tube
Wedge and conical shaped probes for the measurement of free-stream flow quantities at supersonic speeds have been tested in both wind tunnel and flight. These probes have improved capabilities over similar ones used in the past. Through the use of miniature pressure sensors, that are located inside the probes, they are able to provide instantaneous measurements of a time-varying environment. Detailed herein are the results of the tests in NASA Langley Researcher Center s Unitary Plan Wind Tunnel (UPWT) at Mach numbers of 1.6, 1.8 and 2.0, as well as flight tests carried out at the NASA Dryden Flight Research Center (DFRC) on its F-15 aircraft up to Mach numbers of 1.9. In the flight tests the probes were attached to a fixture on the underside of the F-15 fuselage. Problems controlling the velocity of the flow through the conical probe, required for accurate temperature measurements, are noted, as well as some calibration problems of the miniature pressure sensors that impact the accuracy of the measurements.
Surface cooling and heating effects on low speed laminar boundary layer gas flow with constant free stream velocity
Systematic numerical analysis of two dimensional and axisymmetric laminar jet of incompressible fluid with and without free stream
Strongly accelerated turbulent boundary layer, investigating free stream turbulence effect on heat transfer
The paper describes an investigation of the velocity fluctuations in the free stream above an incompressible turbulent boundary layer developing at constant pressure. Measurements were performed in a wind tunnel with an rms level of axial velocity fluctuations of about 0.2 percent. The possibility of assessing the contributions to the rms level of the velocity fluctuations without using the high-pass filtering technique is demonstrated.
Skin friction and heat transfer on flat plate in high velocity laminar motion with mass transfer for any free stream and injected gas combination
Nozzle wall hypersonic turbulent boundary layers at free stream Mach number using pitot, hot wire, and wall fluctuation and static pressure measurements
Flow field from forward facing nozzles exhausting near large cylindrical body at free-stream Mach numbers of 3.0 and 6.0
Experimental and theoretical studies of low speed leading edge boundary layer receptivity to free-stream vorticity produced by upstream wires normal to the leading edge are discussed. Data include parametric variations in leading edge configuration and details of the incident disturbance field including single and multiple wakes. The induced disturbance amplitude increases with increases in the leading edge diameter and wake interactions. Measurements agree with the theory of M. E. Goldstein.
Influence of interaction of flow field with boundary layer on heat transfer in vicinity of interior corner aligned with free stream velocity
A spike-type nose inlet with sharp-lip cowl was investigated at a free-stream Mach number of 2.5 with water injection in its 16-inch diameter, 11-foot-long subsonic diffuser section. Inlet total temperature of exit with liquid-air ratios of about 0.04 with no apparent change in the critical pressure recovery. The observed temperature drops were less than the theoretically predicted values, and the amount of water evaporated was 35 to 50 percent less than that theoretically possible.
Laminar boundary-layer heat transfer and shear stress predictions from similarity solutions, emphasizing large free-stream velocity gradients and highly cooled walls