Inlet Boundary Layer Research Articles

Large-Eddy Simulations of Flow over a Backward-Facing Ramp with a Wall-Mounted Cube

Passive flow control devices, such as vortex generators (VGs), can effectively modulate the turbulent boundary layer flow near regions of adverse pressure gradients, but the interactions between the salient flow structures produced by VGs and those of the separated flow are not fully understood. In this study, a spatially evolving turbulent boundary layer interacting with a wall-mounted cube ahead of a backward-facing ramp is investigated using wall-resolved large-eddy simulations for a Reynolds number of 19,600, based on the inlet boundary layer thickness and freestream velocity. Different cube configurations are examined to isolate the effects of cube height and streamwise position. The large counter-rotating flow produced by the larger cubes generated more turbulent kinetic energy, thereby resulting in smaller separated regions over the ramp. Although significantly lower levels of turbulent kinetic energy dissipation than production are observed in the outer flow regions of the horseshoe vortex and in the induced counter-rotating flow, the spatial distribution of dissipation is similar to that of the turbulent kinetic energy. When the upstream position of an isolated single cube, relative to the leading ramp edge, is more than three cube heights, the streamwise decay of the counter-rotating flow results in lower levels of turbulent kinetic energy.

Design and Characterization of Highly Diffusive Turbine Vanes Suitable for Transonic Rotating Detonation Combustors

In rotating detonation engines the turbine inlet conditions may be transonic with unprecedented unsteady fluctuations. To ensure an acceptable engine performance, the turbine passages must be suited to these conditions. This article focuses on designing and characterizing highly diffusive turbine vanes to operate at any inlet Mach number up to Mach 1. First, the effect of pressure loss on the starting limit is presented. Afterward, a multi-objective optimization with steady RANS simulations, including the endwall and 3D vane design is performed. Compared to previous research, significant reductions in pressure loss and stator-induced rotor forcing are obtained, with an extended operating range and preserving high flow turning. Finally, the influence of the inlet boundary layer thickness on the vane performance is evaluated, inducing remarkable increases in pressure loss and downstream pressure distortion. Employing an optimization with a thicker inlet boundary layer, specific endwall design recommendations are found, providing a notable improvement in both objective functions.

Prediction of the Influence of the Inlet End-Wall Boundary Layer on the Secondary Flow of a Low Pressure Turbine Airfoil Using RANS and Large-Eddy Simulations

Abstract This paper analyzes the effect of the inlet end-wall boundary layer on the secondary flow of a low pressure turbine airfoil cascade at Reynolds number 2 × 105 using RANS and implicit large-eddy simulations (LES). The results are compared against experimental data obtained at two low-speed linear cascade facilities, one located at the Whittle Laboratory of the University of Cambridge and the other at the Polytechnic University of Madrid. The RANS turbulence model is the k−ω−γ−Reθt and no sub-grid scale model has been used in the LES. An unstructured mesh of hexahedra and prisms is used, with high order elements used in the boundary layer region to better describe the airfoil shape in the LES. Two inlet end-wall boundary layers that produce different secondary flow patterns are analyzed: a laminar thin velocity profile and a turbulent thick velocity profile with several inlet turbulent intensities. The agreement between LES numerical predictions and experimental measurements of the position and intensity of the secondary vortices is very good for both cases. RANS simulations are much cheaper in terms of computational cost and reasonably predict most of the flow features, except when the inlet turbulence is low and turbulent transition prediction becomes critical. The effect of the inlet velocity profiles and inlet turbulence on the secondary flow structure is quite pronounced. The velocity profile thickness determines the spanwise penetration of the passage vortex, and different inlet turbulence intensities modify its mixing. Higher inlet turbulence intensities lead to a decrease of the secondary losses due to the passage vortex and an increase of end-wall losses.

Roughness-Induced Transition in Supersonic Boundary Layers with Wall Heating and Cooling Effects

The laminar–turbulent transition of a supersonic boundary layer induced by an isolated roughness element is investigated using direct numerical simulation, BiGlobal linear stability theory, and the three-dimensional parabolized stability equation. Cylindrical and diamond-shaped roughness elements are investigated in combination with different wall temperatures. Direct numerical simulations show that the cylinder configuration induces an earlier transition than the diamond configuration, with the interaction between the separated shear layer and the counter-rotating vortices causing the transition. BiGlobal analysis and the parabolized stability equation confirm the existence of two unstable instability modes in the wake region, namely, a symmetric mode and an antisymmetric mode, with the former being strongly dominant. The wall cooling and heating effects are studied by changing the wall temperature. Wall heating lifts the inlet boundary layer and weakens the separated shear layer. This, in turn, weakens the wake instability and delays the transition. The antisymmetric model disappears as the wall heating increases. Wall cooling accelerates the transition by enhancing the distortion of the roughness wake, resulting in stronger instabilities. Finally, the Reynolds number based on the momentum defect is used to define the transition criterion, and this is found to be in good agreement with the present simulation results.

A synthetic particle group method for the unsteady interaction between shock wave and boundary layer past a compression ramp

To overcome the lack of resolved turbulent flow structures generated by the improved delayed detached eddy simulation (IDDES) model when simulating the shock wave and boundary layer interaction (SWBLI), a new, nonzonal, embedded synthetic turbulence generation method, named the synthetic particle group method (SPOM), is proposed and validated by simulating the supersonic flow past a flat plate at Mach 2.25. In the attached boundary layer, the SPOs are automatically introduced to trigger the turbulent perturbations. The turbulent boundary layer is successfully resolved and fully developed in a short streamwise range to produce the desired outcome. The effect of grid scale is also studied, and a suitable grid-scale value is recommended. Then, the SPOM is applied to the compression ramp flow at Mach 2.25 with unsteady SWBLI. The results show that the IDDES-SPOM is highly effective, accurate, and able to model scenarios where the original IDDES simulation fails. Furthermore, the effects of the inlet boundary layer thickness on local separation unsteadiness are studied. From the results, the lower frequency motion is attributed to a more fully developed boundary layer.

Nonequilibrium Heating/Cooling Effects in a Mach Number 10 Hydrogen-Fueled Scramjet

The influence of thermal/chemical nonequilibrium on a hydrogen-fueled Ma 10 scramjet was analyzed by combining the JF-2 4 s hock tunnel test and improved delayed detached eddy simulation modeling. A remarkable change in flame stabilization mode when incorporating the two-temperature nonequilibrium model was observed. The nonequilibrium heating and cooling effects were analyzed for different sections of the scramjet. The nonequilibrium heating effect facilitates the upstream flame propagation by inducing an early ignition and thickening the inlet boundary layer. The nonequilibrium heating or cooling effect is generally weak in the nearly constant-area isolator and combustor, where the flow is mainly influenced by the chemical nonequilibrium with a variation of 10% in reaction rate. The nonequilibrium cooling effect mainly exists in the expanding nozzle, where Tv>Tt, but the energy replenishment from vibrational mode to trans-rotational mode is nearly frozen when Tt<1000 K. Under both conditions, the final mixing is nearly complete, and net thrust has been achieved. When considering the nonequilibrium effects, the final combustion efficiency increases from 87.68 to 99%, together with a 53.94% rise in peak pressure ratio and a 137.04% rise in specific impulse.

Flow over a single dimple recessed in a flat plate

Direct numerical simulations have been conducted to investigate a zero-pressure-gradient boundary layer flow over a single shallow dimple. Here, the dimple depth to dimple diameter ratio (d/D) as well as the Reynolds number (based on D and free-stream velocity) are fixed at 0.05 and 20 000, respectively. The effect of inlet boundary layer thickness δ on a given dimple is investigated by considering δ/D∈[0.023,0.1]. The flow within the dimple exhibits either a horseshoe vortex (a continuous core line through the two spirals within the dimple) or a tornado-like vortex pair (discontinuous core line). For the given parameter range, four different flow patterns have been identified within the single dimple: (i) a steady symmetric horseshoe vortex pattern for δ/D∈[0.053,0.1], (ii) a steady asymmetric horseshoe vortex pattern for δ/D=0.04, (iii) a quasi-periodic asymmetric horseshoe vortex pattern for δ/D=0.033, and (iv) a mixed horseshoe and tornado-like vortex pattern for δ/D=0.023. The growth of the streamwise vorticity, mainly caused by the tilting of the vertical vorticity, plays a key role in the transition between the different flow patterns. Dimple-induced velocity streaks above the single dimple have been investigated in detail for the first time, showing four different streaks: (i) a high-speed streak above the dimple, (ii) two side-low-speed streaks located outside the dimple span, (iii) two side-high-speed streaks, and (iv) a mid-low-speed streak in between them. These are mainly caused by a flow acceleration effect and a flow diffuser effect over the dimple, as well as a “lift-up” mechanism within the downstream part of the dimple, tilting the boundary layer upward.

Validation of numerical simulations for investigating flow control in hypersonic inlets

Various flow control devices have been conceived for hypersonic inlets to suppress the inlet unstart and reduce the total pressure loss over the wide range of operating conditions of hypersonic systems. In particular, bleed systems that remove the low-momentum portion of the flow have drawn attention because they effectively reduce shock-induced flow separation and suppress inlet unstart. Numerical simulations are crucial for investigating the effectiveness of different bleed configurations in hypersonic inlets during the design phase. However, the flow structure in these inlets and the effect of the bleed on the flow are challenging to predict due to complicated factors such as shock-wave/boundary-layer interactions and compressibility effects. Moreover, since the boundary conditions, turbulence model, and turbulence corrections affect the numerical results, the most appropriate model can be difficult to determine. Therefore, a numerical solver was developed and evaluated on a series of validation cases with available experimental data to establish the fidelity of the numerical simulations. Seven validation cases, ranging from an elementary configuration involving a flat plate to a complex configuration involving a flat plate with a porous bleed and plenum chamber, were investigated. The flat plate flow numerical results, including the turbulent boundary layer and shock-wave/laminar boundary-layer interactions, were consistent with previous experimental and numerical results. However, in the hypersonic inlet cases, the actual physical phenomena could not be represented without considering compressibility and streamline curvature effects, which depend on the hypersonic inlet configuration. Thus, the effects of turbulence corrections on the numerical results of various hypersonic inlet configurations were examined. For cases with rapid expansion near the inlet throat or inlets with curved walls, compressibility or rotation/curvature corrections were necessary to accurately capture the size of the recirculation zone and reproduce the wall pressure data and shock structure along the flow path. However, for bleed flows on flat plates, the simulations were accurate without applying turbulence corrections. Thus, these turbulence corrections should be applied only when rapid expansion occurs or the wall near the shock-induced separation zone is curved. This paper provides the best practice appropriate for various supersonic, hypersonic flow cases, and can be utilized in future numerical studies on hypersonic inlet boundary layer control.

Influences of the cavity leakage flow on shrouded stator performance at different inlet boundary layer shapes

By using the validated numerical method to simulate the annular shrouded stator with the inter-stage seal cavity, the stator aerodynamic performance related to the cavity leakage flow was examined. The current study also indicated the interactions between the leakage flow and the mainstream. Discussions on the developments of secondary flow movement and hub corner separation were conducted, and the performances at different incoming boundary layer shapes were assessed using both the total pressure loss coefficient and the entropy-based loss coefficient. The results indicate that the cavity leakage flow creates a new vortex close to the leading edge of blade and is crucial to the passage vortex development and the concentrated shedding vortex size. At the same time, the cavity leakage flow weakens the transverse deflection of flow near the end wall and strengthens the three-dimensional flow effect. The two loss coefficients for the shrouded stator with seal cavity change little by thickening the boundary layer. The proper boundary layer skew as well as the interacted cavity leakage flow have stronger resistance to the deflection of the passage vortex and the transverse pressure gradient.

Analysis of the Effect of the Leading-Edge Vortex Structure on Unsteady Secondary Flow at the Endwall of a High-Lift Low-Pressure Turbine

The horseshoe vortex system at the leading edge of a low-pressure turbine (LPT) has unsteady flow characteristics, and the flow field in the downstream cascade channel also has unsteady characteristics. In this study, CFD simulations are performed with the help of commercial software CFX, and experimental checks are performed using a fan-shaped cascade test bench to investigate the flow characteristics of the endwall of the PACKB blade type of a high-lift LPT under the same incoming Reynolds number and two incoming boundary layer thickness conditions. The obtained results show that there are different flow alteration characteristics and alteration frequencies of the horseshoe vortex system, among which the vortex system with a thicker boundary layer is larger in size and less spaced from each other, which is more likely to induce the fusion of vortex systems. The centrifugal instability causes the instability of the horseshoe vortex system, and the instability frequency is inversely proportional to the thickness of the boundary layer. With two inlet boundary layers, the instability frequencies of the vortex system are 125 Hz and 175 Hz, respectively, and the ratio of the frequency to the thickness of the boundary layer is reciprocal to each other. The stimulation effect of the unstable horseshoe vortex system on the downstream secondary flow intensity is greater than that of the steady state. The thin boundary layer case generates a greater unsteady loss in the cascade channel than the thick boundary layer case due to the poor stability of the vortex system.

Research on the Influence of Inflow Conditions on the Aerodynamic Performance of a Tandem Cascade

We perform a thorough numerical analysis of the impact of inflow conditions on the aerodynamic performance of a tandem cascade. In particular, we investigate the effects of the incidence angle and the inlet boundary layer (IBL) thickness on the three-dimensional flow field structure and aerodynamic performance. Our results show that the gap flow strength of the tandem cascade decreases with the increase of incidence angle, and it can effectively reduce the mixing of the wakes of the forward blade (FB) and rear blade (RB). In turn, this prevents the passage vortex (PV) in the RB passage from developing along the circumferential direction. The occurrence of IBL does not modify the effects of the incidence angle on the tandem cascade, however, it reduces the load of the RB and the gap flow strength near the endwall. Under all incidence angles, IBL increases the total pressure loss of the tandem cascade, and decreases the static pressure rise (except for an incidence angle equal to -6°). The maximum loss increment is at 2° incidence angle, and the maximum static pressure rise decrement is at 6° incidence angle (Thick-IBL condition) or 7° incidence angle (Thin-IBL condition). Furthermore, we found that the presence of IBL changes the minimum loss condition from 0° (design condition) to -2° incidence angle. Our results thus indicate that in the practical engineering application of the tandem cascade, the reality that IBL degrades the tandem cascade performance in the full incidence angle range should be considered. And the strong endwall secondary flow effect caused by IBL should be considered in the tandem cascade three-dimensional design, so that the tandem cascade two-dimensional performance advantage can be better played.

Interaction mechanism between the tip leakage flow and inlet boundary layer in a highly loaded compressor cascade based on scale-adaptive simulation

Based on the scale-adaptive simulation, the interaction mechanism between the tip leakage flow (TLF) and the inlet boundary layer (IBL) and its effects on the tip flow field and aerodynamic performance of the compressor cascade were investigated. The time-averaged results show that the IBL reduces the blade tip load near the leading edge region, decreases the axial momentum of the TLF, and inhibits the development of the TLF to a certain extent. On the other hand, the IBL promotes the coupling of the tip leakage vortex, secondary vortex, and separation vortex and induces breakdown, resulting in a large area of severe flow separation in the corner region, greatly reducing the diffusion capacity and significantly increasing the flow loss in the tip region. The analysis of the unsteady transient flow fields indicates that the unsteady fluctuation in the tip region is mainly caused by the tip leakage vortex and flow separation near the blade trailing edge. The former is suppressed under the influence of the IBL, while the latter is amplified. The high-intensity oscillation due to the breakdown and decomposition of the tip vortex structures plays a critical role in the fluctuation of the cascade performance. By means of proper orthogonal decomposition, it is found that the IBL enhances the fluctuation of small-scale vortex structures related to flow separation and leakage flow and makes the stability of the tip flow field worse.

Vortex structure control based bleed in axial compressor cascade with tip clearance using large eddy simulation

Abstract As an essential component, the bleed system plays a critical role in supplying turbine cooling air, guaranteeing stage matching, pressurizing the cabin, and de-icing at the wing and engine inlet. However, the extraction of the bleeding air from the compressor causes the engine efficiency degradation and thrust deficit. Therefore, flow control based on bleed is conducted to compensate the bleed induced disadvantages. The influence of the circumferential bleeding slot location on the tip leakage vortex and passage vortex controlling in a compressor cascade with the tip clearance is numerically studied using large eddy simulation. Three bleed configurations and the smooth casing configuration are investigated. 17.11% loss reduction is obtained through bleeding at 10% c x upstream of the blade leading edge with a bleeding rate of 2.76%. The vortex structures and flow patterns are compared and analyzed to reveal the controlling mechanism. Subsequently, the axial vorticity and loss evolution is discussed, and the interaction between the primary flow and bleeding air is revealed. It’s found that bleeding slot placed within the blade passage is exposed into a highly static pressure gradient, and this causes the bleeding air flows into and spills out the bleeding slot and leads to unnecessary loss. Moreover, the influence of large bleeding rate and inlet boundary layer is assessed.

Numerical Investigation of Inlet Boundary Layer in an Axial Compressor Tandem Cascade

To explore the inlet boundary layer (IBL) influence on the tandem cascade aerodynamic performance, this paper took the high subsonic compressor NACA65 K48 cascade and its modified tandem cascade as the research object. The effects of the IBL thickness and the skewed IBL on the aerodynamic performance of the original cascade and tandem cascade were analyzed based on the numerical method. The results show that the tandem cascade effective design makes it better than the original cascade in the aerodynamic performance under different IBL conditions. Compared with the collateral IBL, the skewed IBL can effectively improve the aerodynamic performance of the original cascade and tandem cascade by suppressing the endwall cross flow, but an increase in the IBL thickness will suppress this advantage. In addition, the increase of incidence angle or the IBL thickness will make the tandem cascade forward blade corner separation more serious and cause the flow passage to be blocked, which seriously affects the rear blade diffusing capacity. In general, the IBL thickness is positively correlated with the tandem cascade total pressure loss and negatively correlated with the static pressure rise (except for the −6° incidence angle). The skewed IBL can effectively reduce the total pressure loss and increase the static pressure rise within −4°~7° incidence angle, but the law is opposite at a −6° incidence angle. At a 0° or 2° incidence angle, the performance improvement effect of the skewed IBL on the tandem cascade is the best, and this positive effect diminishes as it tends towards a larger positive incidence angle or a smaller negative incidence angle.

Nacelle intake flow separation reduction at cruise condition using active flow control

Turbofan engine intakes are designed to provide separation-free flow at the fan face over a wide range of operating conditions. But at some off-design conditions, like at high flight speeds and high angles of attack (AoA), the aero engine intake may encounter flow separation. This boundary layer separation inside the nacelle inlet of an aircraft engine can lead to a large number of undesirable outcomes like reduction in fan efficiency, engine stall and high levels of stress on the fan blades. Active flow control is a promising solution to reduce inlet boundary layer separation and the associated fan-face flow distortion at such off-design conditions. By blowing pressurized air into the intake near the separation point, the boundary layer is energized and separation can be controlled. This study investigates the applicability of lip blowing, an active flow control technique, to control intake separation and flow distortion at the fan-face. First, intake separation was triggered in a 3D CFD model based on the NASA Common Research Model (CRM) using high AoA cases at cruise condition (Mach number 0.85, Mass flow capture ratio ∼0.7) and the features of separated flow were analyzed. Thereafter, active flow control was introduce to the intake in the form of two types of lip blowing, direct and pitched blowing. The efficacy of lip blowing at achieving separation control in an ultra high bypass ratio turbofan engine intake has been established through this study. The present paper also examines the significance of blowing parameters like the type of blowing, blowing pressure ratio, and blowing slot dimension, at different angles of attack to identify the critical control parameters. Our research successfully establishes proof of concept by demonstrating the feasibility of using lip blowing for separation control in aero-intakes, via numerical modelling. Furthermore, this study also provides crucial insights regarding the important variables to be considered for future experimental studies, and also for detailed studies covering a wider range of operating and blowing conditions.

Numerical investigation of flow around two cylinders in tandem above a scoured bed

Flow mechanisms around two cylinders in tandem arrangement above a scoured bed have been investigated using the three-dimensional unsteady Navier–Stokes equations with the Spalart–Allmaras improved delayed detached-eddy simulation model. A turbulent inlet boundary layer generation method is adopted to obtain more realistic inlet boundary conditions. First, uniform flow over a single cylinder at Re = 3900 and flow over a single cylinder above scoured beds at Re = 6000 were simulated to validate the numerical model, boundary layer generation method, and mesh density effect. Second, two cylinders in the tandem arrangement above scoured beds with six different pitch ratios L/D are investigated numerically in terms of instantaneous vortex characteristics, the hydrodynamic force, and time-averaged flow fields. The simulation results in scoured beds are compared with simulations under the near flat wall and wall-free conditions. The major findings can be summarized as follows. (1) When L/D≤2.0, the wake of two tandem cylinders is dominated by the intermittent shedding, and the downstream sand dune in the scoured bed hinders the Kármán vortex formation at the rear of the downstream cylinder. Lift force fluctuations of the two cylinders have small amplitudes, and their spectra show a multi-peak distribution and no dominant peak frequency in the spectrum of the upstream cylinder. A squarish cavity-like recirculation zone is formed between two cylinders at L/D=2.0. (2) When L/D≥3.0, the periodic vortex shedding is evident in the wake of the upstream cylinder, and the small sand berm between two cylinders has an impact on the bottom shear layer of the upstream cylinder. The downstream cylinder is periodically impacted by the vortices shed from the upstream cylinder. Lift force spectra of the upstream and downstream cylinders have the same peak frequency. (3) Due to the influence of scoured bed and the inlet boundary layer, the time-averaged lift coefficient of the upstream cylinder remains negative when L/D≥1.5, and the critical spacing for drag inversion is relatively smaller compared with under wall-free conditions. The negative pressure coefficient values of the upstream cylinder are smaller than the values in near flat wall and wall-free conditions.

The Effect of Inlet Conditions on Turbine Endwall Loss

Abstract There can be significant variation and uncertainty in the flow conditions entering a blade row. This article explores how this variability can affect endwall loss in axial turbines. A computational study of three cascades with collinear inlet boundary layers is conducted. Endwall loss varies by more than a factor of 3 depending on the inlet conditions. This variation is caused by dissipation of secondary kinetic energy (SKE). The results can be understood by observing that the inlet conditions predominantly control how secondary vorticity is distributed within the blade passage. Modestly thick inlet boundary layers with high shape factor tend to displace vorticity toward the center of the passage. This displacement reduces vorticity cancellation, increasing secondary velocities, and SKE. A general method is formulated to estimate SKE in preliminary design. Optimum aspect ratio is shown to depend on the inlet boundary condition. Strategies to reduce endwall loss and minimize sensitivity to inlet conditions are then highlighted.

Effects of the Installation Location of a Dielectric Barrier Discharge Plasma Actuator on the Active Passage Vortex Control of a Turbine Cascade at Low Reynolds Numbers

Because axial flow turbines are widely used as the main components of jet engines and industrial gas turbines, their energy reduction effect is significant, even with a slight performance improvement. These turbines operate over a wide range of Reynolds numbers. However, at low Reynolds numbers below 1 × 105, the aerodynamic characteristics deteriorate greatly, due to the flow separation of the boundary layer on the blade suction surface and an increase in the secondary flow. In this study, an experiment to reduce the passage vortex was conducted using a dielectric barrier discharge plasma actuator, which is expected to operate with a new innovative active flow control technology. The plasma actuator was installed on the endwall of a linear turbine cascade in the test section of a wind tunnel. From the velocity distribution measured using particle image velocimetry, the secondary flow vector, turbulence intensity, and vorticity were analyzed. The input voltage and frequency of the plasma actuator were fixed at 12 kVp-p and 10 kHz, respectively. In particular, the optimum installation location of the plasma actuator was examined from upstream to mid-passage positions of the turbine cascade (normalized axial location of Z/Cax = −0.35 to 0.51). In addition, the effect of the Reynolds number was examined by varying it between Reout = 1.8 × 104 and 3.7 × 104. From the experimental results, it was found that the optimum location of the plasma actuator was immediately before the blade leading edge (Z/Cax = −0.20 to −0.06). This is because the inlet boundary layer can be accelerated near the blade leading edge, weakening the horseshoe vortex which initially causes the passage vortex. At a higher Reynolds number, the passage vortex suppression effect of the plasma actuator is weakened, because the flow induced by the plasma actuators becomes relatively weaker as the mainstream velocity increases with an increase in the Reynolds number.

Unsteady Motion of Shock Wave for a Supersonic Compression Ramp Flow Based on Large Eddy Simulation

A large eddy simulation (LES) is conducted to investigate shock wave/turbulent boundary layer interaction in a 24° compression ramp at a high inlet Mach number of Ma=2.9. The recycling/rescaling Method is used as the inflow turbulence generation technique. The shock wave structure and boundary layer flow in the interaction region are studied by flow visualization methods, such as vortex recognition and numerical schlieren. The distributions of turbulent kinetic energy and Reynolds normal stresses at different streamwise locations are analyzed. The results show that a strong anisotropy turbulent flow appears in the reattached boundary layer after the interaction of the shock wave. The large-scale unsteady motion of the separation shock wave is analyzed by using the intermittent factor and power spectrum. It is found that the shock moves around the averaged separation position, and the length scale is equal to 72% of the inlet boundary layer thickness. The power spectrum analysis reveals the existence of low-frequency instability in the separation region.

Influence of the inlet boundary layer on non-axisymmetric endwall contouring effects in a linear compressor cascade

The modern aero-engine compressors are commonly designed to be highly loaded which can introduce strong three-dimensional flow separation, and effective flow control is of vital importance in improving compressor performance. Non-axisymmetric endwall contouring nowadays is conducted on axial compressors to control the flow separation. In this paper, a contoured endwall is designed based on a linear CDA compressor cascade. The genetic algorithm and ANN surrogate model are used in the optimization process of the hub contouring, based on RANS simulations. Three inlet boundary layers which follow the fully turbulent boundary layer assumption are generated with different thickness, and the influence mechanism is numerically investigated. The optimal endwall with streamwise concave near the suction side provides great performance improvement at all inlet boundary layers at the aerodynamic design point. The endwall flow near the suction side is accelerated by the contoured endwall with cross-flow pressure gradient altered, suppressing the corner separation. The main pressure loss is closely related to one streamwise vortex that blocks the endwall flow. A thicker inlet boundary layer can induce early forming and increased strength on the streamwise vortex, enhancing the cross-flow movement and its interaction with the blade surface flow. In addition, the pressure loss near endwall in the inlet boundary layer is found more dominant on the endwall contouring effect than the boundary layer thickness.