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The analysis of the efficiency of turbulent boundary-layer control on low-speed airfoils using mass transfer through a permeable wall is presented. Particular attention is paid to the consideration of the physical properties of the wall flow in the presence of positive (from the wall) and negative (to the wall) mass transfer and the aero-dynamic characteristics of airfoils under conditions of distributed or localized action of mass transfer of varying intensity on the boundary layer.

Interaction of an external artificial disturbance with a blunted leading edge of a swept wing is studied under the conditions of a model experiment. Specific features of disturbances of a three-dimensional boundary layer arising due to significant bluntness of the leading edge and transverse flow are identified. It is shown that localized disturbances generated in the incident flow interact with the boundary layer on the swept wing and generate unsteady streaky structures; high-frequency wave packets are formed near these streaky structures. The dynamics of the development of wave packets and localized streaky structures in a three-dimensional boundary layer in a gradient flow is quantified.

The experimental results on heat transfer in the gas coolant flow in a space formed by a dense packing of seven heated tubes are presented. Eight separating inserts with longitudinal displacers were used to fix the tubes and ensure a uniform gas flow field in the internal and external channels of the tube bundle. Gas mixtures with a large difference in the Prandtl number were used as the working fluid: air (Pr = 0.7) and helium-xenon mixture (Pr = 0.23). The experiments were carried out in the range of Reynolds numbers of 2218 - 12900.

The experimental data on the effect of a pair of weak shock waves on the flow in the boundary layer of swept plates with initial angles χ = 35° and 40° at Mach number M = 2 are presented. The incident flow was disturbed using a weak shock wave (SW) generator made in the form of a two-dimensional sticker on the side wall or on the nozzle surface in the test section of the wind tunnel. For the latter case, shadow visualization of the flow past the models was carried out and the inclination angles of the weak SWs were determined. Measurements with a constant temperature hot-wire anemometer allowed us to record for the first time the effect of a weak SW from the leading edge of the sticker on the flow in the boundary layer of a flat plate with large sweep angles of the leading edge. On the model with χ = 35°, in the vicinity of the maximum effect of a pair of weak SWs, the flow characteristics were measured with a continuous change in the rotation angle of the model. The measurement results suggest that the sweep angle χ = 48° is the critical sweep angle of the blunt leading edge, at which longitudinal vortices are not generated in the boundary layer by a “co-directional or overtaking” weak shock wave. The conclusions of previous studies that with an increase in the sweep angle along the leading edge, there is a decrease in the intensity of the effect of weak shock waves on the flow in the boundary layer, and that flow turbulization occurs for a model sweep angle of 50°, have been confirmed.

The experimental results on thermal destruction processes during pyrolysis and the degree of fragmentation caused by pulsed aerodynamic impact on a demonstration sample (DS) simulating a separable structural element of the Soyuz-2 launch vehicle at the launch site have been obtained. A polymer based on aramid fibers and epoxy binder is used as the structural material of the DS. The process of pyrolysis and subsequent thermal destruction occur due to heating the DS with an electric heater. A pulsed aerodynamic impact takes place in a vacuum chamber due to a gas-dynamic blow of compressed air with an assessment of the fragmentation degree. An experimental program and a corresponding test rig with the metrological support have been developed. Based on the obtained results, a further direction of research on thermal destruction and aerodynamic fragmentation of DS has been mapped out.

The flow pattern and integral flow parameters in confusors implementing different methods of a supersonic flow compression are numerically investigated.. Model configurations are considered, consisting of a tapering inlet section and a constant cross-section, in which different types of flows take place: two-dimensional, axisymmetric or three-dimensional flow. The flow is completely turbulent and computations are based on Navier-Stokes equations and the k-ω SST turbulence model. The flow in the confusor is assumed to be supersonic everywhere, the range of free-stream Mach numbers is М_H = 2 - 4.5.

The paper presents the experimental results on the distribution of the outer wall and coolant temperatures in a simplified model of the peripheral region of a fuel assembly (FA) consisting of two parallel cells. The experimental setup was equipped with three fuel element simulators with a diameter of 10 mm and a length of 500 mm. The power of each fuel element simulator was up to 2000 W (4 kW/m). During operation, the fuel elements were connected to the power source in turn, which allowed determination of the features of temperature distribution in the model with non-uniform energy release. A movable thermocouple and an IR camera were used for measurements. The data obtained for a water coolant and a lead-bismuth alloy are compared. It is shown that the main patterns of temperature distribution are generally similar for different coolants.

The processing of experimental dependences for pitching vs. time curve for free oscillations of a cone with the attached hemisphere rear part gives the pitching moment coefficients, equivalent coefficients for pitch damping, spectral coefficients for pitching angle and pitching moment coefficient for self-induced oscillations flow mode. It was shown that a dependency of pitch damping vs. oscillation amplitude presents a hyperbolic behavior. The frequency spectrum for pitching and pitching moment coefficient demonstrates a domination of one frequency. There is no hysteresis in a dependency of the pitching moment on the attack angle for self-induced oscillations.

A low-parameter equation has been obtained to describe the viscosity coefficient of liquid and gaseous neon at temperatures from 24.6 to 700 K and pressures from 0.044 to 50 MPa. This equation allows data obtaining within the experimental error. It is shown that this equation, proposed for calculating the viscosity coefficient of liquid and gas, allows reliable extrapolation beyond the studied area.

The paper presents characteristics of coherent structures that develop in the shock wave / boundary layer interaction zone. Experiments were performed on a flat plate for two options on the free flow: laminar and turbulent flow modes. The shock wave was generated by a wedge-type body placed ahead the plate. The data are obtained from hot-wire velocimetry technique. Flow structures in a boundary layer were studied using the coherence spectral analysis. It was demonstrated that the transversal size of coherent patterns varies insignificantly in the interaction zone between shock waves and boundary layer.