Boundary Layer Removal Research Articles

Evaluation of Three-Dimensional Transonic Codes Using New Correlation-Tailored Test Data

A comprehensive research program was conducted for the specific purpose of acquiring benchmark test data suitable for evaluations of three-dimensional transonic codes. High-quality data were obtained for three advanced technology wings by using a unique test apparatus and by devoting careful attention to details of the experiment. The test apparatus included provisions for removal of the wind tunnel boundary layer and measurements of far-field pressures. The test data were used in preliminary evaluations of three selected transonic computational methods. For these limited evaluations, nonconservative formulations more closely predicted measured pressures than conservative formulations. In addition, full potential solutions were found to be uniformly in better agreement with experiment than small disturbance results.

Flow Computations in Inlets at Incidence Using a Shock Fitting Bicharacteristic Method

An analysis is presented for calculating the flowfield in supersonic mixed-compression aircraft inlets operating at angle of attack. The flowfield is computed by a steady three-dimensional bicharacterist ic method. The bow shock wave and the reflected internal shock wave system are computed by a three-dimensional, discrete shock wave fitting procedure. Viscous and thermal diffusion may be included as source terms in the bicharacteristic method. A production type of computer program capable of determining the flowfield in a variety of axisymmetric mixed-compression supersonic inlets is available. The results of the present analysis agree well with those produced by the two-dimensional method of characteristi cs when axisymmetric flowfields are computed. For three-dimensional flowfields, the results of the present analysis agree well with experimental data except in regions of high viscous interaction and boundary-layer removal. The present analysis does not compute the boundary layer, nor does it account for boundary-layer bleed.

Reduction of fan noise by annulus boundary layer removal

Much of the noise produced by a fan is due to varying forces caused by the unsteady flow field through which it passes. In the absence of inlet guide vanes, support struts and other mechanical obstructions, the flow irregularities are caused by large scale intake turbulence and temporal and spatial variations in the annulus boundary layer. In many cases the most effective source is the annulus boundary layer, as this interacts with the blade tip which is the fastest part of the blade and is therefore the most effective noise source. This noise can be reduced if the annulus boundary layer is smoothly bled away from the rotor so that the blade tip sees a thin, uniform boundary layer. A one metre diameter ventilation fan has been run in a duct with and without a boundary layer bleed system and a reduction of far field sound power of approximately 5 dB has been obtained for a bleed of 5 % of the main flow with some far field tone reductions of more than 15 dB. Measurements of far field directivities, in-duct acoustic modes and aerodynamic distortions have confirmed the suggested explanation for this phenomenon. The tests have shown that great care must be taken in the design of a bleed system so that residual boundary layer distortions are not in the correct wavelength range to produce propagating acoustic waves at important blade passing harmonic frequencies. If such distortions are present, the far field noise can increase instead of decrease.

Investigation of Propeller Vortex Noise Including the Effects of Boundary Layer Control

An experimental investigation has been conducted on the vortex noise produced by a two-bladed 4-ft diam model propeller capable of boundary layer removal. The propeller had a spinner comprising 70% of the total propeller radius. A porous section on both surfaces of the symmetric section airfoil allowed removal of the boundary layer. Free-field measurements were made in an anechoic chamber at three field points for three tip speeds (209.5, 314.2, and 366.5 ft/sec) and four blade angles (0°, 2.5°, 5°, and 10°) both without and with boundary layer control. Agreement with theory was good (within 2 dB) showing a sixth power of tip velocity relationship and a classical dipole radiation pattern for the over-all SPL of the vortex noise. Boundary layer removal primarily affected sound levels above 3150 Hz, with no ordered effect on the over-all level.

Effect of boundary-layer removal on high velocity flame stabilization

Effect of boundary-layer removal on high velocity flame stabilization

U.S. Patent Specifications

An aircraft has a fuselage and a plurality of lifting surfaces, each of the lifting surfaces comprise a portion fixedly attached to the fuselage and a portion pivotally mounted on the fixed portion for angular displacement about a spanwise axis, a power means mounted on the pivoted portion. The power means includes a compressor. The pivotally mounted portion is provided with a boundary layer removal slot and duct, means leading from the compressor to the slot for removing the boundary layer air.