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Freestream
About: Freestream is a research topic. Over the lifetime, 3428 publications have been published within this topic receiving 56147 citations.
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TL;DR: In this article, the results of an experimental study of tangential supersonic slot injection with tangential subsonic injection through an additional slot above the main slot is presented. And the major conclusions drawn from the tandem injection results, are rst, that a near-separation boundary layer can be simulated in this way.
Abstract: The results of an experimental study of tangential supersonic slot injection into a supersonic airstream with tangential subsonic injection through an additional slot above the main slot is presented. Such a ow is of interest for at least two propulsion-related applications. First, one can simulate a near-separation boundary-layer pro le by subsonic injection through the upper slot, which is then to be re-energized by supersonic injection through the lower slot. Second, this ow is also of interest for application to fuel/ oxidizer/pilot injection from multiple overlaid slots. The experiments were performed in an intermittent, vacuum wind tunnel at a freestream Mach number of 2.85. The supersonic injectant had a Mach number of 2.00, and the subsonic injection was at Mach numbers of 0.26 and 0.72. The results are presented in the form of spark schlieren photographs, interferograms, and wall-static pressure measurements. Density pro les at several axial locations determined from the interferograms are presented, as well as streamwise and spanwise static pressure distributions. The major conclusions drawn from the tandem injection results, are rst, that a near-separation boundary-layer pro le can be simulated in this way. Next, tandemly injected subsonic and supersonic ow can be divided into separate components that closely resemble the respective individual injections into an undisturbed freestream. Also, the effects of the subsonic injection were completely mixed out at a downstream location of six slot heights. Therefore, adverse-pressuregradient-inducing devices should be positioned at least six slot heights downstream of the supersonic injection station if the effects of supersonic injection into a near-separation boundary layer for the purpose of re-energizing it are to be studied.
34 citations
TL;DR: In this article, it was shown that for a known discontinuity in upwash (as in the case of the control surface) the form, as well as the strength, of the singularity are determined uniquely.
Abstract: In the linearized formulation of the oscillating-surf ace problem, singularities in the lift distribution occur at subsonic leading edges, at control surface leading edges, and in general wherever the up wash prescribed by the wing deformations is discontinuous. These singularities are examined by use of the method of matched asymptotic expansions. It is shown that for a known discontinuity in upwash (as in the case of the control surface) the form, as well as the strength, of the singularity are determined uniquely. For subsonic leading edges only the form, but not the strength, of the singularity can be determined. A discussion is also given of the proper shape of the loading functions near side edges. Nomenclature b = reference length (root semi chord) Cp = pressure coefficient k = reduced frequency, ub/Um M = Mach number of freestream MN = M cos A, Mach number normal to edge p — pressure amplitude, Cp = peikt t = time Um = freestream velocity w = amplitude of prescribed upwash on the wing X) y} z = Cartesian coordinates with x in the freestream and in the span wise direction xc = location of control-surfa ce leading edge and hinge Xj y, z = stretched variables
33 citations
TL;DR: In this article, the secondary flow pattern developed at the exit cross-sectional plane was mapped out in detail by a three-dimensional velocity measurement technique using a 3D model of the transition duct.
Abstract: Experiments were made for three circular-to-rectangular transition ducts with different transition lengths at Reynolds numbers which ranged from 4 × 10 3 to 2 × 10 4 . The Reynolds number is based on the inlet boundary-layer thickness and a reference freestream velocity measured upstream of the transition duct. The secondary flow pattern developed at the exit cross-sectional plane was mapped out in detail by a three-dimensional velocity measurement technique
33 citations
TL;DR: In this article, the authors present mid-span measurements for a turbine cascade with active flow control, where steady blowing through an inclined plane wall jet has been used to control the separation characteristics of a high-lift low-pressure turbine airfoil at low Reynolds numbers.
Abstract: The paper presents mid-span measurements for a turbine cascade with active flow control. Steady blowing through an inclined plane wall jet has been used to control the separation characteristics of a high-lift low-pressure turbine airfoil at low Reynolds numbers. Measurements were made at design incidence for blowing ratios from approximately 0.25 to 2.0 (ratio of jet-to-local freestream velocity), for Reynolds numbers of 25000 and 50000 (based on axial chord and inlet velocity), and for freestream turbulence intensities of 0.4% and 4%. Detailed flow field measurements were made downstream of the cascade using a three-hole pressure probe, static pressure distributions were measured on the airfoil suction surface, and hot-wire measurements were made to characterize the interaction between the wall jet and boundary layer. The primary focus of the study is on the low-Reynolds number and low-freestream turbulence intensity cases, where the baseline airfoil stalls and high profile losses result. For low freestream turbulence (0.4%), the examined method of flow control was effective at preventing stall and reducing the profile losses. At a Reynolds number of 25000, a blowing ratio greater than 1.0 was required to suppress stall. At a Reynolds number of 50000, a closed separation bubble formed at a very low blowing ratio (0.25) resulting in a significant reduction in the profile loss. For high freestream turbulence intensity (4%), where the baseline airfoil has a closed separation bubble and low profile losses, blowing ratios below 1.0 resulted in a larger separation bubble and higher losses. The mechanism by which the wall jet affects the separation characteristics of the airfoil is examined through hot-wire traverse measurements in the vicinity of the slot.Copyright © 2004 by ASME
33 citations
TL;DR: In this paper, the authors examined the spanwise development of the interaction region and determined its dependence on leading edge geometry and the incoming-flow parameters, and showed that for blunt fin models, there is a region of interaction in the vicinity of the centerline where the scale and characteristics of the disturbed flowfield are controlled primarily by the leading edge diameter.
Abstract: In this paper, results are presented from an experimental study of fin-induced shock wave/turbulent boundary-layer interaction. Semi-infinite fin models, with sharp and hemicylindrically blunted leading edges, were tested at Mach 3 in two high Reynolds number, adiabatic wall, turbulent boundary layers. Detailed streamwise surface pressure distributions were measured at several spanwise stations for angles of attack from 0 to 12 deg. The objective was to examine the spanwise development of the interaction region and determine its dependence on leading edge geometry and the incoming-flow parameters. The results show that for blunt fins there is a region of the interaction in the vicinity of the centerline where the scale and characteristics of the disturbed flowfield are controlled primarily by the leading edge diameter. Outboard of this inner, or "leading-edge dominated," region the interaction properties and spanwise development are essentially the same as if the leading edge were sharp. Thus there is an outer region of the blunt fin-induced flowfield in which the properties are effectively "independent of leading edge blunting." Nomenclature D = blunt fin leading-edge diameter h - fin height Lu = upstream influence measured relative to the freestream shock wave M = Mach number P = static pressure P2 = wall pressure downstream of interaction Re = Reynolds number X = coordinate in the plane of the test surface and aligned with the tunnel axis, with X= 0 at the fin leading edge Xs = coordinate in the X direction measured relative to the freestream shock wave Y = coordinate normal to the X axis in the plane of the test surface, with Y= 0 at the fin leading edge
33 citations