Experimental study of nonlinear processes in a swept-wing boundary layer at the mach number M=2

Results of experiments aimed at studying the linear and nonlinear stages of the development of natural disturbances in the boundary layer on a swept wing at supersonic velocities are presented. The experiments are performed on a swept wing model with a lens-shaped airfoil, leading-edge sweep angle of 45°, and relative thickness of 3%. The disturbances in the flow are recorded by a constant-temperature hot-wire anemometer. For determining the nonlinear interaction of disturbances, the kurtosis and skewness are estimated for experimentally obtained distributions of the oscillating signal over the streamwise coordinate or along the normal to the surface. The disturbances are found to increase in the frequency range from 8 to 35 kHz in the region of their linear development, whereas enhancement of high-frequency disturbances is observed in the region of their nonlinear evolution. It is demonstrated that the growth of disturbances in the high-frequency spectral range (f > 35 kHz) is caused by the secondary instability.     

Effect of Heat Supply on the Stability of the Supersonic Swept Atiachment-Line Boundary Layer

The stability of a boundary layer with volume heat supply on the attachment line of a swept wing is investigated within the framework of the linear theory at supersonic inviscid-free-stream Mach numbers. The results of numerical calculations of the flow stability and neutral curves are presented for the flow on the leading edge of a swept wing with a swept angle χ=60° at various free-stream Mach numbers. The effect of volume heat supply on the characteristics of boundary layer stability on the attachment line is studied at a surface temperature equal to the temperature of the external inviscid flow. It is shown that in the case of a supersonic external inviscid flow volume heat supply may result in an increase in the critical Reynolds number and stabilization of disturbances corresponding to large wave numbers. For certain energy supply parameters the situation is reversed, the unstable disturbances corresponding to the main flow-instability zone are stabilized but another zone of flow-instability with small wave numbers and a significantly lower critical Reynolds number appears.     

Experimental study of evolution of disturbances in a supersonic boundary layer on a swept-wing model under controlled conditions

Experimental data on stability of a three-dimensional supersonic boundary layer on a swept wing are presented. Evolution of artificial wave trains was studied. The experiments were conducted for Mach numberM=2.0 and unit Reynolds numberRe ₁=6.6·10⁶m⁻¹ on a swept-wing model with a lenticular profile and a40° sweep angle of the leading edge at zero incidence. Excitation of high-frequency disturbances caused by secondary-flow instability at a high initial amplitude was observed. It is shown that the evolution of disturbances at frequencies of10, 20, and30 kHz is similar to the development of travelling waves for the case of subsonic velocities. Institute of Theoretical and Applied Mechanics, Siberian Division, Russian Academy of Sciences, Novosibirsk 630090. Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 41, No. 1, pp. 50–56, January–February, 2000.     

Linear stability of three-dimensional boundary layers

Stability of compressible three-dimensional boundary layers on a swept wing model is studied within the framework of the linear theory. The analysis based on the approximation of local self-similarity of the mean flow was performed within the Falkner-Skan-Cooke solution extended to compressible flows. The calculated characteristics of stability for a subsonic boundary layer are found to agree well with the measured results. In the case of a supersonic boundary layer, the results calculated for a Mach number M = 2 are also in good agreement with the measured spanwise scales of nonstationary vortices of the secondary flow. The calculated growth rates of disturbances, however, are substantially different from the measured values. This difference can be attributed to a high initial amplitude of disturbances generated in the experiment, which does not allow the linear stability theory to be applied. The evolution of natural disturbances with moderate amplitudes is fairly well predicted by the theory. The effect of compressibility on crossflow instability modes is demonstrated to be insignificant. __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 49, No. 2, pp. 3–14, March–April, 2008.     

Linear evolution of controlled disturbances in the supersonic boundary layer on a swept wing

The linear development of controlled disturbances in the three-dimensional supersonic boundary layer on a swept model wing with a sharp leading edge is experimentally investigated at the Mach number 2. The spatial-temporal and spectral-wave characteristics of the wave train of unstable disturbances are obtained. The asymmetry of these characteristics, due to the secondary flow in the three-dimensional boundary layer, is confirmed.     

Experimental study of laminar-turbulent transition on a swept wing with local boundary-layer suction at subsonic velocities

The effect of local suction of the boundary layer gas near the leading edge of the upper surface of a swept wing (with the sweep anglex=35°) on the extent of the laminar flow region is studied on a large half-model at subsonic velocities. Various experimental methods were used. The data are subjected to a comparative analysis. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 4, pp. 65–73, July–August, 1998.     

Influence of intense surface cooling on hypersonic viscous flow past a delta wing

The influence of intense surface cooling on the parameters of a laminar boundary layer flow on a thin delta wing in a hypersonic viscous perfect-gas stream is studied for the strong viscous-inviscid interaction regime. The effect of the power-law shape of the wing cross-section and the wing thickness to boundary-layer displacement thickness ratio on the local and total aerodynamic characteristics is numerically investigated. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 4, pp. 57–64, July–August, 1998.     

Stability and Receptivity of the Swept-Wing Boundary-Layer Flow Controlled by a Periodic Sequence of Plasma Actuators

Fluid Dynamics - The process of flow control in the boundary layer on a swept wing using a span-periodic sequence of plasma actuators mounted at an angle to the leading edge is modeled. The...     

Receptivity of the Swept Wing Boundary Layer to a Steady Flow Inhomogeneity

The boundary layer on a plate with an inclined blunt leading edge is investigated for a free-stream flow with a small span-periodic velocity inhomogeneity. This flow simulates the penetration of the outer turbulence into the swept wing boundary layer. It is shown that the boundary layer perturbations generated by the inhomogeneity generally have a streamwise velocity component significantly greater than the initial inhomogeneity amplitude. The dependence of the perturbations on the distance from the leading edge and the spanwise inhomogeneity period is found. It is shown that the swept wing boundary layer is less sensitive to the perturbation type in question than the straight wing boundary layer.     

Vortex diffraction on a swept wing

The forces acting on a swept wing in the presence of a vortex induced by a delta wing, as well as the velocity field in the vicinity of the swept wing, have been measured. By means of the “frozen,” vortex model and a specially-developed numerical panel method, the forces and moments acting on the wing are calculated from the known velocity field. Comparison of the calculated and measured force characteristics makes it possible to determine the extent to which the model fits the physical flow pattern. It is shown that for the intense vortex considered in this study the model gives results which disagree sharply with the experimental data. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 6, pp. 98–105, November–December, 1998. The study was supported by the International Scientific and Technical Center under grant No. 201.     

Nonlinear effect of external low-frequency acoustics on eigen-oscillations in a supersonic boundary layer

A method of simulation and results of numerical calculation of the evolution of hydrodynamic disturbances in a supersonic boundary layer on a flat plate under the influence of external acoustic waves at Reynolds numbersRe=220–640 and Mach numberM=2 are described. The solution is constructed by the method of expansion with respect to the small parameter; the contribution of linear and quadratic terms to the solution is taken into account. The method developed allows one to estimate the admissible level of the acoustic field, which does not affect the development of eigen-oscillations in the boundary layer. Institute of Theoretical and Applied Mechanics, Siberian Division, Russian Academy of Sciences, Novosibirsk 630090. Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 40, No. 5, pp. 99–105, September–October, 1999.     

Heat transfer and supersonic flow structure in surface cavities

The characteristics of the flow and heat transfer in two- and three-dimensional open cavities on plane and cylindrical surfaces in a supersonic stream in the presence of a turbulent boundary layer have been investigated experimentally. The effects of the Mach number, boundary layer thickness, the shape of the cavity, and its angle of inclination to the free-stream direction on the flow parameters in the mixing layer above the cavity and the heat flux and pressure distribution on the surface of the cavity and its bottom are descirbed. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 4, pp. 74–80, July–August, 1998.     

Numerical simulation of stabilization of the boundary layer on a surface with a porous coating in a supersonic separated flow

Stability of a supersonic (M _∞ = 5.373) boundary layer with local separation in a compression corner with a passive porous coating partly absorbing flow perturbations is considered by solving two-dimensional Navier-Stokes equations numerically. The second mode of disturbances of a supersonic boundary layer is demonstrated to be the most important one behind the boundary-layer reattachment point. The possibility of effective stabilization of these disturbances behind the reattachment point with the use of porous coatings is confirmed. __________ Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, Vol. 48, No. 2, pp. 39–47, March–April, 2007.     

高速三维边界层的横流不稳定性

用两点四阶差分格式研究旋转圆锥超音速三维边界层的横流不稳定性和壁面冷却对稳定性的影响数值结果表明，与二维边界层相比横流使三维边界层第一模式增长率增大，对第二模式影响很小；Ｍｅ＜４３第一模式最不稳定，Ｍｅ＞４３第二模式最不稳定；三维边界层最不稳定第二模式是三维波，二维边界层则为二维波；壁面冷却对第一模式起稳定作用，对第二模式起不稳定作用     

Origin of instability waves in a boundary layer with moderate free-stream turbulence

The characteristics of the transition mechanism in the boundary layer on a wing section are experimentally investigated for moderate free-stream turbulence. It is shown that the frequency and wavelength of the instability depend on the thickness of the boundary layer in the region of the point of inflection on the average velocity profile. Novosibirsk. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 1, pp. 49–54, January–February, 1994.     

Stability of the Nonstationary Swept Attachment-Line Boundary Layer with Time Variation of the Surface Temperature and Gas Suction Velocity

The possibility of controlling the stability of a nonstationary boundary layer on the attachment line of a high-aspect-ratio swept wing by means of periodic variations of the surface temperature or the gas suction velocity at sub- or supersonic free-stream velocities is considered. The characteristic time scale of the variations of the surface temperature or the gas suction velocity on the attachment line is assumed to be equal to the characteristic aerodynamic time. On this assumption the stability characteristics of quasisteady attachment-line boundary layer flows are studied, the minimum values of the critical Reynolds numbers Re^* of loss of stability are determined as functions of the temperature and the suction velocity, and examples of the periodic dependence of the surface temperature and the suction velocity for which, in the case of nonstationary flow, the time-average values of Re^* exceed the analogous values for the steady-state boundary layer are constructed.     

Swept-Wing Boundary Layer Receptivity to a Steady Free-Stream Vortex Disturbance

A single trailing vortex developed behind a micro-wing immersed in a free stream was used to study the vortex receptivity of a swept-wing boundary layer. As a result of the interaction, longitudinal-velocity disturbances develop in the boundary layer. On the swept wing, disturbance transformation occurs near the leading edge and is accompanied by the formation of a wave packet consisting of waves typical of cross-flow instability. Disturbances with other characteristics are also detected. These disturbances may be attributable to distributed boundary-layer receptivity to the free-stream vortex disturbance considered.     

Pressure oscillation on a plane upstream of an obstacle in a supersonic flow

The levels and spectra of pressure oscillation on a plane upstream of a vertical cylinder and a step in an M=3 supersonic flow are measured in the presence of a turbulent boundary layer. Moscow. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 1, pp. 69–74, January–February, 1998.     

Infrared thermography investigations in transitional supersonic boundary layers

The study of boundary-layer transition in supersonic flows is conducted employing infrared thermography (IRT). Several models of swept wings are tested in a blow-down facility at Mach number 2.4. The effects of wing sweep and other parameters (angle of attack, leading-edge contour, presence/absence of surface roughness) are successfully observed. The transition front is clearly identified, demonstrating the utility of IRT for this type of study. The technique is particularly indicated for flows that are sensitive to surface alterations (roughness), such as transitional boundary layers, because it does not require interaction with the model or the flow under investigation. The additional advantage of no need for special apparatus, except for the infrared camera, makes IRT well suited for both wind-tunnel and in-flight testing. Practical problems and limitations encountered when dealing with IRT in high-speed flows are also discussed.     

“Non-free” interaction of plane shock waves and the boundary layer in the neighborhood of the leading edge of a plate with slip

The interaction between a normally impinging shock wave and the boundary layer on a plate with slip is studied in the neighborhood of the leading edge using various experimental methods, including special laser technology, to visualize the supersonic conical gas flows. It is found that in the “non-free” interaction, when the leading edge impedes the propagation of the boundary layer separation line upstream, the structure of the disturbed flow is largely identical to that in the developed “free” interaction, but with higher parameter values and gradients in the leading part of the separation zone. The fundamental property of developed separation flows, namely, coincidence of the values of the pressure “plateau” in the separation zone and the pressure behind the oblique shock above the separation zone of the turbulent boundary layer, is conserved. Moscow. e-mail: ostap@inmech.msu.su. Translated from Izvestiya Rossiiskoi Akademii Nauk, Mekhanika Zhidkosti i Gaza, No. 3, pp. 57–69, May–June, 2000. The work was carried out with financial support from the Russian Foundation for Basic Research (project No. 97-01-00099).