Performance Impact of Tip Clearance Variation on a Transonic, Low Aspect Ratio, Axial Compressor Stage

2014 ◽  
Author(s):  
Rebecca Howard ◽  
Steven L. Puterbaugh
Keyword(s):  
Aspect Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Compressor Stage ◽  
Performance Impact ◽  
Low Aspect Ratio

Aerodynamic Effects of an Unshrouded Low Pressure Turbine on a Low Aspect Ratio Exit Guide Vane

2012 ◽  
Author(s):  
Thorsten Selic ◽  
Davide Lengani ◽  
Andreas Marn ◽  
Franz Heitmeir
Keyword(s):  
Aspect Ratio ◽  
Guide Vane ◽  
Low Pressure ◽  
Tip Clearance ◽  
Pressure Probe ◽  
Model Parameters ◽  
Leakage Flow ◽  
Low Pressure Turbine ◽  
Low Aspect Ratio ◽  
The Impact

This paper presents the effects of an unshrouded low pressure turbine (LPT) onto the following exit guide vane row (EGV). The measurement results were obtained in the subsonic test turbine facility at Graz University of Technology by means of a fast response pressure probe in planes downstream of the rotor as well as oil flow visualisation. The test rig was designed in cooperation with MTU Aero Engines and represents the last 1.5 stages of a commercial aero engine. Considerable efforts were put into the adjustment of all relevant model parameters to reproduce the full scale LPT situation. Different tip clearances were evaluated by means of CFD obtained using a commercial Navier-Stokes code and validated with experimental results. The goal is to evaluate the effect of the varying leakage flow on the flow in the low aspect ratio EGV. Special attention is given to the impact on the development of secondary flows as well as the flow structures downstream of the EGV. The effect of the leakage flow causes a change of the flow structure of the EGV, particularly losses. Considering the largest investigated tip-clearance, the losses increased by 71% when compared to a zero-leakage case.

Influence of Blade Sweep on the Energetic Behavior of Axial Flow Turbomachinery Rotors at Design Flow Rate

2004 ◽  
Author(s):  
Ja´nos Vad ◽  
Ali R. A. Kwedikha ◽  
Helmut Jaberg
Keyword(s):  
Aspect Ratio ◽  
Flow Rate ◽  
Axial Flow ◽  
Pressure Rise ◽  
Design Flow ◽  
Tip Clearance ◽  
Total Efficiency ◽  
Low Aspect Ratio ◽  
Axial Velocity Profile ◽  
Bladed Rotor

Experimental and computational studies were carried out in order to survey the energetic aspects of forward and backward sweep in axial flow rotors of low aspect ratio blading for incompressible flow. It has been pointed out that negative sweep tends to increase the lift, the flow rate and the ideal total pressure rise in the vicinity of the endwalls. Just the opposite tendency was experienced for positive sweep. The local losses were found to develop according to combined effects of sweep near the endwalls, endwall and tip clearance losses, and profile drag influenced by re-arrangement of the axial velocity profile. The forward-swept bladed rotor showed reduced total efficiency compared to the unswept and swept-back bladed rotors. This behavior has been explained on the basis of analysis of flow details. It has been found that the swept bladings of low aspect ratio tend to retain the performance of the unswept datum rotor even in absence of sweep correction.

Stall Inception in Axial Compressors

1990 ◽  
Vol 112 (1) ◽  
pp. 116-123 ◽  
Author(s):  
N. M. McDougall ◽  
N. A. Cumpsty ◽  
T. P. Hynes
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Compressor Stage ◽  
Inlet Flow ◽  
Stall Inception ◽  
Axial Compressors ◽  
Multistage Compressor

Detailed measurements have been made of the transient stalling process in an axial compressor stage. The stage is of high hub-casing ratio and stall is initiated in the rotor. If the rotor tip clearance is small stall inception occurs at the hub, but at clearances typical for a multistage compressor the inception is at the tip. The crucial quantity in both cases is the blockage caused by the endwall boundary layer. Prior to stall, disturbances rotate around the inlet flow in sympathy with rotating variations in the endwall blockage; these can persist for some time prior to stall, rising and falling in amplitude before the final increase, which occurs as the compressor stalls.

Effect of Tip Clearance on Stall Evolution Process in a Low-Speed Axial Compressor Stage

2004 ◽  
Author(s):  
Masahiro Inoue ◽  
Motoo Kuroumaru ◽  
Shinichi Yoshida ◽  
Takahiro Minami ◽  
Kazutoyo Yamada ◽  
...  
Keyword(s):  
Spectral Power ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Rotating Stall ◽  
Short Length ◽  
Tip Clearance ◽  
Evolution Process ◽  
Length Scale ◽  
Compressor Stage ◽  
Low Speed

Effect of the tip clearance on the transient process of rotating stall evolution has been studied experimentally in a low-speed axial compressor stage with various stator-rotor gaps. In the previous authors’ experiments for the small tip clearance, the stall evolution process of the rotor was sensitive to the gaps between the blade rows. For the large tip clearance, however, little difference is observed in the evolution processes independently of the blade row gap. In the first half process, it is characterized by gradual reduction of overall pressure-rise with flow rate decreasing, and the number of short length-scale disturbances is increasing with their amplitude increasing. In the latter half a long length-scale disturbance develops rapidly to result in deep stall. Just before the stall inception the spectral power density of the casing wall pressure reveals the existence of rotating disturbances with broadband high frequency near a quarter of the blade passing frequency. This is caused by the short length-scale disturbances occurring intermittently. A flow model is presented to explain mechanisms of the rotating short length-scale disturbance, which includes a tornado-like separation vortex and tip-leakage vortex breakdown. The model is supported by a result of a numerical unsteady flow simulation.

Performance Prediction for High Turning Low Aspect Ratio Stator Cascades in the Transonic Regime

1970 ◽  
Vol 92 (4) ◽  
pp. 390-398
Author(s):  
H. F. L. Griepentrog
Keyword(s):  
Boundary Layer ◽  
Aspect Ratio ◽  
Flow Field ◽  
Performance Prediction ◽  
Transonic Flow ◽  
Secondary Flows ◽  
Compressor Stage ◽  
Low Aspect Ratio ◽  
Aspect Ratios ◽  
Boundary Layer Interaction

This paper describes a method for the prediction of the transonic flow field in a high solidity, high turning cascade, suitable for use as stator of a shock-in-rotor supersonic compressor stage. Effects of shock boundary layer interaction is taken into account by empirical correlation, valid for blade aspect ratios below unity. Use of partial slots for reduction of the secondary flows is briefly discussed and a correlation on slot efficiency is presented.

Experimental and Numerical Study of Stall Flutter in a Transonic Low-Aspect Ratio Fan Blisk

2003 ◽  
Author(s):  
A. J. Sanders ◽  
K. K. Hassan ◽  
D. C. Rabe
Keyword(s):  
Aspect Ratio ◽  
Numerical Study ◽  
Fluid Dynamic ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Strain Gages ◽  
Pressure Transducers ◽  
Low Aspect Ratio ◽  
Benchmark Data

Experiments are performed on a modern design transonic shroudless low-aspect ratio fan blisk that experienced both subsonic/transonic and supersonic stall-side flutter. High-response flush mounted miniature pressure transducers are utilized to measure the unsteady aerodynamic loading distribution in the tip region of the fan for both flutter regimes, with strain gages utilized to measure the vibratory response at incipient and deep flutter operating conditions. Numerical simulations are performed and compared with the benchmark data using an unsteady three-dimensional nonlinear viscous computational fluid dynamic (CFD) analysis, with the effects of tip clearance, vibration amplitude, and the number of time steps-per-cycle investigated. The benchmark data are used to guide the validation of the code and establish best practices that ensure accurate flutter predictions.

Aerodynamic Performance of Rear Axial Compressor Stage With Annular Diffuser and Outlet Chamber

1996 ◽  
Author(s):  
Václav Cyrus
Keyword(s):  
Aspect Ratio ◽  
Velocity Profile ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Compressor Stage ◽  
Guide Vanes ◽  
Inlet Velocity ◽  
Aspect Ratios ◽  
Annular Diffuser ◽  
Outlet Chamber

A detailed investigation of aerodynamic performance of three low-speed rear axial compressor stage bladings with the aspect ratios: 0.75, 1.0 and 1.25 was carried out. The bladings of industrial type consist of rotor, stator and outlet guide vanes. The outer and inner diameters of the stage are constant. The hub/tip ratio is 0.871 and the outer diameter is 800 mm. Stage blading is followed by an annular diffuser with outlet chamber. The effect of blade aspect ratio on compressor stage performance was also analysed with the use of straight cascade data. This data supported the test stage experimental results. We found that the effect of aspect ratio on stage performance is not remarkable in the considered range. There are some differences at off-design conditions. The lowest value of blading efficiency was obtained in the case with the lowest aspect ratio value. Three inlet velocity profiles were modelled with the use of lengthened inlet annulus and a screen specially designed. It was found that there is a significant effect of inlet velocity profile distortion on rear compressor stage blading performance for all aspect ratios. Aerodynamic characteristics of compressor stage blading with annular diffuser and outlet chamber were determined. During the investigation we also removed the outlet guide vanes. Therefore the effect of swirl and inlet velocity profile could be investigated.

Sign in / Sign up

Export Citation Format

Share Document