Investigation on stall inception of axial compressor under inlet rotating distortion

2015 ◽  
Vol 231 (10) ◽  
pp. 1859-1870 ◽  
Author(s):  
Mingming Zhang ◽  
Anping Hou
Keyword(s):  
Flow Field ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Numerical Approach ◽  
Flow Coefficient ◽  
Stall Inception ◽  
Stall Margin ◽  
Inlet Distortion ◽  
Flow Through ◽  
Inflow Conditions

This paper applies a numerical approach to improve the understanding of reaction to various inflow conditions for the compressor system and the mechanism of stall inception under rotating inflow distortions. Full annulus, unsteady, three-dimensional computational fluid dynamics has been used to simulate an axial low-speed compressor operating under rotating distorted inflow conditions. The development of the flow through the rotor is then explained in terms of the redistribution of the flow field and the process of stall inception. The results suggest that the increased flow incidence close to the tip region under co-rotating inflow distortion plays an important role on the stall inception process. The presence of a strong modal wave is observed under co-rotating inflow distortions. This leads to a significant impact on the loss of stall margin, as compared with other distorted inflow conditions. There is a significant peak in the flow coefficient at stall for co-rotating inlet distortion. It can be interpreted as a resonant behavior of the compressor under a strong interaction between the flow field and inlet distortion. It indicates that the stall inception is triggered by the perturbation of the rotating distorted inflow through the long length scale disturbances.

scholarly journals Optimization Design of a 2.5 Stage Highly Loaded Axial Compressor with a Bezier Surface Modeling Method

2020 ◽  
Vol 10 (11) ◽  
pp. 3860
Author(s):  
Song Huang ◽  
Jinxin Cheng ◽  
Chengwu Yang ◽  
Chuangxin Zhou ◽  
Shengfeng Zhao ◽  
...  
Keyword(s):  
Flow Field ◽  
Dimensional Space ◽  
Three Dimensional ◽  
Axial Compressor ◽  
High Dimensions ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Second Stage ◽  
Highly Loaded ◽  
Three Dimensional Space

Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.

Impact of Sweep on Part Speed Performance of an Axial Compressor Rotor With Circumferential Casing Grooves

2019 ◽  
Author(s):  
Shraman Goswami ◽  
M. Govardhan
Keyword(s):  
Flow Field ◽  
High Performance ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Field Investigation ◽  
Performance Comparison ◽  
Rotor Blades ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Design Speed

Abstract High performance and increased operating range of an axial compressor is obtained by employing three-dimensional design features, such as sweep, as well as shroud casing treatments, such as circumferential casing grooves. A number of different rotor blades with different amounts of sweeps and different sweep starting spans are studied at design speed. Different swept rotors, including zero sweep, are derived from Rotor37 rotor geometry. In the current study the best performing rotor with sweep is analyzed at part speed. The analyses were done for baseline rotor, devoid of any sweep, and with and without circumferential casing grooves. A detailed flow field investigation and performance comparison is presented to understand the changes in flow field at part speed. It is found that that at 100% design speed, stall margin improvement is achived by both sweep and casing grooves, but at 90% speed improvement in stall margin due to sacing groove is very minimal over and above the gain due to sweep. It is also noticed that due to reduced shock loss efficiency is higher at 90% speed than at 100% speed.

Experimental Investigation of the Influence of Inlet Distortion on the Stall Inception in a Low Speed Axial Compressor

2009 ◽  
Author(s):  
Huabing Jiang ◽  
Yajun Lu ◽  
Wei Yuan ◽  
Qiushi Li
Keyword(s):  
Flow Field ◽  
Dynamic Response ◽  
Flow Separation ◽  
Flow Condition ◽  
Axial Compressor ◽  
Flow Conditions ◽  
Inlet Flow ◽  
Stall Inception ◽  
Inlet Distortion ◽  
Compressor Performance

Inlet distortion is one of the major concerns for high maneuverability airplanes. An experiment is performed to investigate the influence of inlet distortion on the stall inception in a low speed axial compressor, where the distorted inlet flow field is simulated with a flat baffle placed upstream of the compressor. The flow field around a rotor blade is measured using 2D Digital Particle Image Velocimetry (DPIV) under both uniform and distorted inlet flow conditions. A comparison of flow fields reveals that the distorted inlet flow condition makes the compressor flow fields asymmetric. Flow separation and blockage within Distorted Sector A and Transition Sector B are more severe compared to Sector C. The distorted Sector A and Transition Sector B are the key regions that degrade compressor performance and stability. The large axial velocity makes the flow field within the Undistorted Sector C vigorous, which helps to suppress flow separation and blockage. Compressor rotor blades experience loading and unloading in a revolution period and the compressor exhibits strong dynamic response when it operates under distorted inlet flow conditions. Time-related pressure signals acquired at the rotor exit are utilized to analyze the development of the stall disturbance and the stall characteristic of the compressor. The development period of the stall disturbance is prolonged by the dynamic response of the compressor flow field under the distorted inlet flow condition. Dynamic development of the stall disturbance induced by inlet distortion reduces the compressor stall intensity. The frequency associated with the rotating stall cell is related to the rotating velocity of stall cells, which keeps the same value for uniform and distorted inlet flow conditions. Consequently, the stall inception of the compressor is influenced by the distorted inlet flow condition. The disturbance initiated in Distorted Sector A will experience development and damping when it propagates circumferentially, and might fail to survive the damping within Undistorted Sector C. Stall inception occurs only when the damping within Undistorted Sector C is insufficient to prevent its growth. The dynamic development of the disturbance can reasonably explain the influence of inlet distortion on compressor performance.

Nonaxisymmetric Flow through Annular Actuator Disks: Inlet Distortion Problem

1978 ◽  
Vol 100 (4) ◽  
pp. 604-617 ◽  
Author(s):  
W. R. Hawthorne ◽  
N. A. Mitchell ◽  
J. E. McCune ◽  
C. S. Tan
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Recent Experimental Data ◽  
Recent Theory ◽  
Inlet Distortion ◽  
Actuator Disk ◽  
Beltrami Flow ◽  
Compressor Rotor ◽  
Flow Through ◽  
Absolute Coordinates

The passage of distorted flow through an annular axial compressor rotor or stator is analyzed in the actuator disk limit. In such a description the flow is steady in absolute coordinates. The resulting analysis yields an overall description of the blade row performance in the presence of inlet flow defects. The present (actuator-disk) analysis is compared successfully with Dunham’s earlier three-dimensional analysis as well as with recent experimental data. In addition, comparison with a recent theory by Greitzer, employing an alternative approach, is made. The analysis shows that at least two important distinct types of vorticity arise, the one being directly analogous to the trailing vorticity (shed circulation) of classical wing theory, and the second being analogous to the vorticity occurring in “secondary flow” theory. The latter arises directly as a result of inlet distortion. The resulting varying flow angles produce spanwise variation in blade loading, and consequent trailing vorticity (Beltrami Flow). The two types of vorticity are therefore interrelated. The static pressure field is also affected by this coupling in agreement with experiment. This problem provided a striking example for which three-dimensional and two-dimensional analyses disagree qualitatively. The vorticity as viewed in coordinates fixed in the rotor is discussed in the Addendum.

Effect of hub clearance size and shape of a cantilevered stator on the performance of a small-scale transonic axial compressor

2021 ◽  
pp. 095765092110606
Author(s):  
Botao Zhang ◽  
Bo Liu ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Hejian Wang
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
Small Scale ◽  
Leakage Flow ◽  
Flow Area ◽  
Stall Inception ◽  
Stall Margin ◽  
Flow Capacity ◽  
Flow Loss ◽  
Transonic Axial Compressor

To deeply understand the hub leakage flow and its influence on the aerodynamic performance and flow behaviors of a small-scale transonic axial compressor, variations of the performance and the flow field of the compressor with different hub clearance sizes and clearance shapes were numerically analyzed. The results indicate that the hub clearance size has remarkable impacts on the overall performance of the compressor. With the increase of the hub clearance, the intensity of the hub leakage flow increases, resulting in more intense flow blockage near the stator hub, which reduces the compressor efficiency. However, the flow field near the blade mid-span is modified due to the more convergent flow as the reduced effective flow area caused by the passage blockage, and the flow separation range is narrowed, thus the flow stability of the compressor is enhanced. On this basis, two kinds of non-uniform clearance cases of expanding clearance and shrinking clearance with the same circumferential leakage area as the design clearance were investigated. The occurrence position of the double leakage flow which is closely connected with the flow loss and blockage is shifted backward by the expanding clearance, the flow capacity near the stator hub is enhanced, and the unsteady fluctuation intensity of the flow field is attenuated but fluctuation frequency remains. Similarly, the modification of the stator blade root flow field may result in the reduction of stall margin. The effect of the shrinking clearance on compressor performance is opposite to that of the expanding clearance, which reduces the peak efficiency and delays the stall inception.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Influence of a Honeycomb Facing on the Flow Through a Stepped Labyrinth Seal

2000 ◽  
Vol 124 (1) ◽  
pp. 140-146 ◽  
Author(s):  
V. Schramm ◽  
K. Willenborg ◽  
S. Kim ◽  
S. Wittig
Keyword(s):  
Flow Field ◽  
Three Dimensional ◽  
Theoretical Work ◽  
Labyrinth Seal ◽  
Honeycomb Structure ◽  
Experimental Investigations ◽  
Labyrinth Seals ◽  
Numerical Predictions ◽  
Flow Through ◽  
Aero Engines

This paper reports numerical predictions and measurements of the flow field in a stepped labyrinth seal. The theoretical work and the experimental investigations were successfully combined to gain a comprehensive understanding of the flow patterns existing in such elements. In order to identify the influence of the honeycomb structure, a smooth stator as well as a seal configuration with a honeycomb facing mounted on the stator wall were investigated. The seal geometry is representative of typical three-step labyrinth seals of modern aero engines. The flow field was predicted using a commercial finite volume code with the standard k-ε turbulence model. The computational grid includes the basic seal geometry as well as the three-dimensional honeycomb structures.

Numerical Investigation of a Cantilevered Compressor Stator at Varying Clearance Sizes

2015 ◽  
Author(s):  
Chengwu Yang ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
Shengfeng Zhao ◽  
Junqiang Zhu
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
High Flow ◽  
Leakage Flow ◽  
Streamwise Direction ◽  
Stall Margin ◽  
Axial Compressors ◽  
Pressure Surface ◽  
The Impact ◽  
Highly Loaded

The clearance size of cantilevered stators affects the performance and stability of axial compressors significantly. Numerical calculations were carried out using the commercial software FINE/Turbo for a 2.5-stage highly loaded transonic axial compressor, which is of cantilevered stator for the first stage, at varying hub clearance sizes. The aim of this work is to improve understanding of the impact mechanism of hub clearance on the performance and the flow field in high flow turning conditions. The performance of the front stage and the compressor with different hub clearance sizes of the first stator has been analyzed firstly. Results show that the efficiency decreases as clearance size varies from 0 to 3% of hub chordlength, but the operating range has been extended. For the first stage, the efficiency decreases about 0.5% and the stall margin is extended. The following analysis of detailed flow field in the first stator shows that the clearance leakage flow and elimination of hub corner separation is responsible for the increasing loss and stall margin extending respectively. The effects of hub clearance on the downstream rotor have been discussed lastly. It indicates that the loss of the rotor increases and the flow deteriorates due to increasing of clearance size and hence the leakage mass flow rate, which mainly results from the interaction of upstream leakage flow with the passage flow near pressure surface. The affected region of rotor passage flow field expands in spanwise and streamwise direction as clearance size grows. The hub clearance leakage flow moves upward in span as it flows toward downstream.

Characteristics of Stable Rotating Stall Cells in an Axial Compressor

2017 ◽  
Author(s):  
Adam R. Hickman ◽  
Scott C. Morris
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Pressure Ratio ◽  
Pressure Sensors ◽  
Axial Compressor ◽  
Field Measurements ◽  
Rotating Stall ◽  
Flow Coefficient ◽  
Stable Operation ◽  
Double Phase

Flow field measurements of a high-speed axial compressor are presented during pre-stall and post-stall conditions. The paper provides an analysis of measurements from a circumferential array of unsteady shroud static pressure sensors during stall cell development. At low-speed, the stall cell approached a stable size in approximately two rotor revolutions. At higher speeds, the stall cell developed within a short amount of time after stall inception, but then fluctuated in circumferential extent as the compressor transiently approached a stable post-stall operating point. The size of the stall cell was found to be related to the annulus average flow coefficient. A discussion of Phase-Locked Average (PLA) statistics on flow field measurements during stable operation is also included. In conditions where rotating stall is present, flow field measurements can be Double Phase-Locked Averaged (DPLA) using a once-per-revolution (1/Rev) pulse and the period of the stall cell. The DPLA method provides greater detail and understanding into the structure of the stall cell. DPLA data indicated that a stalled compressor annulus can be considered to contained three main regions: over-pressurized passages, stalled passages, and recovering passages. Within the over-pressured region, rotor passages exhibited increased blade loading and pressure ratio compared to pre-stall values.

Numerical Investigation of Effect of Inlet Distortion on Compressor Flow Field and Stability

2017 ◽  
Author(s):  
HaoGuang Zhang ◽  
Kang An ◽  
Feng Tan ◽  
YanHui Wu ◽  
WuLi Chu
Keyword(s):  
Numerical Simulation ◽  
Flow Field ◽  
Total Pressure ◽  
Tip Clearance ◽  
Numerical Work ◽  
Stall Margin ◽  
Numerical Investigations ◽  
Inlet Distortion ◽  
Compressor Rotor ◽  
The Stability

The compressor aerodynamic design is conducted under the condition of clean inlet in general, but a compressor often operates under the condition of inlet distortion in the practical application. It has been proven by a lot of experimental and numerical investigations that inlet distortion can decrease the performance and stability of compressors. The circumferential or radial distorted inlet in mostly numerical investigations is made by changing the total pressure and total temperature in the inlet ring surface of the compressors. In most of inlet distortion experiments, distorted inlets are usually created by using wire net, flashboards, barriers or the generator of rotating distortion. The fashion of generating distorted inlet for experiment is different from that for numerical simulation. Consequently, the flow mechanism of affecting the flow field and stability of a compressor with distorted inlet for experiment is partly different than that for numerical simulation. In the numerical work reported here, the inlet distortion is generated by setting some barriers in the inlet ring surface of an axial subsonic compressor rotor. Two kinds of distorted inlet are investigated to exploring the effect of distorted range on the flow field and stability of the compressor with ten-passage unsteady numerical method. The numerical results show that the inlet distortions not only degrade the total pressure and efficiency of the compressor rotor, but also decrease the stability of the rotor. The larger the range of distorted inlet is, the stronger the adverse effect is. The comprehensive stall margin for the inlet distortion of 24 degrees and 48 degrees of ten-passages is reduced about 3.35% and 5.88% respectively. The detailed analysis of the flow field in the compressor indicates that the blockage resulted from tip clearance leakage vortex (TLV) and the flow separation near the suction surfaces of some blades tip for distorted inlet is more serious than that resulted from TLV for clean inlet. Moreover, the larger the range of distorted inlet is, the larger the range of the blockage is. The analysis of unsteady flow shows that during this process, which is that one rotor blade passes through the region affected by the distorted inlet, the range of the blockage in the rotor passage increases first, then reduces, and increases last.

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