Numerical Investigation of the Unsteady Flow in a Transonic Compressor with Curved Rotors

The unsteady 3D flow fields in a single-stage transonic compressor under designed conditions are simulated numerically to investigate the effects of the curved rotors on the stage performance and the aerodynamic interaction between the blade rows. The results show that, compared to the compressor with unurved rotors, the compressor under scrutiny acquires remarkable increases in efficiency with significantly reduced amplitudes of the time-dependent fluctuation. The amplitude of the pressure fluctuation around the stator leading edge decreases at both endwalls, but increases at the mid-span in the curved rotors. The pressure fluctuation near the stator leading edge, therefore, becomes more uniform in the radial direction of this compressor. Except for the leading edge area, the pressure fluctuatinn amplitude declines remarkably in the tip region of stator surface downstream of the curved rotor, but hardly changes in the middle and at the hub.

Wet Compression Performance of a Transonic Compressor Rotor at its Near Stall Point

In order to study the effects of wet compression on a transonic compressor,a full 3-D steady numerical simulation was carried out under varying conditions.Different injected water flow rates and droplet diameters were considered.The effect of wet compression on the shock,separated flow,pressure ratio,and efficiency was investigated.Additionally,the effect of wet compression on the tip clearance when the compressor runs in the near-stall and stall situations was emphasized.Analysis of the results shows that the range of stable operation is extended,and that the pressure ratio and inlet air flow rate are also increased at the near-stall point.In addition,it seems that there is an optimum size of the droplet diameter.

Numerical Investigation into the Transient Behavior of the Spike-Type Rotating Stall for a Transonic Compressor Rotor

In this paper,a numerical investigation into a spike-type rotating stall process is carried out considering a transonic compressor rotor(the NASA Rotor 37).Through solution of the Unsteady Reynolds-Averaged Navier-Stokes(URANS)equations,the evolution process from an initially circumferentially-symmetric near-stall flow field to a stable stall condition is simulated without adding any artificial disturbance.At the near-stall operating point,periodic fluctuations are present in the overall flow of the rotor.Moreover,the blockage region in the channel periodically shifts from middle span to the tip.This fluctuating condition does not directly lead to stall,while the full-annulus calculation eventually evolves to stall.Interestingly,a kind of“early disturbance”feature appears in the dynamic signals,which propagates forward ahead of the rotor.

Design optimization of transonic compressor stage using CFD and response surface model

In order to shorten the design period, the paper describes a new optimization strategy for computationally expensive design optimization of turbomachinery, combined with design of experiment (DOE), response surface models (RSM), genetic algorithm (GA) and a 3-D Navier-Stokes solver(Numeca Fine). Data points for response evaluations were selected by improved distributed hypercube sampling (IHS) and the 3-D Navier-Stokes analysis was carried out at these sample points. The quadratic response surface model was used to approximate the relationships between the design variables and flow parameters. To maximize the adiabatic efficiency, the genetic algorithm was applied to the response surface model to perform global optimization to achieve the optimum design of NASA Stage 35. An optimum leading edge line was found, which produced a new 3-D rotor blade combined with sweep and lean, and a new stator one with skew. It is concluded that the proposed strategy can provide a reliable method for design optimization of turbomachinery blades at reasonable computing cost.

Non-synchronous vibration of rotor blade in a six-stage transonic compressor

This paper presents an experimental study on the Non-Synchronous Vibration(NSV)in a six-stage transonic compressor.The first part of the paper describes the NSV phenomenon of Rotor 1,which occurs when both Stator 1(S1)and Stator 2(S2)or S1 only are closed.Detailed measurements and analysis are carried out for the former case through the unsteady wall pressure and the Blade Strain(BS).The spinning mode theory used in the rotor/stator interaction noise is employed to explain the relation between the circumferential wave number of the aerodynamic disturbance and the Nodal Diameter(ND)of the blade vibration.The variations of the vibration amplitudes of different blades and the Inter-Blade Phase Angles(IBPAs)at different moments suggest that the evolution of NSV is a highly nonuniform phenomenon along the circumferential direction.In addition,the difference between the wall-pressure spectra generated by the NSV and the classic flutter has been discussed.In the second part,the variations of aerodynamic loading due to the adjustment of the staggers of the Inlet Guide Vane(IGV),S1 and S2 have been investigated.It is found that closing S1 only can result in a great fluctuation to the performance of the front stages,which might be detrimental to the flow organization and increase the risk of NSV.In contrast,the effect of closing S2 only on the performance of the first two stages appears to be slighter relatively.

Numerical Simulation on Unsteady Flow Mechanism of a 1.5-Stage Axial Transonic Compressor

In this paper,a numerical simulation method is used to calculate a 1.5-stage axial transonic compressor to explore its unsteady flow mechanism.The performance curve is compared with the experimental data to verify the calculation method with a high numerical accuracy,which shows that the unsteady calculation has good reliability.According to the analysis of the data from the monitoring points under the near-stall condition,the unsteady disturbances originate from the tip region of blade and perform the strongest at the blade pressure surface with a broadband characteristic.Further analysis is conducted by combining with the characteristics of the transient flow field at the tip of blade.The results show that the unsteady pressure fluctuations are caused by the migration of the new vortex cores.These new vortex cores are generated by the breakdown of leakage vortex in the downstream,which is induced by the leakage vortex and shock wave interference.Moreover,the relationship between the unsteady flow characteristics and the working conditions is also studied.The leakage vortex intensity and the shock wave strength gradually increase with the decrease of flow rate.When the combination of the leakage vortex intensity and shock wave strength reaches the first threshold,a single frequency of unsteady disturbances appears at the blade tip.When the combination of the leakage vortex intensity and shock wave strength reaches the second threshold,the frequency of unsteady disturbances changes to a broadband.

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.

Effects of Number of Bleed Holes on Shock-Wave/Boundary-Layer Interactions in a Transonic Compressor Stator

An extensive numerical investigation is conducted to characterize the flow separation control in a transonic compressor cascade with a porous bleed.The bleed holes are arranged on the suction surface in a single row,two staggered rows and three staggered rows.For each bleed scheme,five bleed pressure ratios are examined at an inlet Mach number of 1.0.The results indicate that the aerodynamic performance of the cascade is significantly improved by the porous bleed.For the single-row scheme,the maximum reduction in total pressure losses is 57%.For the two-staggered-row and three-staggered-row schemes,there is an optimal bleed pressure ratio of 1.0,and the maximum reductions in total pressure loss are 68% and 75%,respectively.The low loss in the cascade is due to the well-controlled boundary layer.The new local supersonic region created by the bleed hole is the key reason for the improved boundary layer.The vortex induced by side bleeding provides another mechanism for delaying flow separation.Increasing the bleed holes could create multiple local supersonic regions,which reduce the range of the adverse pressure gradient that the boundary layer needs to withstand.This is the reason why cascades with more bleed holes perform better.

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.

Numerical Investigation of Circumferential Groove Casing Treatment in a Highly-Loaded Low-Reaction Transonic Compressor Rotor

In this paper, a computational investigation of circumferential groove casing treatment in a highly-loaded low-reaction transonic compressor rotor is conducted, in which the stage reaction is significantly reduced due to a larger meridional contraction with respect to conventional transonic compressors. Steady computation at near-stall point is performed first to capture the stall inception of the rotor with smooth casing. Detailed observations, which mainly focus on the tip leakage flow behavior, obstruction and vortical structures in the tip region, determine the reason for the compressor stall. There is tip leakage vortex breakdown in the tip region. Moreover, it yields passage obstruction, and finally leads to the compressor stall. Then, attempts are made to investigate how the circumferential grooves can be applied for the compressor’s stall margin enhancement without compromising efficiency. Three configurations are obtained and analyzed by changing axial position and the number of the circumferential grooves. The results of computational parametric study indicate the optimal location of the groove is near the leading edge and the downstream grooves combine their influence on the compressor’s stabilization and performance in a cumulative manner. The optimal circumferential groove configuration produces an increase of 1% in total pressure ratio at the near-stall point and a gain of 3.7% in stall margin, without any penalty in efficiency. Furthermore, the impact the grooves will exert on the flow mechanisms between the grooves and the main flow is also considered.

Numerical Research on Inlet Total Pressure Distortion in a Transonic Compressor with Non-Axisymmetric Stator Clearance

Inlet total pressure distortion has great adverse effects on the aero-engine performance. The distorted flow passes through the compressor and becomes non-uniform in the downstream blade rows. Different from previous studies based on the assumption of circumferential uniformity, this study aims to improve circumferential non-uniform flow with the non-axisymmetric structure. Non-axisymmetric stator clearance was adopted to resolve the effects of non-uniform flow caused by inlet total pressure distortion in this paper. The 9 stators with tip clearance were installed in the distorted region and the flow field structure and performance under different operating conditions was studied. The study finds that the non-axisymmetric compressor with 9 tip clearance stators can ensure compressor efficiency while improving compressor stability margin. What’s more, the separation range and strength in the distorted region can be reduced significantly and the anti-distortion capability of compressor can be enhanced.

Aerodynamic Sweeping Study and Design for Transonic Compressor Rotor blades

The objective of the present paper is to study the sweep effect on the blade design performance of a transonic compressor rotor.The baseline to be modified and swept is a designed well efficient transonic single rotor compressor. The first part of the present study is concerning the sweep effect with straight leading edge.In this case fixing the hub section the swept blade is formed by tilting the leading edge with whole blade forwards and backwards axially.The second part is to use an optimization strategy with simple gradient-based optimum-searching method and multi-section blade parameterization technique to search and generate an optimal swept rotor with curved arbitrary leading edge.Its adiabatic efficiency is a little bit greater than that of the reference un-swept rotor.

Transition from Unsteady Flow Inception to Rotating Stall and Surge in a Transonic Compressor

Since the transition from rotating stall to surge in a transonic compressor at high speed is very quick,quite often there is no time to take measures to prevent the surge.Therefore,it is desired to find any rotating stall precursors,of which the occurrence can offer sufficient time for stall or surge prevention.In this study,a series of unsteady flow analyses were performed on a transonic compressor under operating conditions before rotating stall with unsteady results scrutinized to find rotating stall precursors.Particular attention is paid to the spatial modes and time modes of static pressure near the casing and around the blade leading and trailing edges.The results show that the characteristics of the precursor in both spatial and time domains can be used as rotating stall warnings.

A Study of 3D Aerodynamic Design for a Transonic Compressor Blading Optimized by the Locations of Aerofoil Maximum Thickness and Maximum Camber

A design procedure for improving the efficiency of a transonic compressor blading was proposed based on a rapid generation method for three-dimensional blade configuration and computational meshes, a three-dimensional Navier-Stokes solver and an optimization approach. The objective of the present paper is to design a transonic compressor blading optimized only by selection of the locations of maximum camber and maximum thickness for the airfoils at different span heights and to study how do these two design parameters affect the blade performance. The blading configuration and the computational meshes can be obtained very rapidly for any given combination of maximum camber and maximum thickness. The computational grid system generated is used for the Navier-Stokes solution to predict adiabatic efficiency, total pressure ratio and flow rate. As a main result of the optimization, adiabatic efficiency was successfully improved.

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.

Numerical Investigation of the Unsteady Tip Leakage Flow and Rotating Stall Inception in a Transonic Compressor

It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception.Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor.For steady analysis,the predicted results agree well with the experimental data for the estimation of compressor rotor global performance.For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage.On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale-spike type rotating stall inception at blade tip region.It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased,the rotating stall cell was further developed in the blade passage.

Experiments of effects of inlet-air distortion on aerodynamic performance in transonic compressor

The inlet-air distortion which was caused by high angle-of-attack flight was simulated by plugboard.Experiments were conducted on a transonic axial-flow compressor's rotor at 98% rotating speed.The flow-field characteristics and mechanism of performance degradation were analyzed in detail.The compressor inlet was divided into four sectors at circumference under inlet-air distortion.They were undistorted sector,transition sector A where the rotor was rotating into the distortion sector,distorted sector and transition sector B where the rotor was rotating out of the distortion sector.The experimental results show that compared with undistorted sector,there is a subsonic flow in transition sector A,so the pressure ratio is decreased by a large margin in this sector.However,the shock wave is enhanced in distortion sector and transition sector B,and thus the pressure ratio increases in these sectors.Because of the different works at circumference,the phase angle of total pressure changes 90° when the inlet total pressure distortion passes through compressor rotor.In addition,the frequency and amplitude of disturbances in front of the rotor strengthenes under inlet distortion,so the unstable flow would take place in advance.In addition,the position of stall inception is in one of the transition sectors.

Experimental and Numerical Investigation of Non-synchronous Blade Vibration Excitation in a Transonic Axial Compressor

The complex flow phenomenon of rotating instability(RI) and its induced non-synchronous vibration(NSV) have become a significant challenge as they continuously increase aerodynamic load.This study aims to provide an understanding of the non-synchronous blade vibration phenomenon caused by the rotating instability of a transonic axial compressor rotor.In this case,blade vibrations and non-synchronous excitation are captured by strain gauges and unsteady wall pressure transducer sensors.Unsteady numerical simulations for a full-annulus configuration are used to obtain the non-synchronous flow excitation.The results show that the first-stage rotor blade exhibits an NSV close to the first bending mode;NSV is accompanied by a sharp increase in pressure pulsation;amplitude can reach 20%,and unsteady aerodynamic frequency will lock in a structural mode frequency when the blade vibrates in a large-amplitude motion.The predicted NSV frequency aligns well with the experimental results.The dominant mode of circumferential instability flow structure is approximately 47% of the number blades,and the cell size occupies 2-3 pitches in the circumferential direction.The full-annulus unsteady simulations demonstrate that the streamwise oscillation of the shedding and reattachment vortex structure is the main cause of NSV owing to the strong interaction between the tip leakage and separation vortices near the suction surface.

Effects of circumferential configuration concerning splitter blade on the aerodynamic performance in a transonic axial fan

For a certain type of transonic axial fan, the flow field of a fan rotor with splitter blade was computed by numerical simulation, and the shape of the rotor was modified. The effects of different circumferential distributions concerning the splitter cascades upon the aerodynamic performance were investigated. The studies show that the optimum splitter cascade is not very close to the suction side of main blade. The load between the main blade and the splitter blade can be soundly distributed in terms of the adjustment of circumferential position of the splitter blade. The best aerodynamic performance can be successfully obtained according to the optimum shape of the expanding fluid channel reasonably formed by the splitter blade and the main blade.

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.