Freestream

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

Planar Velocity and Temperature Measurements in Rarefied Hypersonic Flow Using Iodine LIF

Abstract: A planar laser-induced fluorescence (PLIF) technique is discussed and applied to measurement of time-averaged values of velocity and temperature in an I2-seeded N2 hypersonic free jet facility. A low temperature, non-reacting, hypersonic flow over a simplified model of a reaction control system (RCS) was investigated using the PLIF technique. Data are presented of rarefied Mach 12 flow over a sharp leading edge flat plate at zero incidence, both with and without an interacting jet issuing from a nozzle built into the plate. The velocity profile in the boundary layer on the plate was resolved. The slip velocity along the plate, extrapolated from the velocity profile data, varied from nearly 100% down to 10% of the freestream value. These measurements are compared with results of a DSMC solution. The velocity variation along the centerline of a jet issuing from the plate was measured and found to match closely with the correlation of Ashkenas and Sherman. The velocity variation in the oblique shock terminating the jet was resolved sufficiently to measure the shock wave thickness. I. Iodine fluorescence measurement technique Numerous methods have been reported in the literature for measurement of thermodynamic properties in supersonic flows using laser-induced fluorescence. This paper presents results of a technique based on narrowband excitation of fluorescence from nuclear hyperfine components in the B↔X electronic transition of molecular iodine. This approach was developed to study low temperature flows of I2-seeded N2 gas. A tunable single frequency laser beam is expanded and collimated into a thin sheet and passed into a low density wind tunnel flow, causing I2 molecules to fluoresce within a selected planar section of the flow-field. A series of images are taken of the fluorescence with a charge-coupled device (CCD) array as the laser frequency is incrementally tuned over selected lines in the I2 absorption spectrum. The fluorescence signal versus frequency is thereby obtained for any point in the plane of the laser sheet, within the resolution of the digital image. These absorption spectra can provide information on the local thermodynamic conditions in the flow. This paper describes planar laser-induced fluorescence (PLIF) measurements of velocity and temperature made in a low density, low temperature, hypersonic flow over a simplified model of a reaction control system (RCS) jet, 1 intended to support assessment of a hybrid numerical solution method for such flows. 2, 3

Pressure Fluctuation Measurements in the NASA Langley 20-Inch Mach 6 Wind Tunnel

Abstract: Experiments have been carried out in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel to measure the freestream pressure fluctuations, or tunnel noise, using a pitot rake. These experiments are part of an on-going effort to characterize the freestream disturbances of the Langley hypersonic wind tunnels along with other facilities around the country. Once the freestream disturbances have been characterized, a better understanding of the effect of these disturbances on boundary layer instability and transition measurements can be gained. The current experiments use a multi-probe pitot rake instrumented with both Kulite and PCB pressure transducers. Data were obtained over a range of Reynolds numbers and test section axial and radial positions. In general, noise levels were consistent spatially across the test section and ranged from 1% at the highest Reynolds numbers tested to approximately 1.6% at the lowest Reynolds number tested.

Design considerations of advanced supercritical low drag suction airfoils

Abstract: Supercritical low drag suction laminar flow airfoils were laid out for shock-free flow at design freestream Mach = 0.76, design lift coefficient = 0.58, and t/c = 0.13. The design goals were the minimization of suction laminarization problems and the assurance of shock-free flow at freestream Mach not greater than design freestream Mach (for design lift coefficient) as well as at lift coefficient not greater than design lift coefficient (for design freestream Mach); this involved limiting the height-to-length ratio of the supersonic zone at design to 0.35. High design freestream Mach numbers result with extensive supersonic flow (over 80% of the chord) on the upper surface, with a steep Stratford-type rear pressure rise with suction, as well as by carrying lift essentially in front- and rear-loaded regions of the airfoil with high static pressures on the carved out front and rear lower surface.

Comparison of computations and experimental data for leading edge vortices - Effects of yaw and vortex flaps

Abstract: Computations are presented using the conical Euler equations for swept delta wings with leading edge vortices. All the wings have sharp leading edges swept at 75 degrees to the freestream. In addition to an idealized flat plate model, geometrical features also included are thickness, centerbody, and two vortex flaps. Freestream Mach numbers of 1.7 to 2.8, angles of attack of 10 and 12 degrees, and angles of yaw of 0 and 8 degrees are considered. The computations are compared with pitot pressure traverses for one case. Other calculations are compared with pitot pressure traverses for one case. Other calculations are compared with surface pressure data and vapor screen pictures recently obtained at NASA Langley Research Center. The comparisons indicate that the dominant features of these flows are adequately modeled by the Euler equations, but viscous models are needed for the surface boundary layer and secondary separations.

Uncertainty Analysis of Mars Entry Flows over a Hypersonic Inflatable Aerodynamic Decelerator

Abstract: A detailed uncertainty analysis for high-fidelity flowfield simulations over a fixed aeroshell of hypersonic inflatable aerodynamic decelerator scale for Mars entry is presented for fully laminar and turbulent flows at peak stagnation-point heating conditions. This study implements a sparse-collocation approach based on stochastic expansions for efficient and accurate uncertainty quantification under a large number of uncertainty sources in the computational model. The convective and radiative heating and shear stress uncertainties are computed over the hypersonic inflatable aerodynamic decelerator surface and are shown to vary due to a small fraction of 65 flowfield and radiation modeling parameters considered in the uncertainty analysis. The main contributors to the convective heating uncertainty near the stagnation point are the CO2–CO2, CO2–O, and CO–O binary collision interactions, freestream density, and freestream velocity for both boundary-layer flows. In laminar flow, exothermic recombination rea...