Near-stall Conditions Research Articles

1205 軸流圧縮機における旋回不安定擾乱の解析(OS12-2 流体機械の諸問題,オーガナイズドセッション)

The rotating instability has been investigated at a near-stall condition in an axial flow compressor rotor with large tip clearance by EFD and CFD. This flow phenomenon has disturbances that move relative to the rotor blades and cause high amplitude rotor blade vibrations and tip clearance noise. Furthermore, it has closely related to rotating stall inception. In the present study, it has been found that rotating instability disturbances propagate in cyclic process of strong breakdown and weak breakdown of the tip leakage vortex.

1711 遷音速軸流圧縮機動翼列の失速点近傍における翼先端漏れ流れの非定常挙動に関する数値解析(S19-2 流体機械に関連した流動現象とその解析・計測(2)不安定流動・非定常流動の計測と解析,21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)

1711 遷音速軸流圧縮機動翼列の失速点近傍における翼先端漏れ流れの非定常挙動に関する数値解析(S19-2 流体機械に関連した流動現象とその解析・計測(2)不安定流動・非定常流動の計測と解析,21世紀地球環境革命の機械工学:人・マイクロナノ・エネルギー・環境)

Turbulence Models Assessment for Large-Scale Tip Vortices in an Axial Compressor Rotor

Six turbulence models frequently used in compressor aerodynamics were employed in the detailed numerical investigations of a low-speed large-scale axial compressor rotor, for which the tip flows were measured in detail with stereoscopic particle image velocimetry, to assess the predictive capabilities of the turbulence models for large-scale vortices in the tip region of the rotor. The six turbulence models include: the mixing-length model, the Spalart-Allmaras model, the standard k - e model, the shear-stress transport k - w model, the v 2 - f model, and the Reynolds stress model. Their results were carefully discussed and compared with the measurements both on velocity fields and turbulence stresses. It was found that the Reynolds stress model is superior to the others in the prediction of the tip-leakage vortex at the design condition, whereas the standard k - e model shows the best results in the prediction of the corner vortex at the near-stall condition. Although the simulation could predict the large-scale tip vortices well in the mean flowfield, the computed flow mechanism has large discrepancy with the reality.

Multi-Objective Aerodynamic Exploration of Elements' Setting for High-Lift Airfoil Using Kriging Model

A multi-objective design exploration for a three-element airfoil consisting of a slat, a main wing, and a flap was carried out. The lift curve improvement is important to design high-lift system, thus design has to be performed with considered multi-angle. The objective functions considered here are to maximize the lift coefficient at landing and near-stall conditions simultaneously. Genetic algorithm is used as an optimizer. Although it has the advantage of global exploration, its computational cost is expensive. To reduce the computational cost, the Kriging surrogate model, which was constructed based on several sample designs, is introduced. The solution space was explored based on the maximization of expected improvement (EI) value corresponding to objective functions on the Kriging surrogate models. The improvement of Kriging surrogate model and the exploration of the optimum can be advanced at the same time by maximizing EI value. In this study, a total of 90 sample points are evaluated using the Reynolds-averaged Navier-Stokes simulation for the construction of the Kriging surrogate model. Through the present exploration process, several designs were obtained with better performance than the baseline setting in each objective function. To obtain the information of the design space, functional analysis of variance, which is one of the data mining techniques showing the effect of each design variable on the objectives, is applied. Main effects of the design variables are calculated to recognize which design variable has the effect on the objective functions. This result suggests that the gap and the deflection of the flap have a remarkable effect on each objective function and the gap of the slat has an effect at near-stall condition.

Characteristics of the Tip Leakage Vortex in a Low-Speed Axial Compressor

The flowfields in the tip region of the rotor passage are investigated from a large-scale low-speed axial compressor test at both the design and near-stall conditions by using stereoscopic particle image velocimetry. The characteristics of the tip leakage vortex (TLV) are analyzed based on the variations of its trajectory, concentration, size, streamwise velocity, and the blockage effect, which are extracted from both the ensemble- and conditional-averaged results. According to the ensemble-averaged performance of the TLV, its evolution procedures can be divided into several distinct flow phases. As the flow phase of the TLV changes, its mean flow characteristics also change accordingly. The TLV is inherently unsteady and features unsteady behaviors of vortex wandering, splitting, and breakdown, which have profound effects to its mean flow characteristics.

Stereoscopic PIV measurement of unsteady flows in an axial compressor stage

Stereoscopic particle-image velocimetry (SPIV) measurements are conducted in a Low-speed Large-scale Axial Compressor. During the experiment the two CCD cameras are placed at the different sides of the laser light sheet and it is proved that this configuration is more suitable for the investigation in multi-stage turbomachines. The measured results, including the overall performances of many typical flow structures near the rotor tip region and the phase locked unsteady flows inside the stator passage at both the design and near-stall conditions, are introduced. Some new features of the complicated flow structures, such as the breakdown of the tip leakage vortex, the formation of the compound corner vortex at the rotor suction tip corner, the interactions between the hub stall and the tip separation and the rotor wakes, and the evolutions of the tip corner anti-rotating streamwise vortices inside the stator passage, are revealed.

1628 軸流圧縮機動翼列の失速点近傍における非定常流動現象の数値解析(J12-1 流体機械に関連した流体解析と数値最適化技術(1),J12 流体機械に関連した流体解析と数値最適化技術)

The tip clearance of rotor blades is one of the important factors to be more investigated, because the rotating stall inception originates at a tip leakage flow region in most of axial flow compressors. However, the effect of the tip clearance on unsteady flow field at near-stall condition has not been investigated in detail. In the present work, the unsteady flow phenomena in a low-speed axial flow compressor stage with the large and small tip clearances at near-stall condition have been investigated numerically by URANS (Unsteady Reynolds-averaged Navier-Stokes) simulations. The results show the clear difference caused by the effect of the tip clearance on the flow field at near-stall condition. Namely, the occurrence of separation at the blade suction surface induced instability for small tip clearance, while the occurrence of tip leakage vortex breakdown induced instability for large tip clearance.

G403 軸流圧縮機動翼列の失速点近傍における非定常流れ挙動(1)(GS4 流体構造,界面,一般セッション)

G403 軸流圧縮機動翼列の失速点近傍における非定常流れ挙動(1)(GS4 流体構造,界面,一般セッション)

G403 軸流圧縮機動翼列の失速点近傍における非定常流れ挙動(2)(GS4 流体構造,界面,一般セッション)

G403 軸流圧縮機動翼列の失速点近傍における非定常流れ挙動(2)(GS4 流体構造,界面,一般セッション)

Application of SPIV in turbomachinery

Stereoscopic particle-image velocimetry (SPIV) has been successfully used in a low-speed large-scale axial compressor. A configuration in which two CCD cameras were placed at different sides of the light sheet was employed. It is demonstrated that the results measured with such a configuration are significant for the study of unsteady flow structures of the streamwise vortices and secondary flows in the test rotor, and that such a configuration is easy to use in multi-stage turbomachinery. The instantaneous snapshots, ensemble-averaged results and turbulence statistics in the rotor passage were obtained at both the design and near-stall conditions. The representative flow structures, such as the tip leakage vortex, the corner vortex and the inlet guide vane wake, can be depicted clearly. Moreover, according to experimental and theoretical analyses, some guidance is provided for the application of SPIV in turbomachinery.

Unsteady Navier-Stokes Simulation of a Transonic Flutter Cascade Near-Stall Conditions Applying Algebraic Transition Models

Due to the trend in the design of modern aeroengines to reduce weight and to realize high pressure ratios, fan and first-stage compressor blades are highly susceptible to flutter. At operating points with transonic flow velocities and high incidences, stall flutter might occur involving strong shock-boundary layer interactions, flow separation, and oscillating shocks. In this paper, results of unsteady Navier-Stokes flow calculations around an oscillating blade in a linear transonic compressor cascade at different operating points including near-stall conditions are presented. The nonlinear unsteady Reynolds-averaged Navier-Stokes equations are solved time accurately using implicit time integration. Different low-Reynolds-number turbulence models are used for closure. Furthermore, empirical algebraic transition models are applied to enhance the accuracy of prediction. Computations are performed two dimensionally as well as three dimensionally. It is shown that, for the steady calculations, the prediction of the boundary layer development and the blade loading can be substantially improved compared with fully turbulent computations when algebraic transition models are applied. Furthermore, it is shown that the prediction of the aerodynamic damping in the case of oscillating blades at near-stall conditions can be dependent on the applied transition models.

Experimental Analysis on Tip Leakage and Wake Flow in an Axial Flow Fan According to Flow Rates

The flow characteristics in the blade passage and in the wake region of a low-speed axial flow fan have been investigated by experimental analysis using a rotating hot-wire sensor and a five-hole probe for design and off-design operating conditions. The results show that the tip leakage vortex is moved upstream when the flow rate is decreased, thus disturbing the formation of wake flow near the rotor tip. That is, the tip leakage vortex interfaces with the blade suction surface and results in high velocity fluctuation near the blade suction surface. From axial velocity distributions downstream of the fan rotor, large axial velocity decay near the rotor tip is observed at near-stall condition, which results in a large blockage compared to that at the design condition. Finally, the wake flow downstream of the rotor blade is clearly measured at the design and off-design conditions. However, the trough of the high velocity fluctuation due to Karmann vortex street in the wake flow is observed at a higher flow condition than the design flow rate.

DPIV study of near-stall wake-rotor interactions in a transonic compressor

Wake-rotor interactions at near-stall conditions are studied using digital particle image velocimetry (DPIV) in a transonic compressor at the Compressor Aero Research Lab (CARL) at Wright-Patterson Air Force Base. The wake generators (WGs) are designed to simulate wakes of highly loaded stators without producing flow swirl. Flow visualization, instantaneous, and average velocity field data are presented for near-stall compressor operating conditions. These results describe the interaction of the wake shed from the upstream WG with the downstream potential field of the rotor leading edge for two axial spacings at 75 and 90% span. The advantage of ensemble averaging is demonstrated when operating at challenging conditions such as near-stall, which poses special difficulties in obtaining DPIV data. Ensemble averaging, which is possible because of a natural phase locking provided by the blade potential field, allows robust statistics.

The Application of Pressure- and Temperature-Sensitive Paints to an Advanced Compressor

An innovative pressure-measurement technique that employs the tools of molecular spectroscopy has been widely investigated by the aerospace community. Measurements are made via oxygen-sensitive molecules attached to the surface of interest as a coating, or paint. The pressure-sensitive-paint (PSP) technique is now commonly used in stationary wind-tunnel tests; this paper presents the use of this technique in advanced turbomachinery applications. New pressure- and temperature-sensitive paints (P/TSPs) have been developed for application to a state-of-the-art transonic compressor where pressures up to 1.4 atm and surface temperatures to 90°C are expected for the suction surface of the first-stage rotor. PSP and TSP data images have been acquired from the suction surface of the first-stage rotor at 85 percent of the corrected design speed for the compressor near-stall condition. A comparison of experimental results with CFD calculations is discussed.

Anomalous Flow Phenomena Due to Breakdown of Tip Leakage Vortex in an Axial Compressor Rotor at Near-Stall Condition.

Anomalous flow phenomena in an axial compressor rotor at near-stall conditions have been investigated by unsteady three-dimensional Navier-Stokes flow simulations. The simulation show that the spiral-type breakdown of the tip leakage vortex occurs inside the rotor passage at the near-stall conditions. Downstream of the breakdown onset, the tip leakage vortex twists ant turns violently with time, thus interacting with the pressure surface of the adjacent blade. The vortex breakdown causes significant changes in the nature of the tip leakage vortex, which result in the anomalous phenomena in the tip leakage flow field at the near-stall conditions : no rolling-up of the leakage vortex downstream of the rotor, disappearance of the casing wall pressure trough corresponding to the leakage vortex, large spread of the low-energy fluid accumulating on the pressure side, and large pressure fluctuation on the blade pressure surface. As the flow rate is decreased from the near-stall conditions, the movement of the tip leakage vortex due to the vortex breakdown becomes so large that the leakage vortex interacts with the suction surface as well as the pressure one. The interaction gives rise to the three-dimensional separation of the suction surface boundary layer.

The Role of Tip Leakage Vortex Breakdown in Compressor Rotor Aerodynamics

The breakdown of tip leakage vortex has been investigated on a low-speed axial compressor rotor with moderate blade loading. Effects of the breakdown on the rotor aerodynamics are elucidated by Navier–Stokes flow simulations and visualization techniques for identifying the breakdown. The simulations show that the leakage vortex breakdown occurs inside the rotor at a lower flow rate than the peak pressure rise operating condition. The breakdown is characterized by the existence of the stagnation point followed by a bubblelike recirculation region. The onset of breakdown causes significant changes in the nature of the tip leakage vortex: large expansion of the vortex and disappearance of the streamwise vorticity concentrated in the vortex. The expansion has an extremely large blockage effect extending upstream of the leading edge. The disappearance of the concentrated vorticity results in no rolling-up of the vortex downstream of the rotor and the disappearance of the pressure trough on the casing. The leakage flow field downstream of the rotor is dominated by the outward radial flow, resulting from the contraction of the bubblelike structure of the breakdown region. It is found that the leakage vortex breakdown plays a major role in characteristic of rotor performance at near-stall conditions. As the flow rate is decreased from the peak pressure rise operating condition, the breakdown region grows rapidly in the streamwise, spanwise, and pitchwise directions. The growth of the breakdown causes the blockage and the loss to increase drastically. Then, the interaction of the breakdown region with the blade suction surface gives rise to the three-dimensional separation of the suction surface boundary layer, thus leading to a sudden drop in the total pressure rise across the rotor.

Computations of bladerow stall inception in transonic flows

AbstractA three-dimensional unsteady Navier-Stokes solver has been used to simulate stall inception in a single row ten passage segment of a transonic fan, the NASA rotor-67. At subsonic flow conditions, the 3D results illustrate a rotating stall inception with short scale part-span cells rotating at around 80% rotor speed, similar to that observed in some low speed experiments. However, at a supersonic relative inflow condition, the results show that an isolated blade row tends to stall in a one-dimensional breakdown pattern without first experiencing rotating stall. At near-stall conditions, significant self-excited unsteadiness is generated by the interaction between the tip-leakage vortex and the passage shock wave. Further computations for two-dimensional configurations indicate that it is possible to have a rotating pattern of instability in transonic blade rows associated with circumferential synchronised shock oscillation.

Variations in Upstream Vane Loading With Changes in Back Pressure in a Transonic Compressor

Dynamic loading of an inlet guide vane (IGV) in a transonic compressor is characterized by unsteady IGV surface pressures. These pressure data were acquired for two spanwise locations at a 105 percent speed operating condition, which produces supersonic relative Mach numbers over the majority of the rotor blade span. The back pressure of the compressor was varied to determine the effects from such changes. Strong bow shock interaction was evident in both experimental and computational results. Variations in the back pressure have significant influence on the magnitude and phase of the upstream pressure fluctuations. The largest unsteady surface pressure magnitude, 40 kPa, was obtained for the near-stall mass flow condition at 75 percent span and 95 percent chord. Radial variation effects caused by the spanwise variation in relative Mach number were measured. Comparisons to a two-dimensional nonlinear unsteady blade/vane Navier–Stokes analysis show good agreement for the 50 percent span results in terms of IGV unsteady surface pressure. The results of the study indicate that significant nonlinear bow shock influences exist on the IGV trailing edge due to the downstream rotor shock system. [S0889-504X(00)00303-2]

A Non-Linear Aeroelasticity Analysis of a Fan Blade Using Unstructured Dynamic Meshes

The main objective of this paper is to present a methodology for the three-dimensional aeroelasticity analysis of turbomachinery blades using an unstructured compressible Navier-Stokes solver for the fluid and a modal model for the structure. The basic fluid solver is constructed in the form of a central difference scheme with explicitly added artificial dissipation which is based upon the fourth- and second-order differences of the solution. The temporal discretization uses an implicit time integration scheme based on a Jacobi relaxation procedure. The structural modes of vibration are determined via a finite element model and the mode shapes are interpolated on to the fluid mesh in a manner that is consistent with general unstructured tetrahedral grids. A spring analogy algorithm that can move the mesh according to the instantaneous shape of a deforming blade has been developed for the accurate tracking of the solid boundaries without creating excessive grid distortions. The performance of the resulting system was examined by considering the aeroelastic behaviour of NASA Rotor 67 fan blade and predictions were compared to experimental results wherever possible. Using a three-dimensional cyclic symmetry model, the tip leading edge time histories were predicted under peak-efficiency and near-stall conditions, and the corresponding aeroelastic natural frequencies and aerodynamic damping values were determined. The blade was found to be stable in all cases considered.

Viscous Analysis of Three-Dimensional Rotor Flow Using a Multigrid Method

A three-dimensional code for rotating blade-row flow analysis has been developed. The space discretization uses a cell-centered scheme with eigenvalue scaling for the artificial dissipation. The computational efficiency of a four-stage Runge–Kutta scheme is enhanced by using variable coefficients, implicit residual smoothing, and a full-multigrid method. An application is presented for the NASA rotor 67 transonic fan. Due to the blade stagger and twist, a zonal, nonperiodic H-type grid is used to minimize the mesh skewness. The calculation is validated by comparing it with experiments in the range from the maximum flow rate to a near-stall condition. A detailed study of the flow structure near peak efficiency and near stall is presented by means of pressure distribution and particle traces inside boundary layers.