Numerical investigation on the effects of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade

This paper presents a detailed numerical investigation of the influence of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade. The boundary layer suction (BLS) was used to control the location and strength of shock waves, with the aspirated slot locating at 49% chord, where is just downstream of the impingement point of shock wave at the leading edge. The numerical simulation is based on NUMECA, a commercial software, where the cell-centered control volume approach with third-order spatial accuracy is used to solve the 3-D Reynolds-averaged Navier-Stokes equations under the Cartesian coordinate system. Several conclusions can be made through the observation of the numerical results. (1) Multiple shock waves in cascade passage leaded the velocity deficits of boundary layer on suction surface downstream of shock wave, resulting in seriously separated flow on the suction side of blade, especially when the front shock wave is much stronger than the rest of the shocks. (2) BLS with small mass flow rate can not effectively improve the boundary layer. When the impingement point of oblique shock wave coming from cascade leading edge is bled to downstream of the passage shock wave by BLS, the boundary layer flow is greatly improved. However, if the BLS mass flow rate exceeds a critical value (1.2%), the boundary layer downstream of shock wave would separate from suction surface. (3) At the blade mid-span, the aerodynamic performance of compressor blade is improved as BLS mass flow rate increases. The optimum BLS is about 1.2%. Compared with the baseline case, the BLS with flow rate of 1.2% increases the total pressure recovery coefficient by 12%, and decreases diffusion factor by 18% and deviation angle to 7 ° while keeping the pressure rise constant. (4) The three dimensional flow structure of the compressor cascade ranged from 25% span to 75% span was improved greatly with the 1.2% BLS flow rate. However it could not control the development of the corner boundary layer effectively.     

Two-dimensional numerical simulations of shock waves in micro convergent–divergent nozzles

The steady two-dimensional Navier–Stokes equations with the slip wall boundary conditions were used to simulate the supersonic flow in micro convergent–divergent nozzles. It is observed that shock waves can take place inside or outside of the micronozzles under the earth environment. For the over-expanded flows, there is a boundary layer separation point, downstream of which a wave interface separates the viscous boundary layer with back air flow and the inviscid core flow. The oblique shock wave is followed by the bow shock and shock diamond. The viscous boundary layer thickness relative to the whole nozzle width on the exit plane is increased but attains the maximum value around of 0.5 and oscillates against this value with the continuous increasing of the nozzle upstream pressures. The viscous effect either changes the normal shock waves outside of the nozzle for the inviscid flow to the oblique shock waves inside the nozzle, or transfers the expansion jet flow without shock waves for the inviscid flow to the oblique shock waves outside of the nozzle.     

Transonic instability in entrance part of mixing chamber of high-speed ejector

Introduction The research of flow structure in the entrance part of the mixing chamber of two-dimensional supersonic ejector[1,2] shows, how this structure depends both on stagnation pressure ratio of streams p01/p02[3] and on back pressure ratio pb/p02 [4]. It was found out that the structure of shock waves is not stable, but it oscillates less or more. For the high back pressure ratio a terminal shock wave is in the mixing chamber and due to this shock wave the mixing processes change quali…     

Effect of shock waves on supersonic film cooling with a slotted wall

A slotted wall with a cavity which reduces the effect of the shock wave on the film cooling was developed through understanding of the mechanism by which the shock wave affects the supersonic film cooling. Numerical results show that the supersonic film cooling effectiveness with the slotted wall is improved after the shock wave incidence, even better than that without the shock wave effect. The cooling stream flows into the cavity upstream of the slotted wall and flows out downstream, which bypasses more cooling gas to protect the surface downstream after the shock wave incidence, which weakens the effect of the shock wave on the film cooling. Upstream of the shock wave incidence, the slotted wall reduces the mixing between the mainstream and the cooling stream and the coolant boundary layer thickness, which reduces the film cooling effectiveness for both structures than without the slotted wall, with an effectiveness which nearly the same as or even a little better than without the slotted wall for another structure.     

Effects of incident shock wave on mixing and flame holding of hydrogen in supersonic air flow

The effects of incident shock wave on mixing and flame holding of hydrogen in supersonic airflow have been studied numerically. The considered flow field was including of a sonic transverse hydrogen jet injected in a supersonic air stream. Under-expanded hydrogen jet was injected from a slot injector. Flow structure and fuel/air mixing mechanism were investigated numerically. Three-dimensional Navier–Stokes equations were solved along with SST k-ω turbulence model using OpenFOAM CFD toolbox. Impact of intersection point of incident shock and fuel jet on the flame stability was studied. According to the results, without oblique shock, mixing occurs at a low rate. When the intersection of incident shock and the lower surface is at upstream of the injection slot; no significant change occurs in the structure of the flow field at downstream. However when the intersection moves toward downstream of injection slot; dimensions of the recirculation zone and hydrogen-air mixing rate increase simultaneously. Consequently, an enhanced mixing zone occurs downstream of the injection slot which leads to flame-holding.     

An improvement on the efficiency of a single rotor transonic compressor by reducing the shock wave strength on the blade suction surfaces

It is well known that increasing the rotational velocity is an effective way to increase the total pressure ratio. With increasing flow velocity especially under the condition of transonic flow in the supersonic region, where exist strong shock waves, the shock wave loss becomes main and important. Simultaneously, there occurs boundary layer separation due to the shock wave / boundary layer interaction. In the present paper the transonic compressor blades were studied and analyzed to find a proper and simple way to reduce the shock wave loss by optimizing the suction surface configuration or controlling the gradient of isentropic Mach number on the suction surface. A Navier-Stokes solver combined with a modified design algorithm was developed and used. The NASA single rotor for transonic flow compressor was served as a numerical example to show the effectiveness of this method. Two cases for both original and modified rotors were analyzed and compared.     

The effect of chemical energy release on heat transfer from flames in small channels

Stable combustion in a heated tube, with a radius on the order of the flame thickness, is investigated experimentally and numerically. The downstream portion of the tube is heated by an external heat source resulting in a steady, axially varying temperature gradient along the tube wall. Strongly burning, axisymmetric methane/air flames are stabilized inside this wall temperature profile which are observed to be “flat” for sufficiently small tube dimensions. The position of these flames is dictated by a competition between the energy required to preheat the reactants, that released by combustion, and the heat lost to the wall. To model such flames, an extension to the standard 1-D, volumetric flame formulation is proposed to solve for wall/gas heat transfer by employing a thermal boundary layer. The boundary layer utilizes a non-linear, radially-varying heat source to account for combustion and captures the effect of enhanced interfacial heat transfer inside the reaction zone. The proposed numerical model gives improved quantitative predictions for flame stabilization position than approaches which neglect the effect of heat release by modeling heat transfer with Newton’s law of cooling and a local Nusselt number.     

Breakdown of the boundary-layer approximation for mixed convection above a horizontal plate

The mixed convection flow over a horizontal plate is investigated. Boundary-layer equations, modified to account for the hydrostatic pressure variation across the boundary layer, are solved numerically by a finite-difference scheme. If the (normal) fluid is cooled from below (or heated from above), the numerical results indicate that the derivative of the wall shear stress tends to infinity as a critical distance from the leading edge is approached. This occurs when the wall shear stress is till positive, i.e. before the classical separation criterion is satisfied. A mechanism of self-amplification is shown to be responsible for the breakdown.     

Passive Control of Transonic Flow Fields with Shock Wave Using Non—equilibrium Condensation and Porous wall

When non-equilibrium condensation occurs in a supersonic flow field, the flow is affected by the latent heat released. In the present study, in order to control the transonic flow field with shock wave, a condensing flow was produced by an expansion of moist air on a circular bump model and shock waves were occurred in the supersonic parts of the fields. Furthermore, the additional passive technique of shock / boundary layer interaction using the porous wall with a cavity underneath was adopted in this flow field. The effects of these methods on the shock wave characteristics were investigated numerically and experimentally. The result obtained showed that the total pressure loss in the flow fields might be effectively reduced by the suitable combination between non-equilibrium condensation and the position of porous wall.     

Transition to a periodic flow induced by a thin fin on the sidewall of a differentially heated cavity

The transition to a periodic flow induced by a thin fin on the sidewall of a differentially heated cavity is numerically investigated. The numerical results are compared with a previously reported experiment. It is demonstrated that the transient flow obtained numerically shows features consistent with the experimental flow. Based on the present numerical results, the temporal development and spatial structures of the thermal flow around the fin are described, and the separation of the thermal flow above the fin is discussed. It is found that the presence of the fin changes the flow regime and results in the transition of the thermal flow to a periodic flow. The present numerical results also indicate that the unstable temperature configuration above the fin results in intermittent plumes at the leeward side of the fin, which in turn induce strong oscillations of the downstream boundary layer flow. It is demonstrated that the oscillations of the boundary layer flow significantly enhance the heat transfer through the finned sidewall (by up to 23%).     

Turbulence characteristics and vortical structures in combined-convection boundary layers along a heated vertical flat plate

Time-developing direct numerical simulations are performed for the combined-convection boundary layers created by imposing aiding and opposing freestreams to the pure natural-convection boundary layer in air along a heated vertical flat plate to clarify their structural characteristics. The numerical results reveal that with a slight increase in freestream velocity, the transition region moves downstream for aiding flow and upstream for opposing flow. This fact correlates well with the existing experimental results showing that the transition delays for aiding flow and quickens for opposing flow in the practical space-developing boundary layer. Thereby, for aiding flow, turbulence characteristics indicate the behavior proceeding to the laminarization of the boundary layer. On the other hand, for opposing flow, the large scale fluid motions are apparent and become larger than those for the pure natural-convection boundary layer with increasing freestream velocity. For the occurrence of such fluid motions, the budgets of turbulent energy and two-point spatial correlations in the turbulent combined-convection boundary layers are also examined. Consequently, it is found in the spatial correlations that the turbulence structures are mainly controlled by fluid motions in the outer region of the boundary layer.     

The initiation and propagation of detonation in supersonic combustible flow with boundary layer

Adaptive simulations solving the Navier-Stokes equations have been conducted in order to get a better understanding on the detonation initiation and propagation in a stoichiometric H₂/O₂/Ar supersonic mixture with boundary layer. The detonation is initiated by a continuous hot jet. When reflecting on the wall, the jet induced bow shock interacts with the boundary layer and forms the shock boundary layer interaction phenomena, while in Euler result the bow shock forms Mach reflection. The investigation shows that the Navier-Stokes simulation result is structurally in better agreement with the experiment compared with that of the inviscid Euler simulation result. The bow shock interacts with the separation shock, forming the shock induced combustion behind the interaction zone. Then the combustion front couples with shock and forms Mach stem induced detonation. The Mach stem induced detonation continues to getting higher and propagating upstream, initiating the main flow. The initiated partial detonation exists with the separation shock induced combustion front, forming an “oblique shock induced combustion-partial detonation” structure in the main flow. The investigation on the influence of free stream Mach number further confirms that the boundary layer has an important influence on detonation initiation. The parametric studies also show that there exists a free stream Mach number range to initiate the partial detonation in supersonic combustible flow successfully.     

The effect of air humidity on shock wave induced incipient separation

InttoductionTtansonic flows often aPPear in modem tecbocalaPPlications. The most boleal examples are in civil andmilitary aeroplanes, but it also concerns the advancingblades of a helicopter rotor, supersonic jab engine intakesas well as compressor and tUlbine cascades.The crucial phenomenon Which is inVolved in suchnows is the formation of shock waves and theirinteraction with bounce layers adjacent to boundingwalls. Shock waves Which introduce the rapid increase ofparameters impose an ad…     

跨音速透平叶栅尾缘劈缝射流气动性能分析

基于定常RANS方程,采用Spalart-Allmaras(S-A)湍流模型,数值模拟某跨音速导叶尾缘劈缝射流的定常流动结构,分析尾缘劈缝射流对尾缘激波结构、尾迹流动特性及叶栅气动性能的影响。研究表明:开缝射流显著降低尾缘压力面侧燕尾波强度,并使激波在相邻叶片吸力面入射点向上游移动;当叶栅出口马赫数小于1.35时射流使吸力面燕尾波强度减弱,而达到1.35后射流使该侧激波强度增大;在不同出口马赫数下射流均能降低叶栅动能损失。     

Shock wave boundary layer interactions in hypersonic flows over a double wedge geometry by using conjugate heat transfer

Shock wave boundary layer interaction phenomena play a critical role in the design of supersonic and hypersonic vehicles. Consequently, this paper mainly focuses on hypersonic flow over a double wedge model, flow fields around concave corners are relatively complicated, and produce several classical viscous flow features depending on the combination of the first and second wedge, and the important characteristic phenomena are mainly the shock‐boundary layer and shock‐shock interaction. For these interactions, aerodynamic heating and pressure loads increase greatly when interactions are present. The conjugate heat transfer (CHT) technique is expected to exactly predict the separation bubble length, heat flux, skin friction coefficient, and pressure distributions in double wedge studies in hypersonic applications. In the present CHT studies, the different wall materials used are thermal insulation, Macor, and SiC, it is clearly shown that while using Macor and thermal insulctation wall material in CHT studies, the interface temperature, skin friction coefficient, heat flux distribution along the length change significantly with increase in simulation time. In comparing the CHT results with the fluid flow solver with the wall, considering isothermal and adiabatic boundary results, it is clearly indicated that the fluid flow solver results are either underpredicting or overpredicting the interface properties, but CHT studies give an accurate prediction of the separation length and interface properties.     

Laminar-turbulent transition in supersonic boundary layers with surface heat transfer: A numerical study

Through high-resolution direct numerical simulations, the present study aims to investigate several laminar-to-turbulent transition scenarios in the presence of wall heat transfer for supersonic boundary layers over strongly heated/cooled and adiabatic flat plates. The laminar boundary layer is tripped using a suction and blowing technique with a single-frequency, multiple-spanwise wavenumber excitation. The results are evaluated and compared with linear stability theory to isolate the effect of wall heat transfer, as well as forcing parameters, on the transition. It was found that increasing the disturbance amplitude as well as perturbation frequency moves the transition upstream. Also, the effect of wall heating was seen to stabilize the flow and to postpone the transition, contrary to the wall cooling.     

Effects of boundary layer on wedge-induced oblique detonation structures in hydrogen-air mixtures

Oblique detonation wave (ODW) structures are studied widely in recent years, but most of them are solved by the Euler equations without considering viscosity and then effects of boundary layer. In this study, the Navier-Stokes Equations are used to simulate the wedge-induced ODWs in hydrogen-air mixtures, and the two types of ODW transition structures at different incident Mach number M_i are analyzed to clarify the effects of viscosity and hence the boundary layer. Results show that the effect of boundary layer on ODW structures should be classified by the types of ODW transition patterns. As for the smooth transition pattern of ODW at high Mach numbers, the effect of boundary layer can be neglected, but for the abrupt transition pattern of ODW at low Mach numbers, the effect of boundary layer is large and it changes the ODW structure greatly. Resulting from the interaction of shock and boundary layer, a recirculation zone is formed within the viscous ODW layer at M_i = 7, which leads to the phenomenon that the straight oblique shock wave evolves into two sections, with the downstream one having a larger shock angle. Additionally, the corresponding transition position moves upstream, and the initiation length becomes only one third of that in inviscid ODW. The great importance of considering viscosity in ODW simulations and future designs of combustor of oblique detonation engine has been addressed.     

Study on Shock Wave and Turbulent Boundary Layer Interactions in a Square Duct at Mach 2 and 4

In this paper, the outline of the Mach 4 supersonic wind tunnel for the investigation of the supersonic internal flows in ducts was firstly described. Secondly, the location, structure and characteristics of the Mach 2 and Mach 4 pseudo-shock waves in a square duct were investigated by color schlieren photographs and duct wall pressure fluctuation measurements. Finally, the wall shear stress distributions on the side, top and bottom walls of the square duct with the Mach 4 pseudo-shock wave were investigated qualitatively by the shear stress-sensitive liquid crystal visualization method. The side wall boundary layer separation region under the first shock is narrow near the top wall, while the side wall boundary layer separation region under the first shock is very wide near the bottom wall.     

Natural transition in natural convection boundary layers

The ‘natural transition’ of a natural convection boundary layer adjacent to an isothermally heated vertical surface is investigated by means of three-dimensional direct numerical simulation (DNS). In order to trigger ‘natural transition’ numerically, spatially and temporally random perturbations are introduced into the upstream boundary layer. The propagation of the random perturbations in the streamwise direction is observed. It is found that there exist two competing wavenumbers of spanwise vortical structures, one large and the other small. The large wavenumber dominates in the upstream boundary layer, whereas the small wavenumber dominates in the downstream boundary layer. The streamwise evolution of the mean (time-averaged) streamwise vorticity observed at planes perpendicular to the heated surface in general reveals two- and three-layer longitudinal roll structures. Nonlinear processes in a transitioning natural convection boundary layer are also analysed using Bicoherence method. The transition route and mechanism are discussed based on the power spectra and Bicoherence spectra of the temperature time series obtained in the boundary layer. A spectrum filling process during the ‘natural transition’ to turbulence is also observed, which is qualitatively similar to that observed in Blasius boundary layers.     

Three Dimensional Separation with Spiral—Focus in a Decelerating Duct Flow （Effect of Asymmetric Inlet Boundary Layer Thickness）

An experimental apparatus was developed to study the three dimensional separated flow with spiral-foci. The internal decelerating flow was generated by the air suction from a side wall to produce the separation on an opposite-side wall. The relation between the upstream boundary layer and the generation of spiral-foci in the separation region was observed by a tuft method. As a result, it was clarified that the spiral-focus type separation could be produced on the side wall and its behavior was closely related to the vortices supplied into the separation region from the boundary layer developing along top wall or bottom one.