Boundary Layer Separation Region Research Articles

Numerical Study of the Boundary Layer Separation of a Low‐Pressure Turbine Cascade at Different Incidence Angles

Modern low‐pressure turbine (LPT) pursues high blade loading to improve its efficiency while the possibility of boundary layer separation on the blades is still a challenging problem. The incidence angle is one of the main factors affecting the flow separation and laminar‐turbulent transition, which is of key importance for the design of LPT. In this work, based on the initial simulation results obtained using the Reynolds‐averaged Navier–Stokes (RANS) method, the flow past a single LPT cascade at three representative incident angles (i = −10°, 0°, +10°) is investigated using large‐eddy simulation (LES) method with approximately 26 million structured grids. The simulations are carried out with Open Source Field Operation and Manipulation (OpenFOAM) for an off‐design low Reynolds number condition Re = 40,000 and Ma = 0.5. The time‐averaged aerodynamic performances and three‐dimensional instantaneous flow characteristics of the LPT cascade are discussed in detail, and the results show that the blade loading and the transverse pressure gradient (TPG) from the pressure side to the suction side of leading edge (LE) increase with increasing incidence angle. At the incidence angle of +10°, the LPT cascade undergoes separation and reattachment from the LE to around 20% Cax on the suction surface, which weakens the separation phenomenon at the rear half of the suction surface and the total pressure loss downstream of the trailing edge (TE). It is worth noting that the RANS method underpredicts the blade boundary layer separation region compared with LES under these extreme conditions. These results may offer off‐design condition data for the future improvement of RANS models or experimental studies.

Statistical evaluation of stability margin of a multi-stage compressor with geometric variability using adaptive polynomial chaos-Kriging model

Compressed air energy storage systems must promptly adapt to power network demand fluctuations, necessitating a high surge margin in the compression system to ensure safety. It is challenging to completely eliminate blade geometric variations caused by limited machining precision, the important effects of which should be considered during aerodynamic shape design and production inspection. The present paper explores the uncertainty impact of geometric deviations on the stability margin of a multi-stage axial compressor at a low rotational speed. Initially, an adaptive polynomial chaos expansion-based universal Kriging model is introduced, and its superior response performance in addressing high-dimensional uncertainty quantification problems is validated through rigorous analytical and engineering tests. Then, this model is used to statistically evaluate the stability margin improvement (SMI) of the compressor due to the Gaussian and realistic geometric variabilities separately. The results show that the mean and standard deviation of SMI are −0.11% and 0.5% under the Gaussian geometric variability, while those are 0.33% and 0.39% under the realistic variability. For both the geometric variabilities, the stagger angle and maximum thickness deviations of the first-stage rotor are the most influential parameters controlling the uncertainty variations in the stability margin. Finally, the underlying impact mechanism of the influential geometric deviations is investigated. The variation in the stability margin caused by the geometric deviations primarily results from the alteration of inlet incidences, affecting the size of the tip leakage vortex blockage and boundary-layer separation regions near the blade tip of the first-stage rotor.

Downstream influence of turbulent flow past vortex generators

Abstract Vortex generators (VGs) are typically positioned upstream of a boundary layer separation region. Their effectiveness depends on incoming flow conditions (Mach number, Reynolds number, and boundary layer characteristics), geometrical configuration (vane or ramp, height, width, and angle of incidence), and spacing. Device-induced vortices and following decay allow VGs to be used as a passive control device. This study uses pressure-sensitive paint to determine the global surface pressure pattern for a flat plate flow in the presence of VGs (counter-rotating vanes, co-rotating vanes and ramps). The freestream Mach number is 0.64 and 0.83. The ratio between the height of the VGs and the incoming boundary layer thickness is 0.2, 0.5, and 1.0. The standard deviation in the spanwise pressure in the streamwise direction is used to determine the downstream influence of turbulent flow past VGs. Increasing the height of VGs causes device-induced vortices to propagate farther downstream.

Turbulence Model Comparative Study for Complex Phenomena in Supersonic Steam Ejectors with Double Choking Mode.

Scholars usually ignore the non-equilibrium condensing effects in turbulence-model comparative studies on supersonic steam ejectors. In this study, a non-equilibrium condensation model considering real physical properties was coupled respectively with seven turbulence models. They are the k-ε Standard, k-ε RNG, k-ε Realizable, k-ω Standard, k-ω SST, Transition SST, and Linear Reynolds Stress Model. Simulation results were compared with the experiment results globally and locally. The complex flow phenomena in the steam ejector captured by different models, including shock waves, choking, non-equilibrium condensation, boundary layer separation, and vortices were discussed. The reasons for the differences in simulation results were explained and compared. The relationship between ejector performance and local flow phenomena was illustrated. The novelty lies in the conclusions that consider the non-equilibrium condensing effects. Results show that the number and type of shock waves predicted by different turbulence models are different. Non-equilibrium condensation and boundary layer separation regions obtained by various turbulence models are different. Comparing the ejector performance and the complex flow phenomena with the experimental results, the k-ω SST model is proposed to simulate supersonic steam ejectors.

Stall Behavior in an Ultrahigh-Pressure-Ratio Centrifugal Compressor: Backward-Traveling Rotating Stall

AbstractThis paper describes the stall mechanism in an ultrahigh-pressure-ratio centrifugal compressor, composed of a double-splitter impeller, radial diffuser, and axial diffuser. A model comprising all impeller and diffuser blade passages is used to conduct unsteady simulations that trace the onset of instability in the compressor. Backward-traveling rotating stall waves appear at the inlet of the radial diffuser when the compressor is throttled. Six stall cells propagate circumferentially at approximately 0.7% of the impeller rotation speed. The detached shock of the radial diffuser leading edge and the number of stall cells determine the direction of stall propagation, which is opposite to the impeller rotation direction. Dynamic mode decomposition is applied to instantaneous flow fields to extract the flow structure related to the stall mode. This shows that intensive pressure fluctuations concentrate in the diffuser throat as a result of changes in the detached shock intensity. The diffuser passage stall and stall recovery are accompanied by changes in incidence angle and shock wave intensity. When the diffuser passage stalls, the shock-induced boundary–layer separation region near the diffuser vane suction surface gradually expands, increasing the incidence angle and decreasing the shock intensity. The shock is pushed from the diffuser's throat toward the diffuser leading edge. When the diffuser passage recovers from the stall, the shock wave gradually returns to the diffuser throat, with the incidence angle decreasing and the shock intensity increasing. Once the shock intensity reaches its maximum, the diffuser passage suffers severe shock-induced boundary–layer separation and stalls again.

Parametric Study of Nonequilibrium Shock Interference Patterns over a Fuselage-and-Wing Conceptual Vehicle

Predicting shock/shock and shock/boundary-layer interactions in gas flows that surround high-speed vehicles is key in aerodynamic design. Under typical hypersonic conditions, these flow structures are influenced by complex nonequilibrium phenomena leading to high-temperature effects. In this work, the conceptual Bedford wing-body vehicle is studied to analyze flow patterns in shock/shock and shock/boundary-layer interactions with thermochemical nonequilibrium. A parametric computational fluid dynamics study is carried out for different hypersonic operating conditions, with respect to the freestream Mach number. Simulations are performed with the SU2-NEMO solver coupled to the Mutation++ library, which provides all the necessary thermodynamic, kinetic, and transport properties of the mixture and chemical species. The Adaptive Mesh Generation library is used for automatic anisotropic mesh adaptation. Numerical results show that increasing the freestream Mach number from 4 to 10 leads to changes in the shock layer, locations of shock impingement, and regions of boundary-layer separation. Despite these changes, the change in freestream Mach number has little impact on the overall shock interaction structures.

Passage shock wave/boundary layer interaction control for transonic compressors using bumps

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow, which always accompanies aerodynamic performance penalties. A loss reduction method for smearing the passage shock foot via Shock Control Bump (SCB) located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation. The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model. Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes (RANS) method to reveal flow physics of SCB. Comparisons and analyses on simulation results have also been carried out, showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer. It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100% design rotor speed, and then slightly worsens with decrease of rotatory speed till both equal below 60% rated speed. The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

Stability of space-periodic flow with separation of a laminar boundary layer

An unstable laminar flow over a wavy region of a plate placed parallel to the oncoming air flow is studied. The results have been obtained by the hot-wire method in a low-turbulent wind tunnel at low subsonic velocities. Flow near a wall with transverse undulation, whose amplitude is sufficiently large and can contribute to the formation of local regions of boundary layer separation, is considered. The external harmonic (acoustic) forcing of the periodic flow leads to the suppression of perturbations that dominate in the spectrum of velocity disturbances near the model surface. The results of this work may be useful in developing methods for controlling spatially inhomogeneous flows.

Criteria for designing low-loss and wide operation range variable inlet guide vanes

Critical factors influencing the variable inlet guide vane (VIGV) profile loss at high incidence condition were studied by using numerical methods and a practical design criterion for designing wide low-loss operation range VIGVs in axial-flow compressor was proposed. At first, to acquire research samples, a series of airfoils with different shapes were generated for two selected representative VIGV cascade cases. Steady simulations based on Reynolds-Averaged Navier–Stokes method, carried out by commercial software CFX and validated with experimental data after grid independent study, were first conducted to predict the aerodynamic performances, the surface velocity distributions and the boundary-layer behaviors of the generated airfoils. Based on the simulated results, the influences of geometric parameters on airfoil performances were analyzed and the geometric features of low-loss VIGV airfoil were revealed. Further analysis indicated that the magnitude of airfoil loss at high incidence condition were mainly influenced by the scales of two boundary-layer separation regions: one at the leading edge caused by the high adverse pressure gradient induced by the suction spike and the other one caused by the adverse pressure gradient induced by the re-acceleration flow. To reveal the influence of the suction spike and the re-acceleration flow on the scales of separation regions, two practical parameters Dspike and Are were defined. It was found that there exists an optimized range of the Dspike and Are which could keep the separation flow to a small scale at high incidence condition and can be used as a surface velocity design criterion for designing wide low-loss operation range VIGVs. Moreover, the methods for choosing the airfoil geometric parameters to achieve the preferred surface velocity distribution were discussed. Finally, the developed design criterion was used to guide the airfoil modification of an axial-flow compressor VIGV and achieved an average of 19%, 52% and 73% loss coefficient reduction at three high stagger angle operating points, which confirms the applicability and effectiveness of the design criterion in three-dimensional environment.

Intermittent back-flash phenomenon of supersonic combustion in the staged-strut scramjet engine

The scramjet-powered hypersonic vehicle depends on the sustained and stable supersonic combustion processes, which requires a balance between the flame propagation velocity and the fluid speed. Flame propagation inside a kerosene-fueled staged-strut scramjet engine was experimentally studied in Mach 3.0 incoming flow with the stagnation temperature of 1899 K. The fuel mass flow rate was linearly controlled by an adjustable venturi during the dynamic injection experiment. Two inspection windows were installed on the side and top walls of the combustor respectively. A mirror was utilized to change the light path from the top window to be horizontal to capture the flame structure in the two windows simultaneously. A high-speed imaging camera was employed to capture the flame propagation process described in this short communication. Video imaging showed that the flame was mainly stabilized in the boundary-layer separation region when the upstream equivalence ratio was less than 0.2. In particularly, this short communication reported an intermittent back-flash phenomenon while the fuel equivalence ratio was turned up near to 0.2. The back-flash flame was suggested to be detonation wave by a primary C-J analysis.

Flame Stabilization and Propagation in Dual-Mode Scramjet with Staged-Strut Injectors

To investigate the flame stabilization and propagation in a dual-mode scramjet with two-staged-strut injectors, a series of fuel dynamic adjustment experiments was conducted at the inlet Mach numbers of 2.0 and 3.0, and the corresponding inflow stagnation temperatures were 1231 and 1899 K, respectively. A linearly adjustable venturi was designed to realize the continuous control on the fuel mass flow rate. A high-speed imaging camera was used to record the flame structure during the combustion. The effects of the local equivalence ratio, the adjustment process of the fuel mass flow rate, and the inflow conditions on the combustor performance were experimentally analyzed. The results show that the flame can only be stabilized in the recirculation region or two thin regions of the upstream strut under the lower inflow enthalpy condition, and the flame is blown out when the fuel flow rate is below the lean blowout limit. Comparatively, under the higher inflow enthalpy condition, an additional stabilized flame could be found in the boundary-layer separation region of the side walls, and the combustion waves were observed intermittently. The sudden changes in the wall pressure profiles and an obvious hysteresis phenomenon were found in the fuel dynamic adjustment procedure under the lower inflow enthalpy condition. However, the flame structure and corresponding wall pressure profile changed continuously, illustrating a much smaller hysteresis phenomenon under the higher inflow enthalpy condition.

Investigations Of A Flow Around The Flying Wing Model At Natural Reynolds Numbers

Studies of flow structure near the surface model of drones have been conducted in the subsonic wind tunnel. Experimental data of flow visualization depending on attack angle were obtained by oil film method. Evolution of flow patterns inside the regions of boundary layer separation in different configurations was demonstrated. Ability of separation control and reduction zones of flow separation by local influence (ribs and cones) has been also studied on the wing surface.

Numerical Estimation of Aerodynamic Characteristics of Footbridge Deck

For estimation of aerodynamic characteristics of cable-stayed footbridge deck a computational fluid dynamics (CFD) has been used. An incompressible fluid flow with Navier-Stokes equations has been applied. An adequate numerical model has been created to obtain accurate values of aerodynamic characteristics. Preliminary determination of simulation parameters have been estimated using laminar fluid flow model. Subsequently, Smagorinsky large-eddy simulation (LES) turbulent model has been applied with different simulation parameters to obtain converged values. The boundary layer separation regions and downwind vortex shedding has been observed.

Simulation of Secondary and Separated Flow in Diffusing S Ducts

The focus of this paper is on the numerical simulation of compressible flow in diffusing S-ducts; this flow is characterized by secondary flow as well as regions of boundary-layer separation. The S-duct geometry produces streamline curvature and an adverse pressure gradient resulting in these flow characteristics. Two S-duct geometries are employed in this investigation: one was used in an experimental study conducted at NASA John H. Glenn Research Center at Lewis Field in the early 1990s, and the other is a benchmark configuration proposed by AIAA Propulsion Aerodynamics Workshop to assess the accuracy and best practices of computational fluid dynamics solvers. The computational fluid dynamics flow solver ANSYS-FLUENT is employed in the investigation of compressible turbulent flow through the S duct. A second-order accurate, steady, density-based solver is employed in a finite-volume framework. The three-dimensional Reynolds-averaged Navier–Stokes equations are solved on a structured mesh with a number of turbulence models, namely, the Spalart–Allmaras, , shear stress transport, and transition shear stress transport models, and the results are compared with the available experimental data. The computed results capture the flowfield and pressure recovery with acceptable accuracy when compared to the experimental data. The turbulence model giving the best results is identified.

Simulation of secondary and separated flow in a diffusing S-duct using four different turbulence models

The focus of this article is on the numerical simulation of compressible flow in a diffusing S-duct inlet; this flow is characterized by secondary flow as well as regions of boundary layer separation. The S-duct geometry produces streamline curvature and an adverse pressure gradient resulting in these flow characteristics. The geometry used in this investigation is based on a NASA Glenn Research Center experimental diffusing S-duct that was studied in the early 1990s. The computational fluid dynamics flow solver ANSYS - FLUENT is employed in the investigation of compressible flow through the S-duct. A second-order accurate, steady, density-based solver is employed in a finite-volume framework. The three-dimensional Reynolds-Averaged Navier-Stokes equations are solved on a structured mesh with a number of turbulence models, namely the Spalart–Allmaras (SA), k-ɛ, k-ω SST, and Transition SST models, and the results are compared with the experimental data. The computed results capture the flow field and pressure recovery with acceptable accuracy when compared with the experimental data. The turbulence model giving the best results is identified.

Numerical study of plasma‐fluid behavior and generation characteristics of the closed‐loop MHD electrical power generator

AbstractTime‐dependent r‐z two‐dimensional numerical simulations by the LES technique have been carried out in order to clarify the plasma fluid behavior and power generation characteristics of a disk MHD generator under the rated operating conditions demonstrated in the closed‐loop experimental facility at Tokyo Institute of Technology. The generator currently installed could suffer from nonuniform and low electrical conductivity and from boundary layer separation even under the rated operation conditions. The large amount of generated electric power is consumed in the boundary layer separation region, which reduces the net output power. Reducing the back pressure and improving the inlet plasma conditions reliably provides better generator performance. The influence of a 90° bend downstream duct on the generator performance is found not to be particularly great. © 2010 Wiley Periodicals, Inc. Electr Eng Jpn, 174(4): 37–44, 2011; Published online in Wiley Online Library (wileyonlinelibrary.com). DOI 10.1002/eej.21041

Prograde and retrograde flow past cylindrical obstacles on a β-plane

The problem of viscous prograde (eastward) and retrograde (westward) flow past a cylindrical obstacle on a β-plane is considered. The barotropic vorticity equation is solved using a numerical method that combines finite difference and spectral methods. A modified version of the β-plane approximation is proposed to avoid computational difficulties associated with the traditional β-plane approximation. Numerical results are presented and discussed for flow past a circular cylinder at selected Reynolds numbers (Re )a nd non-dimensional β-parameters ( ˆ β) as well as for flow past an elliptic cylinder of a fixed aspect ratio (r = 0.2) at inclination angles of ±15 ◦ , 90 ◦ and selected Reand ˆ β. In prograde flows, it is found that the β-effect acts to suppress boundary-layer separation and to excite a standing Rossby lee wavetrain, as observed in previ- ous works. In retrograde flows, the boundary-layer separation region is elongated and westward propagating Rossby waves are excited.

Large-Eddy Simulation of Jet Mixing in Supersonic Crossflows

Large-eddy simulation of an underexpanded sonic jet injection into supersonic crossflows is performed to obtain insights into key physics of the jet mixing. A high-order compact differencing scheme with a recently developed localized artificial diffusivity scheme for discontinuity-capturing is used. Progressive mesh refinement study is conducted to quantify the broadband range of scales of turbulence that are resolved in the simulations. The simulations aim to reproduce the flow conditions reported in the experiments of Santiago and Dutton [Santiago, J. G., and Dutton, J. C., Velocity Measurements of a Jet Injected into a Supersonic Crossflow, Journal of Propulsion and Power, Vol.132,1997, pp. 264―273] and elucidate the physics of the jet mixing. A detailed comparison with these data is shown. Statistics obtained by the large-eddy simulation with turbulent crossflow show good agreement with the experiment, and a series of mesh refinement studies shows reasonable grid convergence in the predicted mean and turbulent flow quantities. The present large-eddy simulation reproduces the large-scale dynamics of the flow and jet fluid entrainment into the boundary-layer separation regions upstream and downstream of the jet injection reported in previous experiments, but the richness of data provided by the large-eddy simulation allows a much deeper exploration of the flow physics. Key physics of the jet mixing in supersonic crossflows are highlighted by exploring the underlying unsteady phenomena. The effect of the approaching turbulent boundary layer on the jet mixing is investigated by comparing the results of jet injection into supersonic crossflows with turbulent and laminar crossflows.

Numerical Study of Plasma-Fluid Behavior and Generation Characteristics of the Closed Loop MHD Electrical Power Generator

Time dependent r-z two-dimensional numerical simulations with LES technique have been carried out in order to clarify the plasma fluid behavior and power generation characteristics of the disk MHD generator under the rated operation conditions demonstrated in the closed loop experimental facility at Tokyo Tech. The generator currently installed could suffer from the non-uniform and low electrical conductivity, and the boundary layer separation even under the rated operation conditions. The large amount of generated electric power is consumed in the boundary layer separation region, which reduces a net output power. Reducing the back pressure and improving the inlet plasma conditions surely provide the higher generator performance. The influence of 90 degree bend downstream duct on the generator performance is found to be not marked.

Effect of the Location of an Incident Shock Wave on Combustion and Flow Field of Wall Fuel-Injection

Experimental and numerical studies of the interaction between combustion of a hydrogen jet and an incident shock wave were performed. In the case that an incident shock wave was introduced upstream of the injection slot, the boundary-layer separation region was extensively expanded. The penetration height of the Mach disk with an incident shock wave was less than that without an incident shock wave. Combustion was confirmed when the incident shock wave was introduced downstream of the fuel injection slot, while, with the incident shock wave upstream of fuel injection slot, combustion was not confirmed. The mechanism of these phenomena was discussed based on the results of numerical simulation in terms of the residence time in the separation region near the fuel injection slot.