The Flow Mechanism of How Distorted Flows Deteriorate Stability of an Axial Flow Compressor

2007 ◽  
Author(s):  
J. Zhang ◽  
F. Lin ◽  
J. Chen ◽  
C. Nie
Keyword(s):  
Cfd Simulation ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Resonance Phenomenon ◽  
Flow Mechanism ◽  
Stall Inception ◽  
Blade Passage ◽  
Inlet Distortion ◽  
Flow Compressor

The stalling behavior in a single-stage low-speed axial compressor under inlet distortion is investigated. A blade-passage-scale flow mechanism is proposed to explain the stability deterioration caused by inlet distortion for the tested compressor exhibiting spike stall inception. In contrast to the existing understanding of inlet distortion based on system scale dynamics, the main elements of this flow mechanism are the unsteady behavior of tip leakage vortices (TLV) under inlet distortion; its effect on the initiation of spike flow disturbances, and its interaction with distorted sectors. Rotating inlet distortion (RID) is used as a tool because RID makes it possible to directly compare the flows between distorted and clean flow sectors with fixed measurement stations on the casing, and the fact that the stationary inlet distortion is only a special case of RID makes the results generic. The tests demonstrate that the blade loading in the distorted sector is heavier than that in the non-distorted sector, causing the TLV in the distorted sector move closer to the leading edge of the rotor blade and thus be the first to initiate the spike-like disturbance. The unsteady CFD simulation further confirms that such a disturbance corresponds to a vortex spinning out of the leading edge of the blades. However, the initiation of this spike-like disturbance doesn’t necessarily trigger the full stall immediately. The tracking of the disturbances indicates that most of such spike-like disturbances will be smeared by non-distorted sector and the growth of the spike-like disturbances actually relate closely to how and how often the path of the propagating disturbances come across the path of the rotating distorted sector. The proposed blade-passage-scale flow mechanism also offers an alternative explanation to the “resonance” phenomenon in rotating inlet distortion research, which was explained with excitation-and-response theory for compressors that exhibit modal stall inception.

The Stability-Limiting Flow Mechanisms in a Subsonic Axial-Flow Compressor and Its Passive Control With Casing Treatment

2008 ◽  
Author(s):  
Xingen Lu ◽  
Junqiang Zhu ◽  
Chaoqun Nie ◽  
Weiguang Huang
Keyword(s):  
Flow Instability ◽  
State Of The Art ◽  
Axial Flow ◽  
Flow Mechanism ◽  
Blade Passage ◽  
Casing Treatment ◽  
Compressor Rotor ◽  
Flow Compressor ◽  
The Stability ◽  
End Wall

The phenomenon of flow instability in the compression system such as fan and compressor has been a long-standing “bottle-neck” problem for gas turbines/aircraft engines. With a vision of providing a state-of-the-art understanding of the flow field in axial-flow compressor in the perspective of enhancing their stability using passive means. Two topics are covered in this paper. The first topic is the stability-limiting flow mechanism close to stall, which is the basic knowledge needed to manipulate end-wall flow behavior for the stability improvement. The physical process occurring when approaching stall and the role of complex tip flow mechanism on flow instability in current high subsonic axial compressor rotor has been assessed using single blade passage computations. The second topic is flow instability manipulation with casing treatment. In order to advance the understanding of the fundamental mechanisms of casing treatment and determine the change in the flow field by which casing treatment improve compressor stability, systematic studies of the coupled flow through a subsonic compressor rotor and various end-wall treatments were carried out using a state-of-the-art multi-block flow solver. The numerically obtained flow fields were interrogated to identify complicated flow phenomenon around and within the end-wall treatments and describe the interaction between the rotor tip flow and end-wall treatments. Detailed analyses of the flow visualization at the rotor tip have exposed the different tip flow topologies between the cases with treatment casing and with untreated smooth wall. It was found that the primary stall margin enhancement afforded by end-wall treatments is a result of the tip flow manipulation. Compared to the smooth wall case, the treated casing significantly dampen or absorb the blockage near the upstream part of the blade passage caused by the upstream movement of tip clearance flow and weakens the roll-up of the core vortex. These mechanisms prevent an early spillage of low momentum fluid into the adjacent blade passage and delay the onset of flow instability.

Long-to-Short Length Scale Transition: A Stall Inception Phenomenon in an Axial Compressor With Inlet Distortion

2005 ◽  
Author(s):  
Feng Lin ◽  
Meilin Li ◽  
Jingyi Chen
Keyword(s):  
Axial Compressor ◽  
Axial Flow ◽  
Short Length ◽  
Scale Model ◽  
Future Research ◽  
Length Scale ◽  
Stall Inception ◽  
Inlet Distortion ◽  
Scale Transition ◽  
Asymmetric Flows

A theoretical and experimental study of stall inception processes in a three-stage low-speed axial flow compressor with inlet distortion is presented in this paper. Since inlet distortion provides asymmetric flows imposing onto the compressor, the main goal of this research is to unveil the mechanism of how such flows initiate long and/or short length-scale disturbances and how the compression system reacts to those disturbances. It is found that the initial disturbances are always triggered by the distorted flows, yet the growth of such disturbances depends on system dynamics. While in many cases the stall precursors were the short length scale spikes, there were some cases where the compressor instability was triggered after the disturbances going through a long-to-short length scale transition. A Moore-Greitzer based (system scale) model was proposed to qualitatively explain this phenomenon. It was found that when the compressor operated in a region where the nonlinearity of the characteristics dominated, long length-scale disturbances induced by the inlet distortion would evolve into short length-scale disturbances before they disappeared or triggered stall. However, the model was not able to predict the fact that many disturbances that were triggered by the distorted sector(s) were completely damped out in the undistorted sector(s). It is thus suggested that in future research of compressor instability, one should consider the flows in blade passage scale, the dynamics in system scale and their interaction simultaneously.

A Study of Stall Inceptions in a Low-Speed Axial-Flow Compressor With Various Radial Loadings

2009 ◽  
Author(s):  
J. Zhang ◽  
F. Lin ◽  
J. Chen ◽  
C. Nie
Keyword(s):  
Flow Rate ◽  
Axial Flow ◽  
High Loading ◽  
Separation Region ◽  
Flow Mechanism ◽  
Stall Inception ◽  
Axial Flow Compressor ◽  
Blade Tip ◽  
Flow Compressor ◽  
Large Flow

In this paper, the stall inceptions in a single-stage axial flow compressor with different high loading positions generated by various radial distortions are experimentally and numerically investigated. The results indicate that the stall limit varies with the radial position of the distortion significantly. The closer the position of distortion to the blade tip, the more unstable the compressor becomes. In addition, the results demonstrate that stall inception varies with radial distortion accordingly. While with the hub distortion, the compressor exhibits modal-like disturbances prior to the stall onset, the stall is triggered by the spike-like disturbances directly with the center and tip distortions. The flow mechanism is then further explored with numerical simulations. It is shown that in the hub distortion case, the separation region caused by local high loading in large flow rate can migrate to the tip region along the span as the compressor is throttled to the stall limit. This spanwise migration plays an important role in the formation of the modal-like disturbances. Compared to the hub distortion, the modal-like disturbance in the uniform inlet flow appears in a shorter period of time because it takes less time to initiate stall cell when the separation occurs. In the tip distortion case, the separation at tip dominates so strongly that no modal-like disturbances are found before the stall onset. A discussion is given at the end of this paper to explain why in some compressors, a modal inception emerges first and the stall is triggered by the spike later.

An Experimental and Computational Investigation of Tip Clearance Flow and Its Impact on Stall Inception

2010 ◽  
Author(s):  
Matthew A. Bennington ◽  
Mark H. Ross ◽  
Joshua D. Cameron ◽  
Scott C. Morris ◽  
Juan Du ◽  
...  
Keyword(s):  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Tip Clearance ◽  
Flow Coefficient ◽  
Momentum Balance ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Tip Clearance Flow ◽  
Clearance Flow

A numerical and experimental study was conducted to investigate the tip clearance flow and its relationship to stall in a transonic axial compressor. The CFD results were used to identify the existence of an interface between incoming axial flow and the reverse tip clearance flow. A surface streaking method was used to experimentally identify this interface as a line of zero axial shear stress at the casing. The position of this line, denoted xzs, moved upstream with decreasing flow coefficient in both the experiments and computations. The line was found to be at the rotor leading edge plane when the compressor stalled. Further measurements using rotor offset and inlet distortion further corroborated these results, and demonstrated that the movement of the interface upstream of the leading edge leads to the generation of rotating (“spike”) disturbances. Stall was therefore interpreted to occur as a result of a critical momentum balance between the approach fluid and the tip-leakage flow.

Large-Scale Detached-Eddy Simulation Analysis of Stall Inception Process in a Multistage Axial Flow Compressor

2017 ◽  
Vol 139 (7) ◽  
Author(s):  
Kazutoyo Yamada ◽  
Masato Furukawa ◽  
Yuki Tamura ◽  
Seishiro Saito ◽  
Akinori Matsuoka ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Large Scale ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Detached Eddy Simulation ◽  
Stall Inception ◽  
Eddy Simulation ◽  
Flow Compressor ◽  
Edge Separation

This paper describes the flow mechanisms of rotating stall inception in a multistage axial flow compressor of an actual gas turbine. Large-scale numerical simulations of the unsteady have been conducted. The compressor investigated is a test rig compressor that was used in the development of the Kawasaki L30A industrial gas turbine. While the compressor consists of a total of 14 stages, only the front stages of the compressor were analyzed in the present study. The test data show that the fifth or sixth stages of the machine are most likely the ones leading to stall. To model the precise flow physics leading to stall inception, the flow was modeled using a very dense computational mesh, with several million cells in each passage. A total of 2 × 109 cells were used for the first seven stages (3 × 108 cells in each stage). Since the mesh was still not fine enough for large-eddy simulation (LES), a detached-eddy simulation (DES) was used. Using DES, a flow field is calculated using LES except in the near-wall where the turbulent eddies are modeled by Reynolds-averaged Navier–Stokes. The computational resources required for such large-scale simulations were still quite large, so the computations were conducted on the K computer (RIKEN AICS in Japan). Unsteady flow phenomena at the stall inception were analyzed using data mining techniques such as vortex identification and limiting streamline drawing with line integral convolution (LIC) techniques. In the compressor studied, stall started from a separation on the hub side rather than the commonly observed leading-edge separation near the tip. The flow phenomenon first observed in the stalling process is the hub corner separation, which appears in a passage of the sixth stator when approaching the stall point. This hub corner separation grows with time, and eventually leads to a leading-edge separation on the hub side of the stator. Once the leading-edge separation occurs, it rapidly develops into a rotating stall, causing another leading-edge separation of the neighboring blade. Finally, the rotating stall spreads to the upstream and downstream blade rows due to its large blockage effect.

Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1987 ◽  
Vol 109 (1) ◽  
pp. 142-148 ◽  
Author(s):  
C. H. Law ◽  
A. J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

An Experimental Study of the Inception of Rotating Stall in a Single-Stage Low-Speed Axial Compressor

2007 ◽  
Author(s):  
Alexander K. Simpson ◽  
John P. Longley
Keyword(s):  
Axial Compressor ◽  
Axial Flow ◽  
Rotating Stall ◽  
Single Stage ◽  
Predictive Tool ◽  
Stall Inception ◽  
Low Speed ◽  
Blade Row ◽  
Flow Compressor ◽  
Critical Incidence

There are two established mechanisms, spike and modal inception, by which rotating stall is initiated in an axial flow compressor. Whilst the “Critical incidence hypothesis” and the “Zero slope criterion” are useful ideas in explaining the different stability boundaries for spikes and modes they do not provide the designer with a predictive tool. A detailed experimental investigation utilising a single-stage low-speed compressor is presented in which the aerodynamic environment of a rotor blade row is changed (rotor geometry is held fixed) so that it exhibited both spike and modal inception upon throttling into stall. The dominant mechanism of stall inception was found to be dependent on both the inlet flowfield and the downstream stator. The measurements are analysed and show that the meridional acceleration across the tip region of the rotor influences the mechanism by which rotating stall is incepted. This research is presented as a contribution towards the prediction of the stall inception mechanism.

Large-Scale DES Analysis of Stall Inception Process in a Multi-Stage Axial Flow Compressor

2016 ◽  
Author(s):  
Kazutoyo Yamada ◽  
Masato Furukawa ◽  
Yuki Tamura ◽  
Seishiro Saito ◽  
Akinori Matsuoka ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Large Scale ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Stall Inception ◽  
Eddy Simulation ◽  
Multi Stage ◽  
Flow Compressor ◽  
Edge Separation

The paper describes the flow mechanism of the rotating stall inception in a multi-stage axial flow compressor for an actual gas turbine. Large-scale numerical simulations have been conducted. The compressor investigated is a test rig compressor which was used for development of the industrial gas turbine, Kawasaki L30A. While the compressor consists of 14 stages, the front half stages of the compressor were analyzed in the present study. According to the test data, it is considered that the 5th or 6th stage is the one most suspected of leading to the stall. In order to capture precise flow physics that could happen at stall inception, a computational mesh was made dense, giving at least several million cells to each passage. It amounted to about two billion cells for the first 7 stages (three hundred million cells in each stage). Since the mesh was still not enough for the large-eddy simulation (LES), the detached-eddy simulation (DES) was employed. In the DES, a flow field is calculated by LES except near-wall and near-wall turbulent eddies are modeled by RANS. The computational resource required for such large-scale simulation was still quite large, so the computations were conducted on the K computer (RIKEN AICS in Japan). Unsteady flow phenomena at the stall inception were analyzed by using data mining techniques such as vortex identification and limiting streamline drawing with the LIC (line integral convolution) method. The present compressor has stall started from the separation on the hub side instead of the commonly observed leading-edge separation near the tip. The flow phenomenon first observed in the stalling process is the hub corner separation, which appears in some passage of the 6th stator when approaching the stall point. This hub corner separation expands with time, and eventually leads to the leading-edge separation on the hub side for the stator. Once the leading-edge separation happens, it rapidly develops into the rotating stall, causing another leading-edge separation for the neighboring blade in sequence. Finally, the rotating stall spreads to the upstream and downstream bladerows due to its large blockage effect.

Prediction of Inlet Distortion Induced Axial Compressor Flow Field Instability in Mid-Air Refueling

2003 ◽  
Author(s):  
Ningyu Liu ◽  
Eddie Yin-Kwee Ng ◽  
Hong Ngiap Lim ◽  
Tock Lip Tan
Keyword(s):  
Incidence Angle ◽  
Axial Compressor ◽  
Axial Flow ◽  
Flow Distortion ◽  
Practical Applications ◽  
Inlet Distortion ◽  
Compressor Inlet ◽  
Physical Insight ◽  
Flow Compressor ◽  
Compressor Efficiency

The propagation of strong distortion at inlet of an axial compressor is investigated by applying the critical distortion line and the integral method. The practical applications, such as flaming of leakage fuel during mid-air refueling process, are implemented to show the details of the numerical methodology used in analysis of the axial flow compressor behavior and the propagation of inlet distortion. From the viewpoint of compressor efficiency, the propagation of inlet flow distortion is further described by a compressor critical performance and its critical characteristic. The simulated results present a useful physical insight to the significant effects of inlet parameters on the distortion extension, velocity, and compressor characteristics. The distortion level, the size of distortion area, and the incidence angle at compressor inlet, and the rotor blade speed are found to be the major parameters affecting the mass flow rate of engine.

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