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Experimental data on stability of a three-dimensional supersonic boundary layer on a swept wing are presented. The experiments are performed on a swept wing model with a lenticular profile with a 40 • sweep angle of the leading edge at a zero angle of attack. The supersonic boundary layer on the swept wing was laminarized with the use of distributed roughness. A pioneering study of interaction of traveling and stationary disturbances is performed. Some specific features of this interaction are identified. The main reason for turbulence emergence in a supersonic boundary layer on a swept wing is demonstrated to be secondary crossflow instability.Introduction. The problem of turbulence emergence and development of methods of transition control in three-dimensional boundary layers has been studied by many researchers. Creating a small commercial supersonic aircraft of the new generation is supposed to involve new technologies, in particular, passive control of the laminarturbulent transition (flow laminarization) in the boundary layer with the use of microscale roughness distributed over the surface of a swept wing near the leading edge.It was shown [1, 2] that instability of the boundary layer on a swept wing can be controlled with the use of distributed roughness at subsonic flow velocities. The method of passive control has the following features. Microscale roughness elements are applied parallel to the leading edge of the wing, at a distance equal to 1-5% of the wing chord. Based on calculation results for the most unstable stationary mode, the spanwise step between the roughness elements is chosen, which should be approximately (0.50-0.55)λ st (λ st is the wavelength of the most unstable stationary mode in the direction parallel to the leading edge of the wing). The roughness elements used in [1, 2] were cylinders 6 μm high, which were located near the leading edge of the swept wing. A change in the spanwise distance between the roughness elements was found to affect the position of the laminar-turbulent transition region. For instance, the use of distributed roughness with a 12-mm step along the wing span (or with a step multiple to this value) made the transition region approach the leading edge approximately by 35%, while the use of an 8-mm step increased the laminar flow region by 11%.Distributed roughness was first used for passive control of the transition in a supersonic boundary layer on a swept wing in the experiments [3,4] with the use of the method developed for subsonic velocities [1,2]. Saric and Reed [3,4] reported that they used surface microroughness to delay the transition to turbulence in a three-dimensional boundary layer on a wing model with a subsonic leading edge. For a supersonic leading edge, the boundary layer was observed to remain laminar on the entire model. Even the use of roughness elements with a step λ st did not lead to boundary layer tripping. It should be noted that the values of the transition Reynolds numbers given in [3,4] are overrated by an order of magnitude. Zuccher et...

532.526The instability of three-dimensional boundary layers has become an area of intensive theoretical and experimental research in the last decade [1, 2]. Such boundary layers are encountered in particular in the flow of air past swept wings on aircraft in regions of negative and positive longitudinal pressure gradients.Nearly all of the studies have been conducted for subsonic flow velocities. However, despite the large volume of research, the problem is still not fully resolved: there are qualitative differences between not only the theoretical and experimental data, but also the experimental results obtained by different groups of researchers. The complexity of the problem lies in the fact that together with known Tollmien-Schlichting (TS) waves -the formation of which leads to the transition to turbulence in two-dimensional boundary layers -additional instability waves connected with secondary flows develop in the three-dimensional case. These flows are stationary vortices whose axes are directed mainly along the streamlines of the external flow, as well as some travelling (non-TS) waves. The formation of all of these waves and their relative role in the transition depends heavily on the external conditions.Only two experiments have been conducted at supersonic velocities [3, 4]. The models in them were circular cylinders positioned at a 45 ~ angle to the incoming flow at Math numbers M = 10 and 3.5, respectively. The turbulence transition was studied on the basis of measurements of certain integral quantities -without measurement of the development of processes in the boundary layer. Stationary vortices were detected by visualization on the surfaces.The goal in this investigation is to use measurements made inside a boundary layer to establish if wave processes that are similar to the subsonic case occur in flow about a supersonic airfoil with a sharp leading edge under the "natural" conditions of operation of a wind tunnel. To determine the flow parameters (M**, Re 1, T, U,.), we measured pressure in the afterburner chamber and static pressure in the working part with balances. Stagnation temperature was measured with a thermoeouple. A description of this system was presented in [7]. Measurents of the fluctuation characteristics of the flow during the experiment were performed using instruments and programs developed to measure the transient parameters of a supersonic flow (supersonic boundary layer) [7].Pulsations were measured using a constant-resistance hot-wire anemometer [7] with sensors made of tungsten wire 6 #m in diameter. The length of the sensor wire was 1.2 mm. The sensor was mounted on a holder inserted in the guide bar of