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The paper describes a digital signal processing technique for determining the relationship between disturbances in a supersonic flow and pulsations of the boundary layer of a flat plate model. Estimates of the error of the proposed data processing method are given, the results of an experiment conducted to demonstrate the method on real data are presented.

An experimental study of the effect of longitudinal vortices generated by weak shock waves on the laminar-turbulent transition in the boundary layer of a flat plate at a Mach number 2 is conducted. To amplify the effect of the longitudinal vortices generation, a model with leading edge blunting radius of 0,5 mm was used. The disturbance intensity was measured with a constant temperature hotwire anemometer and a single-wire sensor. To change the Reynolds numbers in the flow, the method of varying the pressure in the pretest section of a supersonic wind tunnel was used. According to the experimental conditions, the level of mass flow perturbations in the test section T325 of ITAM SB RAS did not exceed 0,1 % of the mean flow in the frequency range from 0,2 to 50 kHz. In this work, it is obtained for the first time that stationary streamwise vortices that arise in the boundary layer of a flat plate because of the action of a pair of weak shock waves on its leading edge lead to an early laminar-turbulent transition at different Reynolds numbers. In the case of a homogeneous boundary layer, the transition is achieved at Rex≈ 3,1 × 106, and in measurements in the region of vortices from shock waves at Rex≈ 1,5 × 106and Rex≈ 2 × 106.

A study of the effect of distributed gas with high specific heat capacity blowing into the supersonic boundary layer on its stability to artificial disturbances was conducted. It is shown that the injection of gas with high heat capacity Cp at certain flow values can lead to the stabilization of the supersonic boundary layer.

The influence of external vortex disturbances on the position of the laminar-turbulent transition in boundary layers on swept wings at the Mach number M = 2 is experimentally studied. The experiments are performed with two wing models; the sweep angles of the leading edges are 45° and 72°. The vortex disturbances are generated by a wire placed ahead of the nozzles in the subsonic part of the flow. It is found that such disturbances increase the amplitude and change the spectral composition of oscillations in the boundary layer; moreover, they can induce earlier turbulization of the flow.

The results of experimental studies on the impact of periodic flow modulation on the development of natural disturbances and artificial localized wave packets in the boundary layer of a swept wing with a leading edge swept angle of 40 degrees at a Mach number M = 2.5 are presented. The boundary layer was modulated using periodic roughness on the model surface. Artificial wave packets were generated by a pulsed glow discharge. Hot-wire anemometer measurements showed that, according to the experimental conditions, periodic modulation of the boundary layer leads to the stabilization of the development of controlled localized disturbances and to an increase in the growth of natural disturbances.

An experimental fluctuation analysis is usually conducted via thermal anemometry with hot-wires. However, in the vicinity of oscillating shock waves, this kind of sensor can be destroyed due to strong mechanical loads. On the basis of a sufficient modelling, a wedge-shaped hot-film can be suitable for a quantitative fluctuation analysis across a shock wave. Within this study, the modal analysis according to Kovasznay and Morkovin was adapted from hot-wires to the used wedge hot-film. Additionally, 3-dimensional fluctuation diagrams were derived for the separate as well as for mixed modes. The freestream fluctuations were detected across an oblique shock wave with a wedge hot-film in a constant-temperature mode. The shock was caused by a ramp with a 10° ramp angle placed in a flow with a Mach number of M = 2.5 and a unit Reynolds number of Reunit = 4.96·106·m−1. The recorded perturbations were decomposed according to the modal analysis and found to be dominated by the acoustic mode. The fluctuations’ amplification across the shock wave and the subsequent decay could clearly be detected. They are in good agreement with the literature.

The impact of weak shock waves on the boundary-layer flow of a flat blunt plate installed along the stream was experimentally studied for different sweep angles of the leading edge of the plate at Mach number 2. The shock waves in the incoming flow in the form of an N-wave were generated by 150×7×0.13-mm two-dimensional roughness provided on the side wall of the T-325 wind tunnel. It was confirmed that, when an N-wave was incident onto the plate leading edge with zero sweep angle, generation of longitudinal vortices accompanied by an increase in the pulsation level and by a change in the spectral composition of perturbations, was observed in the boundary layer. Anemometric data showed that, when the sweep angle of the leading edge changed from 0 to 25 degrees under the impact of the “catching-up” incident weak shock wave, the widening of the vortex in the supersonic boundary layer occurred.

The paper presents the experimental results and the results of direct numerical simulation of the development and interaction of two wave trains from two point sources of controlled disturbances in a supersonic boundary layer on a flat plate with an incident flow Mach number of 2.5. Sources were located parallel to the leading edge of the model. For the introduction of controlled disturbances into the boundary layer, the normal component of the mass flow rate was varied in the calculations. In the experiment, periodic glow discharges at a frequency of 20 kHz were used. In both cases, the disturbances sources worked synchronously. Mass flow rate pulsations were measured and recorded in sections, parallel to the leading edge of the model, near the maximum of disturbances along the boundary layer. In the experiment, a constant-temperature hot-wire anemometer was used. After performing a discrete Fourier transform, the spatial distributions of disturbances, the beta-spectra were determined, and the wave characteristics of the development of disturbances downstream were estimated. In addition, direct numerical simulation of the downstream development of disturbances from a single source was performed. The work presents a comparison of experimental and theoretical calculated data. The paper discusses the effects inherent in the interaction of unstable traveling controlled disturbances from two sources operating synchronously.

The paper is devoted to an experimental and theoretical study of effect of moderate angle-of-attack variation on disturbances evolution and laminar-turbulent transition in a supersonic boundary layer on swept wing at Mach 2. Monotonous growth of the transition Reynolds numbers with angle of attack increasing from −2° to 2.7° is confirmed. For the same conditions, calculations based on linear stability theory are performed. The experimental and computational results show a favourable comparison.

The results of experimental studies of the effect of the duration and current of a pulsed glow discharge on the generation of localized disturbances in the supersonic boundary layer of a flat plate at Mach numbers of the oncoming flow M = 3 and 4 have been presented. Flow pulsation measurements were performed by a hot-wire anemometer. It has been found that an increase in the duration of the impulse function leads to an increase in the amplitude and time scales of the perturbations that are generated in the boundary layer. An increase in the current of the pulse discharge leads to an increase in the amplitude of the generated pulsations. The propagation velocities in the boundary layer of localized perturbations at different Mach numbers of the oncoming flow have been estimated.

The generation of two weak shock waves by a two-dimensional roughness mounted on the side wall of the test section of a wind tunnel and their influence on the supersonic boundary layer of a flat blunt plate is experimentally investigated at the Mach number 2.5. The measurements are made using a constant temperature hot-wire anemometer. In the boundary layer of the plate the profiles of the mean mass flow rate and the r.m.s. fluctuations are measured in the region of streamwise vortices generated by the pairwise weak shock wave upon the interaction with the flow in the vicinity of the leading edge of the model. High-intensity fluctuations generated by the influence of the weak shock wave are found to exist.

Experimental study of the effect of a weak shock wave from the protuberance of two-dimensional roughness installed on the side wall of the test section of the wind tunnel on the supersonic boundary layer of the blunted flat plate at the Mach number 2.5 was carried out. The measurements were performed by a constant temperature hot-wire anemometer in the region of stream wise vortices generated by the shock wave from the protuberance during interaction with the flow in the vicinity of the leading edge of the model. The spectral and statistical analyses of the measured disturbances in the boundary layer were carried out. The amplitude-frequency spectra of mass flow pulsations and statistical diagrams of the measured disturbances in the supersonic part of the boundary layer were obtained.

Experimental investigation of the influence of the distributed heavy gas (sulfur hexafluoride, SF6) injection into the near-wall sub-layer of a boundary layer) on the hydrodynamic stability in relation to controlled disturbances of the supersonic flat-plate boundary-layer at free-stream Mach number M = 2 have been performed. It was for the first time in controlled experiments obtained that injection of this foreign gas leads to the boundary-layer stabilization that is manifested in reduction of disturbance amplification rates. Good quantitative agreement of collected experimental data with computations based on the linear stability theory is obtained.

Experimental study of the influence of flow pulsation in settling chamber on the supersonic free stream disturbances was carried out. Data on the pulsations in the settling chamber and the efficiency of deturbulization system as well as the correlation of pulsations of the flow of settling chamber and flow pulsations in test section of T-325 supersonic wind tunnel of ITAM SB RAS were obtained.

The paper presents results of a theoretical study on the properties of a supersonic boundary layer and its linear stability under conditions of surface material sublimation. Calculations were performed for an adiabatic-wall boundary layer over a camphor coated flat-plate model at a free-stream Mach number of M=2. In the boundary layer, surface sublimation generates a binary mixture flow of air and foreign vapors. This flow is studied using boundary layer equations in approximation of local self-similarity. It is shown that increase of the flow stagnation temperature leads to the wall material sublimation. Increasing evaporation causes significant wall cooling and an increase in the near-wall density of the binary mixture. Such modification of the boundary layer profiles leads to a decrease of the amplification rates of the first mode perturbations, which were computed by means of the linear stability theory. It is concluded that application of a camphor coating facilitates the supersonic adiabatic-wall boundary-layer stabilization with increasing flow stagnation temperature.

The theoretical and experimental investigation of the effect of the heavy gas (sulfur hexafluoride) blowing into the boundary layer on the hydrodynamic stability to controlled perturbations of the supersonic flat-plate boundary layer at Mach number 2 have been carried out. Artificial perturbations have been entered in the model’s boundary layer by using a harmonic point glow discharge perturbation source. It is obtained that heavy gas blowing stabilizes the supersonic boundary layer. In experiments at the excitation frequency f = 14 kHz, unit Reynolds number Re1 = 3×106 m−1 and streamwise coordinate x = 90 mm, it was obtained that for gas injection rate Q = 0.12 g/cm2 there is a stabilization of perturbations for all spanwise wave numbers, i.e. the complete stabilization of perturbations – for all wave numbers, the spatial disturbance amplification rates are less than zero.

This work is devoted to the results of an experimental study of the laminar-turbulent transition in the boundary layer of the wing with a sweep angle of the leading edge of 72 degrees. Experiments were performed in a supersonic low noise wind tunnel T-325 ITAM SB RAS at Mach number 4. The measurements were carried out using a hot wire anemometer. Dependences of mass flow pulsations on Rex are obtained, according to which the position of the laminar-turbulent transition was obtained. Amplitude-frequency spectra were obtained and spectral analysis of disturbances was carried out.

The paper is devoted to the main results of experimental studies of the evolution of wave trains in supersonic boundary layers generated by a high-frequency glow discharge. The activity has been conducted out under grants of Russian Foundation of Basic Research (RFBR) and in the framework of budget financing from the mid-90s of 20th century and until recently. Mainly it is concerned to the flat plate supersonic boundary layer with the sharp leading edge.

This article is devoted to the experimental study of perturbations in a modulated swept-wing boundary layer at Mach 2. The modulation of the boundary layer was created using periodic roughness of the surface of the experimental model. The mean flow and pulsations were measured using a hot-wire anemometer. Spectral analysis of the disturbances showed that modulation of the boundary layer can lead to the stabilization of the development of disturbances in a wide frequency spectrum.

An experimental study of the correlation of pulsations of free supersonic flow and the pulsations of the boundary layer of a flat plate with a sharp leading edge was performed using a new multi-channel hot-wire measurement system. The experiments were performed in a low-noise wind tunnel T-325 of the Khristianovich Institute of Theoretical and Applied Mechanics of the Siberian Branch of the Russian Academy of Sciences at the Mach number M=2. Spatial-temporal distributions of the correlation coefficient for cross-sections of the boundary layer parallel to the leading edge were obtained.

The paper is devoted to an experimental study of laminar-turbulent transition in supersonic boundary layers on swept wings. The experiments were fulfilled in the low nose supersonic wind tunnel T-325 of Khristianovich Institute of Theoretical and Applied Mechanics on the models of a swept wing with a swept angle of 72° at Mach 2 which corresponds to a subsonic leading edge. The transition location was determined using a constant temperature hot-wire anemometer. A comparison of the influence of small angles of attack on the transition position with a model with a supersonic leading edge is shown. The transition to a wing model with a subsonic leading edge occurs earlier than on a wing model with a supersonic leading edge. The angle of attack has a significant effect on the position of the transition point from the laminar to turbulent flow. The data presented in the article are among the first results for a swept wing with a subsonic leading edge.

This paper is devoted to the results of an experimental study of heat transfer in the boundary layer of a flat plate with the impact of weak shock waves on the leading edge. The experimental study was conducted on the model of blunt flat plate at Mach 2. To generate weak shock waves into the flow, a roughness on the lower part of the nozzle of the test section of wind tunnel was used. Visualization experiments of heat flows were carried out using an infrared camera. As a result of experiments, data on the features of heat transfer on the surface of a flat plate were obtained. A double longitudinal structure in the form of a temperature change defect caused by the interaction of two vortices with the model surface is shown. Temperature changes in the vortex region downstream in a direction parallel to the leading edge have a similar structure, but different amplitudes.

Wedge-shaped hot-films are a promising alternative to hot-wires for supersonic free flow disturbance measurements if hot-wires cannot be used due to harsh flow conditions. Up to the present, wedge hot-films could not serve as quantitative tools because of an insufficient modelling of the substrate’s influence on the thermal balance. The present paper shows a static calibration of a wedge-shaped hot-film sensor that is based on a combination of an analytical, a numerical and an experimental approach. The substrate’s impact was analysed by the help of CFD simulations and modelled by a newly introduced substrate factor. The obtained sensitivities were discussed and qualitatively explained compared with standard hot-wires. A quantitative comparison of experimentally detected mass flux fluctuations proved the wedge hot-film data to match similar hot-wire results very well. The presented approach has the potential to upgrade hot-films of various shapes to tools for quantitative fluctuation measurements.

An experimental study of the influence of unit Reynolds number on the position of laminar-turbulent transition in swept-wing boundary layer at supersonic flow velocities was carried out. In the experiments, a swept-wing model with 3-% circular arc airfoil and 45-deg gliding angle of wing edges was used. The position of the transition was iden-tified using a hot-wire anemometer. It was found that at М = 2 and 2.5, an increase in the unit Reynolds number (Re1) leads to a transition delay. It was shown that an increase in freestream Mach number and in the level of flow pulsations in the wind-tunnel test section leads to a less pronounced influence of Re1 on the transition position. At a high noise level due to the growth of Mach number or due to the introduction of vortical disturbances, no effect due to unit Reyn-olds number on the transition position was observed.

The evolution of a controlled broadband wave packet in flat-plate and swept-wing supersonic boundary layers was experimentally investigated at Mach number M = 2. The wave packet was introduced into the boundary layer by a localized pulse electrical discharge. The structure and evolution downstream of the wave packet were studied by hot-wire measurements. It was found that the wave packet has a symmetric shape in a flat-plate boundary layer, whereas there is asymmetry in case of a swept-wing one. The spectral analysis of the development of different modes of the wave packet was provided.

Experiments were performed on the influence of distributed injection of a heavy gas (elegas SF6) into the near-wall region of the supersonic (freestream Mach number М∞ = 2) boundary layer on its stability in relation to natural disturbances. Heavy gas injection for the case of linear development of disturbance field results in boundary layer stabilization. It was experimentally proved that the elegas injection can suppress disturbances at the frequencies higher than 15 kHz for the tested range of the streamwise coordinate.

The effect of distributed tangential injection of heavy SF6 gas into the wall zone of a boundary layer on the supersonic flow stability and laminar-turbulent transition is experimentally and theoretically investigated at the freestream Mach number M∞ = 2. For the first time it is experimentally shown that in the case of this injection laminar-turbulent transition is downstream displaced.

The recovery ratio of a wedge-shaped hot-film probe was determined in an experimental as well as numerical study, since this information is still unpublished and essential for using the probe in hot-film anemometry. The experiments were conducted at the Khristianovich Institute of Theoretical and Applied Mechanics (ITAM) in Novosibirsk, Russia, and the simulations were performed with StarCCM+, a commercial 2nd order finite volume code. In the analysis, the Mach number was varied between M = 2 and M = 4, and the unit Reynolds number ranged from Re1 = 3.8•106 to Re1 = 26.1•106 m−1, depending on the Mach number. During the experiment, the stagnation temperature was kept constant for each Mach number at a separate value in the range of T 0 = 289 ± 7 K. Three different stagnation temperatures were used in the simulations: T 0 = 259 K, T 0 = 289 K, and T 0 = 319 K. The difference between the experimental and the numerical results is ≤ 0.5 %, and, therefore, both are in very good accordance. The influence of the Mach number, of the unit Reynolds number, and of the stagnation temperature was analysed, and three different fitting functions for the recovery ratio were established. In general, the recovery ratio shows small variations with all three tested parameters. These dependencies are of the same order of magnitude.

Time-and-space localized disturbances (wave packets), created by pulsed electric discharges in a supersonic boundary layer, were visualized using hot-wire anemometry. Experimental studies were performed on a flat plate model with a sharp leading edge at Mach number M = 2. The evolution of localized artificial disturbances generated by discharges of different duration and power was studied. The velocity of propagation of the wave packets was also determined. Contour lines of the wave packet in the plane (z, t)