Freestream

About: Freestream is a research topic. Over the lifetime, 3428 publications have been published within this topic receiving 56147 citations.

Fluorescence velocimetry of the hypersonic, separated flow over a cone

Abstract: Planar laser-induced fluorescence of nitric oxide is used to measure a component of the velocity field for the Mach 7 flow around a 30-deg half-angle, 50-mm-diam cone mounted to a long, 38-mm-diam shaft, or sting. Transverse velocities are measured in the freestream, the shock layer, and the separated region at the junction between the cone and the sting. For most of the flowfield, the uncertainty of the measurements is between ±50 and ±100 m/s for velocities ranging from -300 to 1300 m/s, corresponding to a minimum uncertainty of ±5%. The measurements are compared with the commercial computational fluid dynamics (CFD) code CFD-FASTRAN . The agreement between the theoretical model and the experiment is reasonably good. CFD accurately predicts the size and shape of the shock layer and separated region behind the cone as well as the magnitude of the gas velocity near the reattachment shock. However, the magnitude of the velocity in the shock layer and gas expansion differ somewhat from that predicted by CFD. The discrepancies are attributed to a small systematic error associated with laser-beam attenuation and also to inexact modeling of the flowfield by CFD

Effectiveness of the Gas Turbine Endwall Fences in Secondary Flow Control at Elevated Freestream Turbulence Levels

Abstract: A secondary flow management technique which employs a boundary layer fence on the endwall of a gas turbine passage is evaluated under freestream turbulence conditions that are representative of turbine conditions. A turbulence generator, which was able to reproduce the characteristics of the combustor exit flow, was used. The horseshoe and passage vortices observed in previous tests with low turbulence level remain coherent and strong within the cascade passage when the intensity is elevated to 10 percent. A boundary layer fence on the endwall remains effective in changing the path of the horseshoe vortex and reducing the influence of the vortex on the flow near the suction wall at the high freestream turbulence level. The fence is more effective in reducing the secondary flow for the high turbulence case than for a low 11 case, probably because the vortex which has been deflected into the core flow diffuses and dissipates faster in the more turbulent flow. The fence decreases aerodynamic losses for streamlines within the core of the channel flow. NOMENCLATURE chord� (mm) specific heat of the air (kJ/kgK) total pressure coefficient. (Pt-Ptr)/(0.5pU.2) secondary kinetic energy coefficient, (v2+w2)/UO2 height of the fence� (mm) pressure side leg of horseshoe vortex suction side leg of horseshoe vortex heat transfer coefficient (W/m2K) power spectral density (e.g. u'2(f, df)/di), � (m2/s2) total pressure total pressure in freestream upstream of the cascade Reynolds number Reynolds number based on the chord length curvilinear distance from stagnation line along suction wall (mm)

Computational/experimental aeroheating predictions for X-33 Phase II vehicle

Abstract: Laminar and turbulent heating-rate calculations from an "engineering" code and laminar calculations from a "benchmark" Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase II flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent flow. For the highest angle of attack some of the measured heating data exhibited a "non-laminar" behavior which caused the heating to increase above the laminar level long before "classical" transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface.

Boundary-Layer Transition on a Flared Cone in the Texas A&M Mach 6 Quiet Tunnel

Abstract: Measurements of boundary layer transition location and boundary-layer profiles on a sharp-tipped 5o-half-angle flared cone were made in a low-disturbance Mach 6 wind tunnel. Following a recent nozzle repolishing effort, an updated flow quality characterization for the facility is presented for freestream unit Reynolds numbers of 8 and 10 × 10/m. At the high Reynolds number condition, transition onset on the flared cone model was observed to occur only very near its base, consistent with the previous observations on the same test article in the quiet tunnel’s former installation at NASA Langley. Uncalibrated boundary-layer profiles of mean and rms mass flux are obtained using constant temperature hotwire anemometry at several axial locations, and are compared with available historical data. Locations of maximum fluctuating content within the boundary layer are consistent with previous experiments and expectation, although boundary-layer thicknesses are observed to be 16% higher than previously, most likely due to a slightly heated-wall condition on the cone. Several imminent improvements to the current hotwire technique are discussed. This early work represents a key milestone toward the acquisition of highly-resolved, calibrated measurements of hypersonic transitional boundary layers to be used as code validation.

Sonic Injection into a Mach 5.0 Freestream Through Diamond Orifices

Abstract: An experimental study to characterize the near-field (x/d<8.0) flow for sonic air injection through 15-deg half-angle diamond-shaped orifices at four incidence angles (10, 27.5, 45, and 90 deg) and two total pressures (0.10 and 0.46 MPa) into a high Reynolds number (53 x 10 6 ) Mach 5.0 freestream was performed. A 90-deg circular injector, with the same exit port area and total pressures, was examined for comparative purposes. The experimental methods included surface oil flow visualization, shadowgraph photography, Mie scattering flow visualization, pressure-sensitive paint, and a pitot-cone five-hole pressure probe. Flowfield documentation, jet penetration, and shock-induced total pressure loss were derived from these data. Attachment of the interaction shock wave was found to depend on both incidence angle and injector pressure. Penetration correlations were developed for both the diamond and circular injectors. An approximate analysis indicated that the shock-induced total pressure loss decreased with decreasing incidence angle and injection pressure, and the largest losses were incurred by the 90-deg circular injector.