Jean Delery

Aspects of vortex breakdown

Abstract Vortex breakdown is of primary importance in many situations met in aeronautical as well as extra-aeronautical applications. During the past 40 years, this problem has been the subject of a large number of investigations, both in the experimental and theoretical domains. This article presents an overview of some of the results obtained by these studies. Carefully executed visualizations and velocity measurements have allowed a clear description of the field associated with breakdown. These experiments executed in a vortex tube, where the vortex is confined, or in arrangements where it is free, show that the existence of a stagnation point forming on the centre-line of the vortical structure is a distinctive feature of breakdown. Different types of breakdown (spiral and bubble types) seem to exist, however this point is subject to controversy. At the same time, the vortex undergoes a brutal dilatation with the generation of large scale fluctuations. Similar features are observed when the breakdown occurs over a delta wing. The two main parameters influencing the phenomenon are the adverse pressure gradient to which the vortex is submitted and its swirl intensity which cannot go beyond a critical value. The most popular theories for vortex breakdown belong to four main classes: the quasi-cylindrical approach and analogy to boundary layer separation, solution of the axisymmetric Navier-Stokes equations, the concept of the critical state and hydrodynamic instabilities. More advanced models, based on the bifurcation theory or using a direct numerical simulation of the problem have also been developed. In most cases, the obtained results are in good agreement with experimental observations, but the predictive capability of these theories is still limited.

Shock wave/turbulent boundary layer interaction and its control

Abstract The interaction between a shock wave and a boundary layer often leads to extremely detrimental effects, especially if the shock is strong enough to separate the boundary layer. When this happens, there occurs a rapid growth of the dissipative region along with a dramatic intensifying of turbulent fluctuations with the frequent occurrence of buffeting. In the present review some fundamental properties of the interaction are first considered for a two-dimensional, adiabatic flow developing on a flat surface. The following features are first examined: upstream interaction length, incipient shock-induced separation and evolution of the boundary layer properties during the interaction process. Supersonic and transonic flows are both considered. In a second part, the means envisaged to control shock/boundary layer interactions are presented, with emphasis on the physics of phenomena involved in this process. Control methods can be classified into two categories: those acting on the boundary layer properties before it enters the shock region (e.g. wall cooling, wall mass transfer, upstream blowing) and those having a local action in the shock foot region itself (e.g. suction or injection, boundary layer removal, passive control). The most appropriate technique depends on the specific application under consideration. Finally, some calculation methods are briefly presented, most of them being restricted to laminar interactions.

Research into vortex breakdown control

Abstract Vortex breakdown remains a significant and intriguing phenomenon that can have detrimental or beneficial effects, depending on the application. Thus there is a strong need to both better understand the phenomenon and to control it, either to prevent breakdown or to promote it. For the past 50 years, multiple flow control techniques have demonstrated the ability to manipulate the vortex breakdown location over slender delta wings at high angles of attack. An extensive historical review of these diverse control methods, mechanical and pneumatic, steady or periodic, is presented and discussed; however, none of these techniques has clearly demonstrated a superior efficiency or effectiveness in controlling either the vortical flow structure or the vortex breakdown location. Each technique, does, on the other hand, provide a unique approach to the control of the vortex breakdown depending on the desired outcome. There are still major obstacles to overcome before the control of vortex breakdown is implemented in flight. For example, oscillations of the vortex breakdown locations are difficult to quantify and to identify. The often poor effectiveness of control techniques can be in great part attributed to insufficient knowledge of breakdown and in an inability to accurately predict breakdown. When considering the large quantity of studies aimed at vortex breakdown control and their relative success, it is clear that decisive progress in this domain will require further basic investigations to clearly elucidate the physics of the phenomenon and to improve the predictive capability.

Oscillation of Vortex Breakdown Location and Blowing Control of Time-Averaged Location

Abstract : The goal of this research is the control of leading-edge vortex breakdown location utilizing along-the-core blowing near the apex on the leeward surface of sharp-edged, slender, delta wings at high angles of attack. In the S2Ch subsonic wind tunnel at ONERA Chalais Meudon, two delta wing models with 70-deg sweep angles and root chords of 950 mm have been configured to collect qualitative and quantitative surface and flowfield data. First, an examination of the streamwise, time-dependent oscillation of the leading-edge vortex breakdown locations without active flow control is presented. These results further the understanding of the vortex breakdown phenomena and provide a more precise basis for evaluating the effectiveness of various flow control methods. Second, open-loop blowing along one of the vortex cores on the leeward surface of the delta wing demonstrates the ability to displace downstream the controlled, time-averaged, vortex breakdown location by 20% of the root chord.

Interference Between a Cylindrical Bow Shock and a Plane Oblique Shock

Hypersonic flows are prone to intense shock waves whose intersection, or interference, gives rise to a system of waves and slip surfaces that can have a large influence on the aerodynamics of a vehicle. It is convenient to distinguish six types of interference associated with very distinct flow structures, depending on the intensity and relative direction of the intersecting shock waves. Among these classes of interference, those of types III and IV are the origin of shear layers or supersonic jets whose impact on a nearby surface creates potentially destructive peaks of pressure and heat flux locally. Type III and IV interference and the corresponding heat transfer distributions were investigated experimentally. The model consisted of a cylindrically blunted plate and a wedge serving as an oblique shock generator. These experiments were carried out in a short-duration wind tunnel at Mach numbers 6 and 16 in air and at Mach number 6.6 in carbon dioxide. The Reynolds number based on the plate bluntness diameter was varied in the range from 2.2 x 10 4 to 1.6 x 10 6 . The influence of the impinging shock location on the interference heat transfer was carefully investigated

Methodes de calcul des ecoulements turbulents decolles en supersonique

Abstract This paper presents a general review of methods at present available for treating problems of supersonic turbulent separated flows. The essential physical characters of these phenomena are first described in detail, on the example of a step. Then we examine the two theoretical avenues of research most frequently used to treat that kind of problem, viz. on the one hand, the overall methods based on the Chapman-Korst flow model; on the other hand, the integral techniques. The study of the overall methods is centred on the theories calling upon the concept of reattachment angle criterion; it is illustrated by many applications concerning widely different cases. The integral methods are first presented in a general way; we then discuss the problems raised by the resolution of the differential system describing the phenomenon (subcritical or supercritical behaviour of the solution) and several examples are presented. Lastly, we show how insufficient may be the theoretical models making use of the Prandtl boundary layer equations.

Code verification/validation with respect to experimental data banks

Abstract The past 40 years have known a spectacular development of CFD capabilities. It is now possible to compute complex three-dimensional unsteady flows even at the design stage by solving the Unsteady Averaged Navier–Stokes Equations (URANS approach) and progress are made every day in still more advanced approaches such as LES and DNS. However, the confidence in CFD methods is still limited because of uncertainties in the numerical accuracy of the codes and of the inadequacy of the turbulence models they use. Thus, there is still a need for well made and well documented experiments to validate the codes and to help in their improvement. Such experiments must also fulfil quality criteria to be considered as safe enough and really useful for code validations. The article presents a discussion of the strategy to be followed to ensure the reliability and accuracy of a code by placing emphasis on the experimental aspects of code validation. The purpose is illustrated by considering recent examples of CFD validation operations based on basic – or building block – experiments. The first case considers an experiment on a purely laminar shock wave/boundary layer interaction used to assess the numerical accuracy of several codes. Other examples deal with the crucial problem of the validation of turbulence models in strongly interacting flows. The conclusion stresses the importance to constitute high quality data banks on typical flows still difficult to predict. The problem of data dissemination is also briefly addressed.

A Study of Turbulence Modelling in Transonic Shock-Wave Boundary-Layer Interactions

Since validation studies of turbulence models require a great number of comparisons with experimental data, solving the boundary-layer equations is an inexpensive numerical tool by which this validation can be performed. This approach has been used extensively to test six turbulence models applied to transonic shock-wave/boundary-layer interactions: the algebraic models of Michel, Quemard and Durant, of Alber, of Baldwin and Lomax and of Johnson and King, the [k, e] transport equation model and the ASM.

A basic experimental investigation of passive control applied to a transonic interaction

Abstract Passive control applied to a turbulent shock wave/boundary layer interaction has been investigated by considering a two-dimensional channel flow. The field resulting from application of passive control has been probed in great detail by using a two-component laser Doppler velocimetry system to execute mean velocity and turbulence measurements. Four different perforated plates have been considered, as also the solid wall reference case. The performed measurements have shown that passive control deeply modifies the inviscid flowfield structure, the unique strong shock being replaced by a lambda shock system. This fractionning of the compression induces a substantial reduction of the wave drag associated with the interaction. On the other hand, the combined injection-suction effect taking place in the control region provokes an important thickening of the boundary layer. There results an increase of the friction drag which nearly outbalances the gain in wave drag. A determination of the total drag in the control region was made. It was found that passive control induced a modest decrease of this drag compared to the solid wall case. Also, the rugosity of the holes is an important source of drag (excrescence drag) which contributes to compromise the potential benefit of the passive control technique.

A detailed experimental analysis of the flow in a highly loaded fixed compressor cascade: the iso-cascade co-operative programme on code validation

Abstract The ISO-Cascade programme, involving IHI, Snecma and ONERA, has been undertaken to perform a thorough experimental investigation of a highly loaded compressor cascade in order to validate Navier–Stokes solvers used in turbomachinery bladings design. The aim of this study was to reproduce in a two-dimensional fixed geometry the main physical features of the flow in a real compressor, including strong viscous effects. The flow produced in the ISO-Cascade has been qualified by surface flow visualisations, schlieren pictures, surface pressure and skin friction measurements, surveys by multi-hole pressure probes and extensive explorations with a three-component LDV system. The large amount of results thus obtained has allowed a consistent description of the flow topology and the constitution of an unrivalled data bank to validate computer codes.