Holger Babinsky

Microramp Control of Supersonic Oblique Shock-Wave/Boundary-Layer Interactions

The potential of microramp sub-boundary-layer vortex generators for flow control in supersonic engine inlets is investigated. In particular, the study focuses on the ability of these devices to beneficially affect oblique shock-wave/ boundary-layer interactions. Experiments have been conducted at Mach 2.5 to determine the nature of flow controlled by microramps and to investigate their ability to delay separation in a reflected shock interaction. Various ramp heights between 30 and 90% of the boundary-layer thickness were investigated. The details of the vortical flow generated by such devices were identified. The general flow features were found to scale with device height and it is suggested that smaller devices need to be placed closer to the expected adverse pressure gradients. When applied to a separated oblique shock-wave/boundary-layer interaction generated with a 7 degree wedge, microramps were not able to completely eliminate flow separation, although they were shown to break up separated regions. Other performance indicators across the shock-wave/boundary-layer interaction were also improved through the application of the devices.

Effect of microvortex generators on separated normal shock/boundary layer interactions

Experiments have been performed in a blowdown supersonic wind tunnel to investigate the effect of subboundary layer vortex generators placed upstream of a normal shock/turbulent boundary layer interaction at a Mach number of 1.5 and a freestream Reynolds number of 28 x 10 6 . The Reynolds number based on the inflow boundary layer displacement thickness was 26,000. Two types of subboundary layer vortex generators were investigated: wedge-shaped and counter-rotating vanes. It was found that the vane-type subboundary layer vortex generators eliminated and the wedge-type subboundary layer vortex generators greatly reduced the shock-induced separation. When placed in the supersonic part of the flow, both types of subboundary layer vortex generators caused a wave pattern consisting of a shock, reexpansion, and shock. The reexpansion and double shocks are undesirable features because they equate to increased total pressure losses. Furthermore there are indications that the vortex intensity is reduced by the normal shock/boundary layer interaction. Overall, the vane-type subboundary layer vortex generators were the more effective devices as they eliminated the shock-induced separation and had the least detrimental effect on the shock structure.

Shock wave-boundary-layer interactions

1. Introduction John K. Harvey and Holger Babinsky 2. Physical introduction Jean Delery 3. Transonic shock wave boundary layer interactions Holger Babinsky and Jean Delery 4. Ideal gas shock wave turbulent boundary layer interactions in supersonic flows and their modeling - two dimensional interactions Alexander A. Zheltovodov and Doyle D. Knight 5. Ideal gas shock wave turbulent boundary layer interactions in supersonic flows and their modeling - three dimensional interactions Doyle D. Knight and Alexander A. Zheltovodov 6. Experimental studies of shock wave/boundary layer interactions in hypersonic flows Michael S. Holden 7. Numerical simulation of hypersonic shock wave boundary layer interactions Graham V. Candler 8. Shock wave/boundary layer interactions taking place in hypersonic flows John K. Harvey 9. Shock wave unsteadiness in turbulent shock wave boundary layer interactions P. Dupont, J. F. Debieve and J. P. Dussauge 10. Analytical treatment of shock/boundary layer interactions George Inger.

Shock-wave/boundary-layer interaction control using three-dimensional bumps for transonic wings

Three-dimensional bumps have been developed and investigated on transonic wings, aiming to fulfill two major objectives of shock-wave/boundary-layer interaction control, that is, drag reduction and buffet delay. An experimental investigation has been conducted for a rounded bump in channel flow at the University of Cambridge and a computational study has been performed for a spanwise series of rounded bumps mounted on a transonic aerofoil at the University of Stuttgart. In both cases wave drag reduction and mild control effects on the boundary layer have been observed. Control effectiveness has been assessed for various bump configurations. A double configuration of narrow rounded bumps has been found to perform best, considerably reducing wave drag by means of a well-established X-shock structure with little viscous penalty and thus achieving a maximum overall drag reduction of about 30%, especially when significant wave drag is present. Counter-rotating streamwise vortex pairs have been produced by some configurations as a result of local flow separation. On the whole a large potential of three-dimensional control with discrete rounded bumps has been demonstrated both experimentally and numerically.

Unsteady Lift Generation on Rotating Wings at Low Reynolds Numbers

The rotating wing experiment is a fully three-dimensional simplification of the flapping-wing motion observed in nature. The spanwise velocity gradient and the wing starting and stopping acceleration that exist on an insectlike flapping wing are generated by the rotational motion of a finite-span wing. The flow development around a rotating wing at Re = 60, 000 has been studied using high-speed particle image velocimetry to capture the unsteady velocity field. Lift and drag forces have been measured for several different sets of wing kinematics and angles of attack. The lift curve shape was similar in all cases. A transient high lift peak, approximately 1.5 times the quasi-steady value, occurred in the first chord length of travel, and it was caused by the formation of a strong attached leading-edge vortex. This vortex then separated from the leading edge, resulting in a sharp drop in lift. As weaker leading-edge vortices continued to form and shed, lift values recovered to an intermediate value. The circulation of the leading-edge vortex has been measured and agrees well with the force data. Wing kinematics had only a small effect on the aerodynamic forces produced by the waving wing. In the early stages of the wing stroke, the velocity profiles with low accelerations affected the timing and the magnitude of the lift peak, but at higher accelerations, the velocity profile was insignificant.

Unsteady shock wave dynamics

An experimental study of an oscillating normal shock wave subject to unsteady periodic forcing in a parallel-walled duct has been conducted. Measurements of the pressure rise across the shock have been taken and the dynamics of unsteady shock motion have been analysed from high-speed schlieren video (available with the online version of the paper). A simple analytical and computational study has also been completed. It was found that the shock motion caused by variations in back pressure can be predicted with a simple theoretical model. A non-dimensional relationship between the amplitude and frequency of shock motion in a diverging duct is outlined, based on the concept of a critical frequency relating the relative importance of geometry and disturbance frequency for shock dynamics. The effects of viscosity on the dynamics of unsteady shock motion were found to be small in the present study, but it is anticipated that the model will be less applicable in geometries where boundary layer separation is more severe. A movie is available with the online version of the paper.

SBLI control for wings and inlets

AbstractFlow control can be applied to shock wave/boundary layer interactions to achieve two different goals; the delay of shock-induced separation and/or the reduction of stagnation pressure losses, which cause wave drag or inlet inefficiencies. This paper introduces the principles and main techniques for both approaches and assesses their relative suitability for practical applications. While boundary layer suction is already in wide use for separation control, the most promising novel device is the micro-vortex generator, which can deliver similar benefits to traditional vortex generators at much reduced device drag. Shock control is not yet used on practical applications for a number of reasons, but recent research has focused on three-dimensional devices which promise to deliver flow control with improved offdesign behaviour. Furthermore, there are some indications that a new generation of control devices may be able to combine the benefits of shock and boundary layer control and reduce shock-induced stagnation pressure losses as well as delay shock-induced separation.

How do wings work

The popular explanation of lift is common, quick, sounds logical and gives the correct answer, yet also introduces misconceptions, uses a nonsensical physical argument and misleadingly invokes Bernoullis equation. A simple analysis of pressure gradients and the curvature of streamlines is presented here to give a more correct explanation of lift.

Vortex Generators for a Normal Shock/Boundary Layer Interaction with a Downstream Diffuser

T HE interaction of a shock wavewith a turbulent boundary layer constitutes a fundamental problem of high-speed fluid mechanics. A detailed survey of past work on high-speed interactions has been carried out by Settles and Dolling [1] and Smits and Dussauge [2]. The shock interaction problem is particularly germane to the design of supersonic inlets. In such supersonic inlets, deceleration of the flow is achieved through a succession of oblique shock waves followed by a terminal normal shock. Boundary layers form on the inlet surfaces and interact with the shock system, giving rise to various shock/boundary-layer interactions (SBLIs). Each interaction of oblique/normal shock waves with the boundary layer causes stagnation pressure losses and downstream spatial distortions seen by the engine. An inlet must be carefully designed to minimize these losses and distortions during the compression process since they affect overall propulsion performance. In mixed-compression inlets, shock-induced separation can lead to engine unstart, which requires that the entire propulsion system undergo a restart sequence during flight. In external compression inlets, specifically axisymmetric configurations, a thick hubside boundary layer increases blockage and can decrease compressor performance. Thus, successfully controlling SBLIs has the potential to significantly improve supersonic inlet performance. As will be discussed in the following, various techniques of flow control for SBLIs have been proposed. However, it is often difficult to interpret the results because the flowfield may be too specific to an individual inlet configuration or too basic such that a relationship to inlet performance is not clear. To address this issue, a newwind-tunnel flowfield has been proposed [3] that captures much of the key shock boundary-layer interaction physics of supersonic external compression inlets. Thisflowfieldwill be used to study the novel flow control methods introduced herein. The conventionalflow control technique for SBLI conditions in an engine inlet employs a bleed of the boundary layer [4,5]. This bleed Presented as Paper 2010-4464 at the 40th AIAA Fluid Dynamics Conference and Exhibit, Chicago, IL, 28 June–1 July 2010; received 31 January 2011; revision received 26 July 2011; accepted for publication 11 August 2011. Copyright©2011 by theAmerican Institute ofAeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the copier pay the

Microvortex Generators Applied to Flowfield Containing a Normal Shock Wave and Diffuser

10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 0748-4658/12 and