Development of magnetohydrodynamic boundary layers

Many authors have studied the problem of the development of a hydrodynamic boundary layer when a body is suddenly set in motion. The results obtained are presented most fully in the monographs of H. Schlichting [1] and L. G. Loitsyanskii [2]. In magnetohydrodynamics the development of the boundary layer over the surface of an infinite flat plate for uniform oncoming flow has been closely studied (for example [3, 4]). Below, the problem of the development of a plane magnetohydrodynamic boundary layer is considered in a different formulation. We shall suppose that the distributions of velocity U(x) and enthalpy h(x) are given along the body contour for t=0. At that moment the viscosity and thermal conductivity mechanisms are instantaneously switched on. Viscous and thermal boundary layers begin to grow in a direction normal to the body. The medium in the boundary layer interacts with the magnetic field. This formulation corresponds to the development of a magnetohydrodynamic boundary layer on a body which is set in motion with a jerk, in the case where the rate of establishment of magnetohydrodynamic flow of the inviscid, thermally nonconducting fluid around the body is much less than the rate of development of the boundary layer. Then U(x) and h(x) are found by solving the problem of stationary magnetohydrodynamic flow of an inviscid thermally nonconducting fluid around a body, or simply the hydrodynamic flow if the medium interacts with the field only in the boundary layer.     

Calculation of the deceleration of a viscous supersonic gas flow in a channel

The deceleration of nonuniform viscous supersonic gas flows in planar and axisymmetric channels is investigated. A modification of Prandtl's formula for the turbulent viscosity is proposed in order to take into account the dependence of the mixing length on the value of the axial Mach number. The results of the calculations are compared with known experimental data on the deceleration of a supersonic flow in a subsonic pseudoshock.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 2, pp. 162–166, March–April, 1982.We thank A. N. Sekundov for discussing the work.     

Visualization of the structure of supersonic turbulent boundary layers

A series of flow visualizations has been performed on two flat-plate zero-pressure-gradient supersonic boundary layers. The two different boundary layers had moderate Mach numbers of 2.8 and 2.5 and Re 's of 82, 000 and 25, 000 respectively. A number of new visualization techniques were applied. One was a variation of conventional schlieren employing selective cut-off at the knife edge plane. Motion pictures of the flow were generated with this technique. Droplet seeding was also used to mark the flow, and high speed movies were made to show structure evolution. Still pictures were also taken to show details within the large-scale motions. Finally, Rayleigh scattering was used to construct planar images of the flow. Together, these techniques provide detailed information regarding the character and kinematics of the large-scale motions appearing in boundary layers in supersonic flow. Using these data, in concert with existing hot-wire data, some suggestions are made regarding the characteristics of the average large-scale motion.This work was supported by the Air Force Office of Scientific Research under Grant 89-0120, monitored by Dr. James M. McMichael. Also, the authors wish to thank Prof. R. B. Miles for his contributions to the Rayleigh scattering portion of this project.     

The influence of boundary layers on supersonic inlet flow unstart induced by mass injection

A transverse jet is injected into a supersonic model inlet flow to induce unstart. Planar laser Rayleigh scattering from condensed CO₂ particles is used to visualize flow dynamics during the unstart process, while in some cases, wall pressure traces are simultaneously recorded. Studies conducted over a range of inlet configurations reveal that the presence of turbulent wall boundary layers strongly affect the unstart dynamics. It is found that relatively thick turbulent boundary layers in asymmetric wall boundary layer conditions prompt the formation of unstart shocks; in symmetric boundary conditions lead to the propagation of pseudo-shocks; and in both cases facilitate fast inlet unstart, when compared with thin, laminar boundary layers. Incident shockwaves and associated reflections are found to affect the speed of pressure disturbances. These disturbances, which induce boundary layer separation, are found to precede the formation of unstart shocks. The results confirm the importance of and need to better understand shock-boundary layer interactions in inlet unstart dynamics.     

Stationary supersonic conducting gas flow in a channel with nonconducting walls under weak magnetohydrodynamic interaction

A solution of the problem of flow in a channel with nonconducting walls for a small magnetohydrodynamic interaction parameter N is obtained by numerical methods. In the 0–10 range of variation of the Hall and magnetic Reynolds number parameters the distributions of the electrical parameters and the average (over the cross section) and local gasdynamic flow parameters are computed for two different geometries of the applied magnetic field. It is shown that an increase in the Hall and magnetic Reynolds number parameters is accompanied by a diminution in the Joule dissipation and the perturbation of the average (over the cross section) gasdynamic flow characteristics. It is disclosed that the distribution of the gasdynamic parameters over the channel cross section is extremely nonmonotonic in the end current zones.Translated from Zhurnal Prikladnoi Mekhaniki i Tekhnicheskoi Fiziki, No. 4, pp. 20–29, July–August, 1970.In conclusion, the author is grateful to A. B. Vatazhin for useful comments and constant attention to the research and to I. U. Tolmach for valuable comments.     

On the use of a two-layer model for large-eddy simulations of supersonic boundary layers with separation

The two-layer modeling approach has become one of the most promising and successful methodology for simulating turbulent boundary layers in the past ten years. In the present study, a mixed wall model for large-eddy simulations (LES) of high-speed flows is proposed which combine two approaches; the thin-Boundary Layer Equations (TBLE) model of Kawai and Larsson (1994) and the analytical wall-layer model of Duprat et al. (2011) for streamwise pressure gradients. The new hybrid model has been efficiently implemented into a three-dimensional compressible LES solver and validated against DNS of a spatially-evolving supersonic boundary layer (BL) under moderate and strong pressure gradients, before being employed for the prediction of nozzle flow separations at different flow conditions, ranging from weakly to highly over-expanded regimes. A good agreement is obtained in terms of mean and fluctuating quantities compared to the DNS results. Particularly, the current wall-modeled LES results are found to perfectly match the DNS data of supersonic BL with/out pressure gradient. It is also shown that the model can account for the effect of the large-scale turbulent motions of the outer layer, indicating a good interaction between the inner and the outer part of the wall layer. In terms of simulations costs and improvements of computing power, the obtained results highlight the capability of the current wall-modeling LES strategy in saving a considerable amount of computational time compared to the wall-resolved LES counterpart, allowing to push further the simulations limits. Furthermore, the application of these computationally low-costly LES simulations to nozzle flow separation allows to clearly identify the origin of the shock unsteadiness, and the existence of broadband and energetically-significant low-frequency oscillations (LFO) in the vicinity of the separation region.     

Cone in a supersonic flow near a surface with a turbulent boundary layer

The published information about the interaction of incident shocks and a turbulent boundary layer relate to cases of a thin boundary layer ( 1–3 mm) on a flat surface. The present study relates to supersonic flow with Mach number M = 3 and stagnation pressure p₀=1.2 MPa past cones near a surface with a thick boundary layer formed on a plate abutting the lower edge of a plane nozzle.Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 4, pp. 177–180, July–August, 1991.     

Investigation of the heat transfer mechanism in supersonic turbulent boundary layers

A method is described for calculating turbulent Prandtl numbers from Mach number and total temperature profiles in supersonic boundary layers. The calculations are based on boundary layer measurements in the Mach number range from 3.5 to 5. The investigations clearly indicate that in addition to accurate profile measurements reliable values of shear stress and heat flux at the wall must exist, in order to be able to calculate the turbulent Prandtl number in the viscous regime of the boundary layer. For flow conditions with and without heat transfer, the derived turbulent Prandtl numbers indicate that the turbulent transport of heat decreases much faster towards the wall than the turbulent transport of momentum. The results of the analysis show that only the unequivocal qualitative result of increasing turbulent Prandtl numbers in the viscous region of the boundary layer, can be expected. The variation of the turbulent Prandtl number can be described successfully using a simple approximation, based on the mixing length concept, and is applied to the calculation of total temperature distribution using the law of the wall for compressible flow.     

MHD deceleration and heat transfer for a sphere in a supersonic flow of partially ionized gas

Translated from Zhurnal Prikladnoi Mekhaniki i Tekhnicheskoi Fiziki, No. 1, pp. 15–19, January–February, 1991.     

Laminar boundary layers at a moving interface generated by counter-current gas-liquid stratified flow

An experimental and theoretical investigation of the velocity distribution in the laminar boundary layers at a moving interface generated by a counter-current gas-liquid stratified flow has been made. No similarity solution of the problem exists, as in the case of co-current motion of the phases. The solution is carried out by means of the perturbation method. A comparison of the theoretical results with experimental ones obtained by the use of hot-wire technique is performed.     

The rapid expansion of a turbulent boundary layer in a supersonic flow

Rapid Distortion Approximations (RDA) may be used to simplify the Reynolds stress equations in rapidly distorted flows, as suggested by Dussauge and Gaviglio (1987). These approximations neglect diffusive and dissipative terms while retaining the production and pressure terms. The retained terms are then modeled as functions of the Reynolds stress tensor and gradients of the mean flow. The models for the pressure-strain term as developed by Lumley (1978) and Shih and Lumley (1985) are evaluated by comparing the calculated results with experimental data for the case of a Mach 2.84 turbulent boundary layer in a 20° centered expansion. The agreement between computed and experimentally obtained Reynolds stresses was found to be encouraging.Dedicated to Professor J.L. Lumley on the occasion of his 60th birthday.This work was supported by the U.S. Air Force under AFOSR Contract 89-0420. Monitored by Dr. James McMichael.     

Analysis of turbulent boundary layer interaction with an external supersonic flow behind a step

An integral method of analyzing turbulent flow behind plane and axisymmetric steps is proposed, which will permit calculation of the pressure distribution, the displacement thickness, the momentum-loss thickness, and the friction in the zone of boundary layer interaction with an external ideal flow. The characteristics of an incompressible turbulent equilibrium boundary layer are used to analyze the flow behind the step, and the parameters of the compressible boundary layer flow are connected with the parameters of the incompressible boundary layer flow by using the Cowles-Crocco transformation.A large number of theoretical and experimental papers devoted to this topic can be mentioned. Let us consider just two [1, 2], which are similar to the method proposed herein, wherein the parameter distribution of the flow of a plane nearby turbulent wake is analyzed. The flow behind the body in these papers is separated into a zone of isobaric flow and a zone of boundary layer interaction with an external ideal flow. The jet boundary layer in the interaction zone is analyzed by the method of integral relations.The flow behind plane and axisymmetric steps is analyzed on the basis of a scheme of boundary layer interaction with an external ideal supersonic stream. The results of the analysis by the method proposed are compared with known experimental data.Notation x, y longitudinal and transverse coordinates - X, Y transformed longitudinal and transverse coordinates - , ^*, ^** boundary layer thickness, displacement thickness, momentum-loss thickness of a boundary layer - , ^*, ^** layer thickness, displacement thickness, momentum-loss thickness of an incompressible boundary layer - u, velocity and density of a compressible boundary layer - U, velocity and density of the incompressible boundary layer - , stream function of the compressible and incompressible boundary layers - , dynamic coefficient of viscosity of the compressible and incompressible boundary layers - r₁ radius of the base part of an axisymmetric body - r radius - R transformed radius - M Mach number - friction stress - p pressure - a speed of sound - s enthalpy - v Prandtl-Mayer angle - P Prandtl number - P_t turbulent Prandtl number - r₂ radius of the base sting - b step depth - =0 for plane flow - =1 for axisymmetric flow Translated from Izvestiya Akademii Nauk SSSR, Mekhanika Zhidkosti i Gaza, No. 3, pp. 33–40, May–June, 1971.In conclusion, the authors are grateful to M. Ya. Yudelovich and E. N. Bondarev for useful comments and discussions.     

Formation of a shock wave at the boundary of a local supersonic zone

Translated from Prikladnaya Mekhanika i Tekhnicheskaya Fizika, No.1, pp. 24–32, January–February, 1993.