Experimental and Numerical Investigations of Shock-Wave Boundary Layer Interactions in a Highly Loaded Transonic Compressor Cascade

Experimental and numerical investigations were conducted to investigate the variations of shock-wave boundary layer interaction(SBLI) phenomena in a highly loaded transonic compressor cascade with Mach numbers.The schlieren technique was used to observe the shock structure in the cascade and the pressure tap method to measure the pressure distribution on the blade surface.The unsteady pressure distribution on blade surface was measured with the fast-response pressure-sensitive paint(PSP) technique to obtain the unsteady pressure distribution on the whole blade surface and to capture the shock oscillation characteristics caused by SBLI.In addition,the Reynolds Averaged Navier Stokes simulations were used to compute the three-dimensional steady flow field in the transonic cascade.It was found that the shock wave patterns and behaviors are affected evidently with the increase in incoming Mach number at the design flow angle,especially with the presence of the separation bubble caused by SBLI.The time-averaged pressure distribution on the blade surface measured by PSP technique showed a symmetric pressure filed at Mach numbers of 0.85,while the pressure field on the blade surface was an asymmetric one at Mach numbers of 0.90 and 0.95.The oscillation of the shock wave was closely with the flow separation bubble on the blade surface and could transverse over nearly one interval of the pressure taps.The oscillation of the shock wave may smear the pressure jump phenomenon measured by the pressure taps.

Experimental Investigation of the Transonic Shock Oscillation Characteristics in a Heavy-Duty Gas Turbine Compressor Cascade

This paper presents an experimental study of the self-sustained transonic shock oscillating behaviors in a heavy-duty gas turbine compressor cascade under the inlet Mach number of 0.85,0.90 and 0.95.The transonic shock patterns and the surface flow structures are captured by schlieren imaging and oil flow visualization.The time-averaged and instantaneous transonic shock oscillating behaviors at the near choke point and the near stall point are investigated by the Anodized Aluminum Pressure-Sensitive Paint(AA-PSP)surface pressure measurement.The normal passage shock dominant pattern and the detached bow shock dominant pattern at the near choke point and the near stall point are experimental characterized,respectively.The passage shock oscillation behaviors at the near choke point have been observed to undergo periodic pressure perturbations of the shock shift between the upstreamλshock feet mode and the downstreamλshock feet mode.The detached bow shock oscillation behaviors at the near stall point have been observed to undergo the pressure perturbations of the shock cycle movement between the upstream detached bow shock mode and the downstream detached bow shock mode.The differences between the shock shift mode and the shock cycle movement mode lead to the different streamwise oscillation travel ranges and different shock intensity variations under the same inlet Mach number.

Effect of tip geometry and tip clearance on aerodynamic performance of a linear compressor cascade

The tip leakage flow between a blade and a casing wall has a strong impact on compressor pressure rise capability, efficiency, and stability. Consequently, there is a strong motivation to look for means to minimize its impact on performance. This paper presents the potential of passive tip leakage flow control to increase the aerodynamic performance of highly loaded compressor blades. Experimental investigations on a linear compressor cascade equipped with blade winglets mounted to the blade tips have been carried out. Results for a variation of the tip clearance and the winglet geometry are presented. Current results indicate that the use of proper tip winglets in a compressor cascade can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction-side winglets are aerodynamically superior to the pressure-side or combined winglets. The suction-side winglets are capable of reducing the exit total pressure loss associated with the tip leakage flow and the passage secondary flow to a significant degree.

Numerical study of corner separation in a linear compressor cascade using various turbulence models

Three-dimensional corner separation is a common phenomenon that significantly affects compressor performance. Turbulence model is still a weakness for RANS method on predicting corner separation flow accurately. In the present study, numerical study of corner separation in a linear highly loaded prescribed velocity distribution （PVD） compressor cascade has been investigated using seven frequently used turbulence models. The seven turbulence models include Spalart Allmaras model, standard k-e model, realizable k-e model, standard k-to model, shear stress transport k co model, v2-fmodel and Reynolds stress model. The results of these turbulence models have been compared and analyzed in detail with available experimental data. It is found the standard k-1： model, realizable k-e model, v2-f model and Reynolds stress model can provide reasonable results for predicting three dimensional corner separation in the compressor cascade. The Spalart-Allmaras model, standard k-to model and shear stress transport k-w model overesti- mate corner separation region at incidence of 0°. The turbulence characteristics are discussed and turbulence anisotropy is observed to be stronger in the corner separating region.

Influence of leading edge with real manufacturing error on aerodynamic performance of high subsonic compressor cascades

To investigate the influence of real leading-edge manufacturing error on aerodynamic performance of high subsonic compressor blades,a family of leading-edge manufacturing error data were obtained from measured compressor cascades.Considering the limited samples,the leadingedge angle and leading-edge radius distribution forms were evaluated by Shapiro-Wilk test and quantile–quantile plot.Their statistical characteristics provided can be introduced to later related researches.The parameterization design method B-spline and Bezier are adopted to create geometry models with manufacturing error based on leading-edge angle and leading-edge radius.The influence of real manufacturing error is quantified and analyzed by self-developed non-intrusive polynomial chaos and Sobol’indices.The mechanism of leading-edge manufacturing error on aerodynamic performance is discussed.The results show that the total pressure loss coefficient is sensitive to the leading-edge manufacturing error compared with the static pressure ratio,especially at high incidence.Specifically,manufacturing error of the leading edge will influence the local flow acceleration and subsequently cause fluctuation of the downstream flow.The aerodynamic performance is sensitive to the manufacturing error of leading-edge radius at the design and negative incidences,while it is sensitive to the manufacturing error of leading-edge angle under the operation conditions with high incidences.

Reduced order model for unsteady aerodynamic performance of compressor cascade based on recursive RBF

Based on Recursive Radial Basis Function(RRBF)neural network,the Reduced Order Model(ROM)of compressor cascade was established to meet the urgent demand of highly efficient prediction of unsteady aerodynamics performance of turbomachinery.One novel ROM called ASA-RRBF model based on Adaptive Simulated Annealing(ASA)algorithm was developed to enhance the generalization ability of the unsteady ROM.The ROM was verified by predicting the unsteady aerodynamics performance of a highly-loaded compressor cascade.The results show that the RRBF model has higher accuracy in identification of the dimensionless total pressure and dimensionless static pressure of compressor cascade under nonlinear and unsteady conditions,and the model behaves higher stability and computational efficiency.However,for the strong nonlinear characteristics of aerodynamic parameters,the RRBF model presents lower accuracy.Additionally,the RRBF model predicts with a large error in the identification of aerodynamic parameters under linear and unsteady conditions.For ASA-RRBF,by introducing a small-amplitude and highfrequency sinusoidal signal as validation sample,the width of the basis function of the RRBF model is optimized to improve the generalization ability of the ROM under linear unsteady conditions.Besides,this model improves the predicting accuracy of dimensionless static pressure which has strong nonlinear characteristics.The ASA-RRBF model has higher prediction accuracy than RRBF model without significantly increasing the total time consumption.This novel model can predict the linear hysteresis of dimensionless static pressure happened in the harmonic condition,but it cannot accurately predict the beat frequency of dimensionless total pressure.

Application of shear-sensitive liquid crystal coating to visualization of transition and reattachment in compressor cascade

The present paper aims at introducing Shear-Sensitive Liquid Crystal Coating(SSLCC)technology into compressor cascade measurement for the first time and serves as a basis for better understanding of the influence from the boundary layers. Optical path layout, which is the most significant difficulty in internal flow field measurement, will be solved in this paper by selfdesigned image acquisition device. Massive experiments with different Mach number and incidence are conducted at a continuous subsonic cascade wind tunnel to capture the boundary layer phenomenon. Image processing methods, such as Three-Dimensional(3-D) reconstruction and Hue conversion, are used to improve the accuracy for transition position detection. The analysis of the color-images indicates that complex flow phenomena including transition, flow separation,and reattachment are captured successfully, and the effect of Mach number and incidence on the boundary layer flow is also discussed. The results show that: the Mach number has a significant effect on transition position; the incidence has little effect on transition position, but it has a great impact on the transition distance and leading-edge separation; influenced by the end-walls, the reattachment occurs in advance under positive angle of attack conditions.

Numerical Study of Dual Sweeping Jet Actuators for Corner Separation Control in Compressor Cascade

Unsteady behaviors are important issues in flow control of turbomachinery.Pulsed excitation or suction is widely investigated in compressor cascades.This paper presents a discussion on the unsteady flow control realized by dual sweeping jet actuator(SJA)located on the blade suction surface.The unsteady numerical simulations were utilized to study the effect of applying dual SJAs on controlling the corner separation.With the numerical results,the following conclusions could be drawn with current compressor cascade.A maximum total pressure loss coefficient reduction of 6.8%was obtained.The analysis of the flow field pointed out that the regulation mechanisms of the corner separation were different with each SJA.The SJA ahead achieved an interruption of the suction side boundary layer development and the rear SJA enhanced the interaction and entrainment between the excitation stream and the secondary flows.Meanwhile,the different unsteadiness structures of the flow field frequency spectrum compared with single SJA cases were identified.The first peak frequency corresponded to the difference of the two SJAs and the rest frequencies could be regulated to a base frequency and its harmonic frequencies.

Effect of Non-Axisymmetric End Wall on a Highly Loaded Compressor Cascade in Multi-Conditions

As an effective method to influence end wall flow field,non-axisymmetric profiled end wall can improve the aerodynamic performance of compressor cascades.For a highly loaded low pressure compressor cascade,called V103,the study found the optimal non-axisymmetric profiled end wall decreases total pressure loss coefficient by 4.57%,5.48%and 3.04%under incidences of–3°,0°,and 3°,respectively,compared with those of the planar end wall.The optimal non-axisymmetric profiled end wall changes the structure of secondary flow in hub region,generating a corner vortex near suction surface,inhibiting the development of the passage vortex towards suction surface and reducing flow separation.When the inlet Mach numbers are 0.62 and 0.72,the total pressure loss coefficient decreases by 3.19%and 4.58%for optimal non-axisymmetric profiled end wall compared with those of the planar end wall.Though optimal non-axisymmetric profiled end wall increases total pressure loss near hub region in blade passage under different inlet Mach numbers,the peak value and region of high loss coefficient above 10%span in blade passage significantly decrease.In addition,different incidences affect the secondary flow streamlines and vortex structure near the cascade hub region,however,different inlet Mach numbers hardly change the secondary flow streamlines and vortex structure.In short,the optimal non-axisymmetric profiled end wall shows better aerodynamic performance than the planar end wall for the highly loaded compressor cascade in multi-conditions.

Parameter study on numerical simulation of corner separation in LMFA-NACA65 linear compressor cascade

Large-eddy simulation（LES） is compared with experiment and Reynolds-averaged Navier-Stokes（RANS）, and LES is shown to be superior to RANS in reproducing corner separation in the LMFA-NACA65 linear compressor cascade, in terms of surface limiting streamlines,blade pressure coefficient, total pressure losses and blade suction side boundary layer profiles. However, LES is too expensive to conduct an influencing parameter study of the corner separation.RANS approach, despite over-predicting the corner separation, gives reasonable descriptions of the corner separated flow, and is thus selected to conduct a parametric study in this paper. Two kinds of influencing parameters on corner separation, numerical and physical parameters, are analyzed and discussed： second order spatial scheme is necessary for a RANS simulation; incidence angle and inflow boundary layer thickness are found to show the most significant influences on the corner separation among the parameters studied; unsteady RANS with the imposed inflow unsteadiness（inflow angle varying sinusoidally with fluctuating amplitude of 0.92°） does not show any non-linear effect on the corner separation.

Influence of Endwall Air Injection with Discrete Holes on Corner Separation of a Compressor Cascade

The aim of this study is to reveal the influence mechanism of endwall air injection with distributed holes on the corner separation of a highly loaded compressor cascade,so as to promote the application of injection in aero-engines.Single-hole and double-hole endwall injection schemes featuring different axial locations,pitchwise locations,injection mass rates and injection directions,were designed and investigated.Results showed that the corner separation was eliminated by endwall injection;the optimal single-hole injection scheme achieved an endwall loss coefficient reduction of 29.7%,with injection coefficient as low as 0.48%.The optimal axial location of single-hole endwall injection was at 82%axial chord,being the center of corner separation.However,as injection hole was located at upstream of it,endwall injection resulted in severer corner separation.The mid-span flow field was deteriorated after endwall injection,which was due to 3D flow effects,i.e.,AVDR(axial velocity density ratio)effect and low-momentum fluid spanwise migration effect.The optimal injection was achieved at low injection angle and from close to the suction surface on pitchwise.Double-hole injection exhibited inferior performance compared with single-hole,which was due to the interaction of the two injection streams and mixing of injection streams with the bulk stream.

Numerical investigation of positive dihedral application conditions in compressor cascades

This study attempts to make a contribution to the understanding of dihedral application conditions and their aerodynamic mechanisms.The present efforts have finished contrastive investigations on several dihedral blades to their corresponding straight ones with different geometric or aerodynamic conditions including aspect ratio,solidity,aerofoil turning angle,inlet boundary layer configuration and inlet Mach number.A dihedral with the angle between the suction side and the endwall to be obtuse,i.e.,positive dihedral,is chosen.The result reveals the dihedral application conditions consist of aerofoil turning angle,inlet boundary layer,inlet Mach number and so on.The further analysis indicates:in a transonic cascade,two considerations are needed on the contrastive relationship between intensities of the two shocks,namely detached shock and passage shock,and the interaction of the shocks with the corner separation.

Experimental Investigation of Flow Separation Control Using Dielectric Barrier Discharge Plasma Actuators

Influence of plasma actuators as a flow separation control device was investigated experimentally. Hump model was used to demonstrate the effect of plasma actuators on external flow separation, while for internal flow separation a set of compressor cascade was adopted. In order to investigate the modification of the flow structure by the plasma actuator, the flow field was examined non-intrusively by particle image velocimetry measurements in the hump model experiment and by a hot film probe in the compressor cascade experiment. The results showed that the plasma actuator could be effective in controlling the flow separation both over the hump and in the compressor cascade when the incoming velocity was low. As the incoming velocity increased, the plasma actuator was less effective. It is urgent to enhance the intensity of the plasma actuator for its better application. Methods to increase the intensity of plasma actuator were also studied.

Numerical and experimental investigation of quantitative relationship between secondary flow intensity and inviscid blade force in axial compressors

The secondary flow attracts wide concerns in the aeroengine compressors since it has become one of the major loss sources in modern high-performance compressors.But the research about the quantitative relationship between secondary flow and inviscid blade force needs to be more detailed.In this paper,a database of 889 three-dimensional linear cascades was built.An indicator,called Secondary Flow Intensity(SFI),was used to express the loss caused by secondary flow.The quantitative relationship between the SFI and inviscid blade force deterioration was researched.Blade oil flow and Computation Fluid Dynamics(CFD)results of some cascades were also used to cross-validate.Results suggested that all numerical cascade cases can be divided into 3 clusters by the SFI,which are called Clusters A,B and C in the order of the increasing SFI indicator.The corner stall,known as the strong corner separation,only happens when the SFI is high.Both calculations and oil flow experiments show that the SFI would stay at a low level if the vortex core at the endwall surface does not appear.The strong interaction of Kutta condition and endwall cross-flow is considered the dominant mechanism of higher secondary flow losses,rather than the secondary flow penetration depth on the suction surface.In conclusion,the inviscid blade force spanwise deterioration is strongly related to the SFI.The correlation of the SFI and spanwise inviscid blade force deterioration is given in this paper.The correlation could provide a quantitative reference for estimating secondary flow losses in the design.

Experimental Investigation of a Linear Cascade with Large Solidity Using Pressure Sensitive Paint and Dual-Camera System

Pressure Sensitive Paint(PSP)technique has been increasingly applied to the experimental research of aerodynamics and thermodynamics due to its strengths of non-contact,high resolution results and large coverage area,etc.However,rarely has this technique been successfully used to the study of internal flow such as compressor cascade,since narrow flow passages would heavily restrict the acquisition of PSP images.In this paper,PSP technique was used to study the pressure distribution on a linear compressor cascade with large solidity of 2.3,where the view of recording camera can be heavily blocked due to adjacent blade surfaces.To help get integrated PSP images of the internal flow passage,dual camera system along with image processing tools like 3D reconstruction and image integration were adopted.The results showed that with the aid of such assistance,image results with good quality and readability could be obtained.Meanwhile,pressure data given by PSP were compared with data from traditional way of pressure taps and showed good consistency.Massive results of the entire cascade passage surface were given with different inlet Mach numbers and incidence angles.The results showed that PSP technique can integrally measure cascade tunnel of large solidity with the help of dual-camera system.

Computational Study on the Boundary Layer Suction Effects in the Supersonic Compressor Flow with a Bowed Blade

Flow control methodologies have been widely used to eliminating flow separation and increasing the blade load in axial compressor.Aiming at revealing the flow mechanism of coupled bowed blading and boundary layer suction in a supersonic compressor cascade,a cascade with a diffusion coefficient of 0.62 is numerically presented.First of all,according to the available experimental data,the numerical method was validated;then,different bowed blading effects on flow field in detail were investigated;at last,based on the flow physics of purely bowed blading,the positively bowed blade was coupled with boundary layer suction on blade suction surface,whereas the negatively bowed blade was coupled with endwall suction.For coupled control method,influence mechanism on flow field,especially on the shock structure was revealed,and different aspect ratios of coupled control method were investigated as well.Results showed that the coupled positively bowed blading and suction surface suction can eliminate the flow separation effectively.Compared with that of baseline supersonic cascade,the total pressure loss coefficient of the coupled scheme was reduced by 37.4%at most.At mid-span,the shock moved downstream and the single shock was separated to a dual-shock structure since the positively bowed blading reduced the static pressure of mid-span.The coupled negatively bowed blading and endwall suction also effectively enhanced the performance of cascade by removing the corner separation,with the loss coefficient reduced by as much as 41.9%.However,the suction coefficient of optimal coupled negatively bowed blading scheme reached 10.5%,which is too high for practical use.After coupled control,the 3 D shock structure became“C”shaped distribution along spanwise because of the difference in influence mechanism of negatively bowed blading on different spanwise location.Due to the opposite influence effect of positively and negatively bowed blading,the shock structure in the two different schemes of cascades were different and showed opposite variation trends as aspect ratio increased.