Search

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.

The experimental investigation of the influence of the distributed heavy gas (sulfur hexafluoride of SF6) injection (into the wall layer of a boundary layer) on the hydrodynamic stability in relation to artificial (controlled) disturbances of the supersonic flat-plate boundary layer at free-stream Mach number M=2 have been performed. Controlled disturbances have been introduced into the boundary layer past the model by means of a harmonic point glow discharge disturbance generator. At different speeds of injection of sulfur hexafluoride, β-spectra (mean-square disturbance amplitudes depending on a longitudinal wave number β) measured at exciting frequency f=14 kHz, were defined and compared. It is shown (and for the first time experimentally for controlled disturbances) that such heavy-gas (sulfur-hexafluoride) injection leads to the boundary-layer stabilization, at least, at an exciting frequency.The experimental investigation of the influence of the distributed heavy gas (sulfur hexafluoride of SF6) injection (into the wall layer of a boundary layer) on the hydrodynamic stability in relation to artificial (controlled) disturbances of the supersonic flat-plate boundary layer at free-stream Mach number M=2 have been performed. Controlled disturbances have been introduced into the boundary layer past the model by means of a harmonic point glow discharge disturbance generator. At different speeds of injection of sulfur hexafluoride, β-spectra (mean-square disturbance amplitudes depending on a longitudinal wave number β) measured at exciting frequency f=14 kHz, were defined and compared. It is shown (and for the first time experimentally for controlled disturbances) that such heavy-gas (sulfur-hexafluoride) injection leads to the boundary-layer stabilization, at least, at an exciting frequency.

Experimental studies of the effect of a weak wave from the ledge 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 wave from the ledge during interaction with the flow in the vicinity of the leading edge of the model. The spectral and statistical analysis of the measured disturbances in the boundary layer was 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.

The paper discusses the experimental results of the laminar-turbulent transition investigations in the boundary layers at Mach 2 and 2.5, conducted in low noise supersonic flow conditions of T-325 wind tunnel of the ITAM SB RAS. The measuring problems of the transition in inhomogeneous boundary layers are discussed.

The results of an experimental study of the natural disturbance development in a flat plate boundary layer with a wavy surface are presented. The model of a flat steel plate with a blunt radius of the leading edge r = 0.34 mm was used. To create a wavy surface on the model, longitudinal roughness elements were applied. They were additionally coated with a film to align the side sections of the roughness and give them a rounded shape. For these experimental conditions, no influence of the surface waviness on the position of the laminar-turbulent transition in the flat plate boundary layer was detected.

We report a theoretical study of the properties of a supersonic boundary layer and its linear stability under conditions of surface material sublimation. Calculations were performed for an adiabatic boundary layer for a flat plate with a naphthalene coating at a free-stream Mach number of M = 3 (for the first instability mode disturbances). In the boundary layer, surface sublimation generates a binary mixture flow (air and foreign vapors). This flow is studied using local self-similar boundary layer equations, and it is shown that the rise in the flow stagnation temperature and the corresponding evaporation of the wall material cause significant wall cooling and an increase in the near-wall density of the binary mixture. This modification of the boundary layer profiles leads to a decrease in the disturbance amplification rates. This is confirmed by calculations based on linear stability theory (LST). Boundary-layer stabilization occurs with an increase in stagnation temperature. The influence of surface sublimation on the position of the laminar–turbulent transition was estimated by means of the LST-based e N method. The possibility of increasing the transition Reynolds number by application of the sublimation coating is demonstrated. The results of pilot boundary layer transition experiments performed in a hot-shot wind tunnel are reported. For the first time, a delay in the transition due to the application of a naphthalene coating was experimentally demonstrated. It is also shown that surface sublimation leads to an increase in the growth rates of the second and third instability modes for a Mach 8 boundary layer.

This study performed a joint theoretical and experimental investigation of the influence of distributed normal-to-the-surface and inclined heavy-gas injection into the near-wall sublayer of a boundary layer through a permeable wall on the laminar-turbulent transition (LTT). Sulfur hexafluoride (SF6) is used as a foreign gas for injection into the boundary layer. It also assessed stability in relation to both natural and artificial (controlled) disturbances of a supersonic flat-plate boundary layer at a free-stream Mach number (M) of 2. It is established, theoretically, that the action of a large molecular weight gas injection on the boundary layer is similar to the action of wall cooling and leads to an increase in boundary layer stability and LTT delay. The influence of injection on the position of transition is estimated by means of the eN method. Principally, the analysis shows the possibility of increasing the transition Reynolds number by means of SF6 injection. Controlled disturbances are introduced in the model boundary layer by means of a point harmonic glow-discharge disturbance generator and are measured by using a hot-wire anemometer. For the first time, it is shown experimentally that distributed injection of the heavy SF6 gas leads to boundary layer stabilization. This is mostly due to the reduction in growth rates of disturbances at higher frequencies, while the LTT shifted to higher Reynolds number values. Good qualitative agreement is achieved between the experimental data obtained with artificially generated disturbances and computations based on linear stability theory.

Experimental results are presented on the linear development of a wave train in a three-dimensional supersonic boundary layer on a swept wing with 45°-sweep angle at Mach number M = 2.0. Artificial disturbances in the boundary layer were excited by periodical glow discharge mainly at frequencies 10 and 20 kHz. The maximum of the controlled pulsation amplitude was localized in the boundary layer in normal to the swept-wing surface direction for spatial amplitude distributions as well as for the amplitude wave spectra. The asymmetry of the wave characteristics of the unstable disturbances in the boundary layer of a swept wing caused by the presence of crossflow is confirmed. The stability characteristics, including dispersion relations and amplification rates over a wide range of wave inclination angles, are obtained for disturbances with prescribed frequencies 10 and 20 kHz.