Experimental observation of transition behavior on a flat plate

In studying transition behavior a shock tube and tunnel were used to produce high temperatures, and thin-film platinum heat gauges were used to measure local heat flux as well as to detect the transition of the laminar boundary layer over a flat plate and a cone. Initial investigations were conducted in the hypersonic shock tunnel to obtain high-temperature information for the development of an ICBM nose cone. Shock Mach numbers as large as 50 with a temperature of 15,000 K after the incident wave were produced in the driven tube. Shock tubes are used to investigate the heat transfer over various surfaces to 2500 K for the development of future gas turbines

Investigation to optimize the passive shock wave/boundary layer control for supercritical airfoil drag reduction

The passive shock wave/boundary layer control for reducing the drag of 14%-thick supercritical airfoil was investigated in the 3 in. x 15.4 in. RPI Transonic Wind Tunnel with and without the top wall insert at transonic Mach numbers. Top wall insert was installed to increase the flow Mach number to 0.90 with the model mounted on the test section bottom wall. Various porous surfaces with a cavity underneath were positioned on the area of the airfoil where the shock wave occurs. The higher pressure behind the shock wave circulates flow through the cavity to the lower pressure ahead of the shock wave. The effects from this circulation prevent boundary layer separation and enthropy increase hrough the shock wave. The static pressure distributions over the airfoil, the wake impact pressure survey for determining the profile drag and the Schlieren photographs for porous surfaces are presented and compared with the results for solid surface airfoil. With a 2.8% uniform porosity the normal shock wave for the solid surface was changed to a lambda shock wave, and the wake impact pressure data indicate a drag coefficient reduction as much as 45% lower than for the solid surface airfoil at high transonic Mach numbers

Investigation to optimize the passive shock wave-boundary layer control for supercritical airfoil drag reduction

The optimization of passive shock wave/boundary layer control for supercritical airfoil drag reduction was investigated in a 3 in. x 15.4 in. Transonic Blowdown Wind Tunnel. A 14% thick supercritical airfoil was tested with 0%, 1.42% and 2.8% porosities at Mach numbers of .70 to .83. The 1.42% case incorporated a linear increase in porosity with the flow direction while the 2.8% case was uniform porosity. The static pressure distributions over the airfoil, the wake impact pressure data for determining the profile drag, and the Schlieren photographs for porous surface airfoils are presented and compared with the results for solid-surface airfoils. While the results show that linear 1.42% porosity actually led to a slight increase in drag it was found that the uniform 2.8% porosity can lead to a drag reduction of 46% at M = .81

Sharp flat plate heat transfer in helium at Mach numbers of 22.8 to 86.8 and in corner flow with air at Mach number of 19

Surface heat transfer rates were measured on a sharp flat plate at zero angle of attack in a hypersonic shock tunnel. The density and leading edge Knudsen number were varied to span the continuum to near free molecule regimes. The strong interaction parameter varied from 11 to 16,000 with Knudsen numbers from 0.56 to 17.1 respectively. Local heat transfer rates in the corner flow region produced by the intersection of two perpendicular flat plates with sharp leading edges were determined for various flow densities. The strength of the shock wave from the vertical plate was varied by adjusting the angle of attack from 0 to 5 deg. The unit Reynolds number varied from 1,000 to 17,200 and the Knudsen numbers from 1.6 to 27. The strong interaction parameter varied from 14 to 500

Heat transfer on a flat plate in helium at Mach numbers 67.3 and 87.6 and in hypersonic corner flow with air at Mach number of 19

Hypersonic heat transfer rates on flat plates in helium and in corner flow region with ai

Heat Transfer on a Flat Plate in Continuum to Rarefied Hypersonic Fows at Mach Numbers of 19.2 and 25.4

Surface heat transfer rates measured on flat plates in hypervelocity shock tunne

Mean and fluctuating velocity contours and acoustic characteristics of subsonic and supersonic jets

Shadowgraph photographs of subsonic and supersonic jet

Subsonic and supersonic jet flow and acoustic characteristics and supersonic suppressors

To study the similarities and differences between subsonic and supersonic jets, velocity and impact pressure fluctuations were determined along the axis over a jet Mach number range of 0.6 to 1.4 for a 2 in. diameter convergent nozzle and for a one inch diameter jet flow. Static pressure distribution fluctuations due to shear and turbulence in the jet flow for subsonic and supersonic jets were related to acoustic radiation to the far field. Also determined were flow and acoustic characteristics of a single shroud, and multiple shroud tube and shroud suppressors for supersonic and subsonic exhaust velocities. A compressor consisting of 191 tubes and 191 shrouds decreased the primary Mach number drastically for both jet Mach numbers of 1.4 and 0.7; rms impact and static pressure fluctuations on the axis were also reduced from values existing for an equivalent area single nozzle

Flow and acoustic characteristics of subsonic and supersonic jets from convergent nozzle

Acoustic and flow characteristics of subsonic and supersonic jets from convergent nozzle