On the Mechanism of Spike Stall Inception and Near Stall Non-Synchronous Vibration in an Axial Compressor

2020 ◽  
Author(s):  
Daniel Möller ◽  
Heinz-Peter Schiffer
Keyword(s):  
Shear Layer ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Experimental Investigations ◽  
Flow Solver ◽  
Stall Inception ◽  
Transonic Compressor ◽  
Transient Data ◽  
Stable Characteristic ◽  
Operating Points

Abstract The aerodynamic and aeroelastic behavior of an engine-like transonic compressor stage is investigated in this paper. Simulations were carried out for the transonic compressor at the Technical University of Darmstadt using the AU3D flow solver. A comparison with previous experimental investigations shows close agreement for both steady and transient data. The simulations enable a detailed flow analysis during spike stall inception and non-synchronous vibrations (NSV) at operating points close to the stall limit. The aerodynamic structure of shear layer fluctuations in the rotor tip region during the initial phase of spike stall inception is investigated. Already before the stall limit is reached such fluctuations can occur. In this case, the compressor continues to operate on the aerodynamically stable characteristic and NSV might be excited due to the unsteady blade force. This paper concludes with a joint analytical interpretation of the shear layer fluctuations observed at the beginning of spike stall and during near stall NSV by basic flow principles.

On the Mechanism of Spike Stall Inception and Near Stall Non-Synchronous Vibration in an Axial Compressor

2020 ◽  
Author(s):  
Daniel Möller ◽  
Heinz-Peter Schiffer
Keyword(s):  
Shear Layer ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Experimental Investigations ◽  
Flow Solver ◽  
Stall Inception ◽  
Transonic Compressor ◽  
Transient Data ◽  
Stable Characteristic ◽  
Operating Points

Abstract The aerodynamic and aeroelastic behavior of an engine-like transonic compressor stage is investigated in this paper. Simulations were carried out for the transonic compressor at the Technical University of Darmstadt using the AU3D flow solver. A comparison with previous experimental investigations shows close agreement for both steady and transient data. The simulations enable a detailed flow analysis during spike stall inception and non-synchronous vibrations (NSV) at operating points close to the stall limit. The aerodynamic structure of shear layer fluctuations in the rotor tip region during the initial phase of spike stall inception is investigated. Already before the stall limit is reached such fluctuations can occur. In this case, the compressor continues to operate on the aerodynamically stable characteristic and NSV might be excited due to the unsteady blade force. This paper concludes with a joint analytical interpretation of the shear layer fluctuations observed at the beginning of spike stall and during near stall NSV by basic flow principles.

Effects of inlet radial distortion on the type of stall precursor in low-speed axial compressor

2016 ◽  
Vol 232 (1) ◽  
pp. 55-67 ◽  
Author(s):  
Ali Arshad ◽  
Qiushi Li ◽  
Simin Li ◽  
Tianyu Pan
Keyword(s):  
Pressure Fluctuations ◽  
Axial Compressor ◽  
Rotating Stall ◽  
Experimental Investigations ◽  
Inlet Flow ◽  
Blade Loading ◽  
Stall Inception ◽  
Low Speed ◽  
Modal Type ◽  
Stall Precursor

Experimental investigations of the effect of inlet blade loading on the rotating stall inception process are carried out on a single-stage low-speed axial compressor. Temporal pressure signals from the six high response pressure transducers are used for the analysis. Pressure variations at the hub are especially recorded during the stall inception process. Inlet blade loading is altered by installing metallic meshed distortion screens at the rotor upstream. Three sets of experiments are performed for the comparison of results, i.e. uniform inlet flow, tip, and hub distortions, respectively. Regardless of the type of distortion introduced to the inflow, the compressor undergoes a performance drop, which is more severe in the hub distortion case. Under the uniform inlet flow condition, stall inception is caused by the modal type disturbance while the stall precursor switched to spike type due to the highly loaded blade tip. In the presence of high blade loading at the hub, spike disappeared and the compressor once again witnessed a modal type disturbance. Hub pressure fluctuations are observed throughout the process when the stall is caused by a modal wave while no disturbance is noticed at the hub in spike type stall inception. It is believed that the hub flow separation contributes to the modal type of stall inception phenomenon. Results are also supported by the recently developed signal processing techniques for the stall inception study.

Numerical Investigations on Partial Surge Initiated Inception and Stall Evolution in a Transonic Compressor

2017 ◽  
Author(s):  
Jiaguo Hu ◽  
Tianyu Pan ◽  
Wenqian Wu ◽  
Qiushi Li ◽  
Yifang Gong
Keyword(s):  
High Performance ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Rotating Stall ◽  
Rapid Decline ◽  
Second Phase ◽  
Stall Inception ◽  
Transonic Compressor ◽  
Transient Simulations ◽  
Two Factors

The instability has been the largest barrier of the high performance axial compressor in the past decades. Stall inception, which determines the route and the characteristics of instability evolution, has been extensively focused on. A new stall inception, “partial surge”, is discovered in the recent experiments. In this paper full-annulus transient simulations are performed to study the origin of partial surge initiated inception and explain the aerodynamic mechanism. The simulations show that the stall inception firstly occurs at the stator hub region, and then transfers to the rotor tip region. The compressor finally stalled by the tip region rotating stall. The stall evolution is in accord with the experiments. The stall evolution can be divided into three phases. In the first phase, the stator corner separation gradually merged with the adjacent passages, producing an annulus stall cell at the stator hub region. In the second phase, the total pressure rise of hub region emerges rapid decline due to the fast expansion of the annulus stall cell, but the tip region maintains its pressure rise. In the third phase, a new rotating stall cell appears at the rotor tip region, leading to the onset of fast drop of the tip region pressure rise. The stall cells transfer from hub region to the tip region is caused by two factors, the blockage of the hub region which transfers more load to the tip region, and the separation fluid fluctuations in stator domain which increase the circumferential non-uniformity in the rotor domain. High load and non-uniformity at the rotor tip region induce the final rotating stall.

Breakdown of Tip Leakage Vortices in Compressors at Flow Conditions Close to Stall

1997 ◽  
Author(s):  
Stefan Schlechtriem ◽  
Michael Lötzerich
Keyword(s):  
Vortex Breakdown ◽  
Stability Limit ◽  
Flow Conditions ◽  
Mixed Flow ◽  
Stall Inception ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Flow Compressor ◽  
The Stability ◽  
Operating Points

The breakdown of tip leakage vortices at operating points close to the stability limit of transonic compressor rotors has been detected. The aerodynamic phenomenon is considered to have a major impact on stall inception. Computations have been carried out and a detailed visualization of the phenomenon is given. In addition the connection of vortex breakdown to rotating instabilities and stall is discussed. Furthermore the tip flow field of the axial rotor is compared to the results for a centrifugal and a mixed flow compressor operating at similar tip speeds.

Time Resolved Experimental Investigations of an Axial Compressor With Casing Treatment

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
R. Emmrich ◽  
H. Hönen ◽  
R. Niehuis
Keyword(s):  
Physical Phenomenon ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Suction Side ◽  
Tip Clearance ◽  
Experimental Investigations ◽  
Pressure Probe ◽  
Time Resolved ◽  
Casing Treatment ◽  
Compressor Rotor

A casing treatment with axial and radial skewed slots ending in a plenum chamber has experimentally been investigated at a highly subsonic axial compressor stage. The aim was to investigate the physical phenomenon of this treatment family that is responsible for the stabilization of the blade passage flow and the drop in efficiency mostly observed. The experimentally gained performance results of this configuration showed an extension of the operating range by approximately 50%, while the efficiency for design conditions is reduced by 1.4%. Apart from this, operating points at part load conditions have been observed nearly without any loss in efficiency. The detailed flow analysis is performed by means of results from a 3D pneumatic probe with temperature sensor and a dynamic total pressure probe. The focus of the investigations is on the incidence flow to the compressor rotor, the tip clearance vortex flow in combination with the wall stall separation region and the blade stall due to suction side separation. The casing treatment configuration is investigated with a special interest in detecting those effects which have an impact on the stability and the compressor overall efficiency, including the interaction of the rotor and the stator flow fields.

Short Length-Scale Rotating Stall Inception in a Transonic Axial Compressor: Criteria and Mechanisms

2006 ◽  
Author(s):  
Chunill Hah ◽  
Jo¨rg Bergner ◽  
Heinz-Peter Schiffer
Keyword(s):  
Axial Compressor ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Stall Inception ◽  
Tip Clearance Flow ◽  
Transonic Compressor ◽  
Eddy Simulation ◽  
Clearance Flow ◽  
Flow Mechanisms ◽  
Tip Clearance Vortex

The current paper reports on investigations aimed at advancing the understanding of the flow mechanism that leads to the onset of short-length scale rotating stall in a transonic axial compressor. Experimental data show large oscillation of the tip clearance vortex as the rotor operates near the stall condition. Inception of spike-type rotating stall is also measured in the current transonic compressor with high response pressure transducers. Computational studies of a single passage and the full annulus were carried out to identify flow mechanisms behind the spike-type stall inception in the current transonic compressor rotor. Steady and unsteady single passage flow simulations were performed, first to get insight into the interaction between the tip clearance vortex and the passage shock. The conventional Reynolds-averaged Navier-Stokes method with a standard turbulence closure scheme does not accurately reproduce tip clearance vortex oscillation and the measured unsteady pressure field. Consequently, a Large Eddy Simulation (LES) was carried out to capture more relevant physics in the computational simulation of the rotating stall inception. The unsteady random behavior of the tip clearance vortex and it’s interaction with the passage shock seem to be critical ingredients in the development of spike-type rotating stall in a transonic compressor. The Large Eddy Simulation was further extended to the full annulus to identify flow mechanisms behind the measured spike-type rotating stall inception. The current study shows that the spike-type rotating stall develops after the passage shock is fully detached from the blade passages. Interaction between the tip clearance vortex and the passage shock creates a low momentum area near the pressure side of the blade. As the mass flow rate decreases, this low momentum area moves further upstream and reversed tip clearance flow is initiated at the trailing edge plane. Eventually, the low momentum area near the pressure side reaches the leading edge and forward spillage of the tip clearance flow occurs. The flows in the affected blade passage or passages then stall. As the stalled blade passages are formed behind the passage shock, the stalled area rotates counter to the blade rotation just like the classical Emmon’s type rotating stall. Both the measurements and the computations show that the rotating stall cell covers one to two blade passage lengths and rotates at roughly 50% of the rotor speed.

scholarly journals Simulation of Casing Treatments of a Transonic Compressor Stage

2008 ◽  
Vol 2008 ◽  
pp. 1-10 ◽  
Author(s):  
M. Hembera ◽  
H.-P. Kau ◽  
E. Johann
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Compressor Stage ◽  
Flow Solver ◽  
Stall Margin ◽  
Transonic Compressor ◽  
Flow Mechanisms ◽  
Multistage Compressor ◽  
Casing Treatments

This article presents the study of casing treatments on an axial compressor stage for improving stability and enhancing stall margin. So far, many simulations of casing treatments on single rotor or rotor-stator configurations were performed. But as the application of casing treatments in engines will be in a multistage compressor, in this study, the axial slots are applied to a typical transonic first stage of a high-pressure 4.5-stage compressor including an upstream IGV, rotor, and stator. The unsteady simulations are performed with a three-dimensional time accurate Favre-averaged Navier-stokes flow solver. In order to resolve all important flow mechanisms appearing through the use of casing treatments, a computational multiblock grid consisting of approximately 2.4 million nodes was used for the simulations. The configurations include axial slots in 4 different variations with an axial extension ranging into the blade passage of the IGV. Their shape is semicircular with no inclination in circumferential direction. The simulations proved the effectiveness of casing treatments with an upstream stator. However, the results also showed that the slots have to be carefully positioned relative to the stator location.

Unsteady Aerodynamic Blade Excitation at the Stability Limit and During Rotating Stall in an Axial Compressor

2006 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Gas Turbines ◽  
Axial Compressor ◽  
Stability Limit ◽  
Rotating Stall ◽  
Operating Conditions ◽  
Experimental Investigations ◽  
Momentum Exchange ◽  
Stall Inception ◽  
The Stability ◽  
Force Excitation

The stable operating range of axial compressors is limited by the onset of rotating stall and surge. These flow conditions endanger the reliability of operation and have definitely to be avoided in compressors of gas turbines. However, there is still a need to improve the physical understanding of these flow phenomena to prevent them while utilizing the maximum available working potential of the compressor. This paper discusses detailed experimental investigations of the rotating stall onset with the main emphasis on the aerodynamic blade excitation in the Dresden four-stage Low-Speed Research Compressor. The stall inception, which is triggered by modal waves, as well as the main flow features during rotating stall operation are discussed. To investigate the unsteady pressure distributions, both the rotor and the stator blades of the first stage were equipped with piezoresistive pressure transducers. Based on these measurements the unsteady blade pressure forces are calculated. Time-resolved results at the stability limit as well as during rotating stall are presented. For all operating conditions rotor-stator-interactions play an important role on the blade force excitation. Furthermore the role of the inertia driven momentum exchange at the stall cell boundaries on the aerodynamic blade force excitation is pointed out.

Aerodynamic Effects in a Transonic Compressor With Nonaxisymmetric Tip Clearance

2018 ◽  
Author(s):  
Maximilian Jüngst ◽  
Samuel Liedtke ◽  
Heinz Peter Schiffer ◽  
Bernd Becker
Keyword(s):  
Mass Flow ◽  
Current Knowledge ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Stability Margin ◽  
Tip Clearance ◽  
Stable Operation ◽  
Stall Inception ◽  
Transonic Compressor ◽  
Eccentric Rotor

Future axial compressor designs tend to be built with larger relative tip gaps and eccentricity, since the core engines are reduced in size. Our knowledge of the aerodynamic effects due to eccentric tip gaps is largely based on low-speed work. The aim of this study is to widen current knowledge by using the 1.5 stage Darmstadt Transonic Compressor, which is representative of the front stage of a high pressure compressor. Efficiency, peak pressure rise and stability margin of the compressor are reduced linearly at design speed when the tip clearance is increased from 0.9% to 2.5% tip chord length. This holds true for configurations with eccentric rotor tip gap, if their circumferentially averaged gaps are considered. For a compressor with 96% eccentricity and 1.7% average tip clearance, corrected mass flow at rotor exit varies locally with up to ±20% and ±10% at stator exit, which can result in inlet distortions for subsequent stages in a multi-stage configuration. Also, the redistribution of flow massively influences stall inception during throttling at constant speed. Propagating disturbances are damped in sectors with higher inlet mass flow and lower incidence. Thus, overall operation remains stable, even though some sectors are highly disturbed. Consequently, the maximum clearance of an eccentric stage is not limiting the stable operation of the whole stage.

Some Aspects of Transonic Compressor Tandem Design

2018 ◽  
Author(s):  
A. Hergt ◽  
S. Grund ◽  
J. Klinner ◽  
W. Steinert ◽  
M. Beversdorff ◽  
...  
Keyword(s):  
Mach Number ◽  
Wake Flow ◽  
Objective Functions ◽  
Flow Solver ◽  
Axial Compressors ◽  
Transonic Compressor ◽  
Blade Row ◽  
High Level ◽  
Transonic Cascade ◽  
Operating Points

The development of modern axial compressors has already reached a high level. Therefore an enlargement of the design space by means of new or advanced aerodynamic methods is necessary in order to achieve further enhancements of performance and efficiency. The tandem arrangement of profiles in a transonic compressor blade row is such a method. It is necessary to address the design aspects a bit more in detail in order to efficiently apply this blading concept to turbomachinery. Therefore, in the current study the known design aspects of tandem blading in compressors will be summed up under consideration of the aerodynamic effects and construction characteristics of a transonic compressor tandem. Based on this knowledge, a new transonic compressor tandem cascade (DLRTTC) with an inflow Mach number of 0.9 is designed using modern numerical methods and a multi objective optimization process. Three objective functions as well as three operating points are used in the optimization. Furthermore, both tandem blades and their arrangement are parameterized. From the resulting database of 1246 members a final best member is chosen as state-of-the-art design for further detailed investigation. The aim of the ensuing experimental and numerical investigation is to answer the question, whether the tandem cascade resulting from the modern design process fulfills the described design aspects and delivers the requested performance and efficiency criteria. The numerical simulations within the study are carried out with the DLR flow solver TRACE. The experiments are performed at the Transonic Cascade Wind Tunnel of DLR in Cologne. The inflow Mach number during the tests is 0.9 and the AVDR [1, 2] is adjusted to 1.3 (design value). Wake measurements with a 3-hole probe are carried out in order to determine the cascade performance. The experimental results show an increase in losses and a reduction of the cascade deflection by about 2 degrees compared to design concept. Nevertheless, the experimental and numerical results allow a good understanding of the aerodynamic effects. In addition, Planar PIV was applied in a single S1 plane located at midspan to capture the velocity field in the wake of blade 1 in order to analyze the wake flow in detail and describe its influence on the cascade deflection and loss behavior. Finally, an outlook will be given on what future tandem compressor research should be focused.

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