scholarly journals Investigation of Design Features of Compressor Casing Treatment

2014 ◽  
Vol 61 (1) ◽  
pp. 153-161
Author(s):  
Vitaliy Nezym
Keyword(s):  
Rotor Blade ◽  
Axial Compressor ◽  
Design Features ◽  
Casing Treatment ◽  
Groove Configurations ◽  
Compressor Performance ◽  
Operational Range ◽  
Blade Tips

Abstract Casing treatment in the form of circumferential grooves over a rotor blade tips is used for improvement of an axial compressor performance. Usually, these grooves extend compressor’s stall range (stable operational range) but decrease its efficiency. In the paper, there are presented main results of investigations on grooves that influence positively efficiency of compressor. There were investigated traditional (typical) and newly developed groove configurations. Certain grooves combine increase in efficiency with extension in stall range.

scholarly journals A Study of Casing Treatment Stall Margin Improvement Phenomena

1975 ◽  
Author(s):  
D. C. Prince ◽  
D. C. Wisler ◽  
D. E. Hilvers
Keyword(s):  
Rotor Blade ◽  
Axial Compressor ◽  
Operating Conditions ◽  
Blade Angle ◽  
Normal Surface ◽  
Peak Efficiency ◽  
Stall Margin ◽  
Casing Treatment ◽  
Circumferential Groove ◽  
Wall Boundary Layer

The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot, and blade angle slot treatments were tested at low speeds. With the circumferential groove treatment the stalling flow was reduced 5.8 percent at negligible efficiency sacrifice. The axial-skewed slot treatment improved the stalling flow by 15.3 percent; 1.8 points in peak efficiency were sacrificed. The blade angle slot treatment improved the stalling flow by 15.0 percent; 1.4 points in peak efficiency were sacrificed. These values are consistent with previous experience at transonic speeds. The favorable stalling flow situations correlated well with observations of higher-than-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annulus wall pressure gradients, especially in the 50 to 75 percent chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations. Rotor blade wakes may be somewhat thinner in the presence of good treatments, particularly under operating conditions close to the baseline stall. Annulus wall boundary layer profiles are only slightly influenced by casing treatment.

Review—Axial Compressor Stall Phenomena

1980 ◽  
Vol 102 (2) ◽  
pp. 134-151 ◽  
Author(s):  
E. M. Greitzer
Keyword(s):  
Fluid Dynamic ◽  
Axial Compressor ◽  
Rotating Stall ◽  
Empirical Correlations ◽  
Flow Distortion ◽  
Stall Margin ◽  
Compression System ◽  
Casing Treatment ◽  
Linearized Stability ◽  
Compressor Performance

Stall in compressors can be associated with the initiation of several types of fluid dynamic instabilities. These instabilities and the different phenomena, surge and rotating stall, which result from them, are discussed in this paper. Assessment is made of the various methods of predicting the onset of compressor and/or compression system instability, such as empirical correlations, linearized stability analyses, and numerical unsteady flow calculation procedures. Factors which affect the compressor stall point, in particular inlet flow distortion, are reviewed, and the techniques which are used to predict the loss in stall margin due to these factors are described. The influence of rotor casing treatment (grooves) on increasing compressor flow range is examined. Compressor and compression system behavior subsequent to the onset of stall is surveyed, with particular reference to the problem of engine recovery from a stalled condition. The distinction between surge and rotating stall is emphasized because of the very different consequences on recoverability. The structure of the compressor flow field during rotating stall is examined, and the prediction of compressor performance in rotating stall, including stall/unstall hysteresis, is described.

Application of Casing Circumferential Grooves for Improved Stall Margin in a Transonic Axial Compressor

2002 ◽  
Author(s):  
D. C. Rabe ◽  
C. Hah
Keyword(s):  
Incidence Angle ◽  
Axial Compressor ◽  
Numerical Procedure ◽  
Leading Edge ◽  
Navier Stokes ◽  
Stall Margin ◽  
Casing Treatment ◽  
Transonic Compressor ◽  
Flow Mechanisms ◽  
Groove Configurations

Experimental and numerical investigations were conducted to study the fundamental flow mechanisms of circumferential grooves in the casing of a transonic compressor and their influence on compressor stall margin. Three different groove configurations were tested in a highly loaded transonic compressor. Experimental results show that circumferential grooves increase the stall margin of the compressor at the tested operating condition. Grooves with a much smaller depth than conventional designs are shown to be similarly effective in increasing the stall margin. Steady-state Navier-Stokes analyses were performed to study flow structures associated with each casing treatment. The numerical procedure calculates the overall effects of the circumferential grooves correctly. Detailed investigation of calculated flow fields indicates that losses are generated by interaction between the main passage flow and flow exiting the grooves. The grooves increase the stall margin by reducing the flow incidence angle on the pressure side of the leading edge, despite an overall increase in the endwall boundary layer thickness. This is due to complex interaction of the main passage flow with the additional radial and tangential flows created by the grooves.

scholarly journals Axial Compressor Mean-Line Analysis: Choking Modelling and Fully-Coupled Integration in Engine Performance Simulations

2021 ◽  
Vol 6 (1) ◽  
pp. 4
Author(s):  
Ioannis Kolias ◽  
Alexios Alexiou ◽  
Nikolaos Aretakis ◽  
Konstantinos Mathioudakis
Keyword(s):  
Axial Compressor ◽  
Engine Performance ◽  
Engine Model ◽  
Multi Stage ◽  
Transient Simulations ◽  
Full Knowledge ◽  
Compressor Performance ◽  
First Time ◽  
Performance Calculation ◽  
Fully Coupled

A mean-line compressor performance calculation method is presented that covers the entire operating range, including the choked region of the map. It can be directly integrated into overall engine performance models, as it is developed in the same simulation environment. The code materializing the model can inherit the same interfaces, fluid models, and solvers, as the engine cycle model, allowing consistent, transparent, and robust simulations. In order to deal with convergence problems when the compressor operates close to or within the choked operation region, an approach to model choking conditions at blade row and overall compressor level is proposed. The choked portion of the compressor characteristics map is thus numerically established, allowing full knowledge and handling of inter-stage flow conditions. Such choking modelling capabilities are illustrated, for the first time in the open literature, for the case of multi-stage compressors. Integration capabilities of the 1D code within an overall engine model are demonstrated through steady state and transient simulations of a contemporary turbofan layout. Advantages offered by this approach are discussed, while comparison of using alternative approaches for representing compressor performance in overall engine models is discussed.

Recirculation-groove coupled casing treatment for a transonic axial compressor

2021 ◽  
Vol 111 ◽  
pp. 106556
Author(s):  
Tien-Dung Vuong ◽  
Kwang-Yong Kim ◽  
Cong-Truong Dinh
Keyword(s):  
Axial Compressor ◽  
Casing Treatment ◽  
Transonic Axial Compressor

Compressor Performance and Operability in Swirl Distortion

2010 ◽  
Author(s):  
Yogi Sheoran ◽  
Bruce Bouldin ◽  
P. Murali Krishnan
Keyword(s):  
Gas Turbine ◽  
Complex Geometry ◽  
Full Range ◽  
Axial Compressor ◽  
Cfd Model ◽  
Generator System ◽  
Sliding Mesh ◽  
Wide Range ◽  
Compressor Performance ◽  
The Impact

Inlet swirl distortion has become a major area of concern in the gas turbine engine community. Gas turbine engines are increasingly installed with more complicated and tortuous inlet systems, like those found on embedded installations on Unmanned Aerial Vehicles (UAVs). These inlet systems can produce complex swirl patterns in addition to total pressure distortion. The effect of swirl distortion on engine or compressor performance and operability must be evaluated. The gas turbine community is developing methodologies to measure and characterize swirl distortion. There is a strong need to develop a database containing the impact of a range of swirl distortion patterns on a compressor performance and operability. A recent paper presented by the authors described a versatile swirl distortion generator system that produced a wide range of swirl distortion patterns of a prescribed strength, including bulk swirl, twin swirl and offset swirl. The design of these swirl generators greatly improved the understanding of the formation of swirl. The next step of this process is to understand the effect of swirl on compressor performance. A previously published paper by the authors used parallel compressor analysis to map out different speed lines that resulted from different types of swirl distortion. For the study described in this paper, a computational fluid dynamics (CFD) model is used to couple upstream swirl generator geometry to a single stage of an axial compressor in order to generate a family of compressor speed lines. The complex geometry of the analyzed swirl generators requires that the full 360° compressor be included in the CFD model. A full compressor can be modeled several ways in a CFD analysis, including sliding mesh and frozen rotor techniques. For a single operating condition, a study was conducted using both of these techniques to determine the best method given the large size of the CFD model and the number of data points that needed to be run to generate speed lines. This study compared the CFD results for the undistorted compressor at 100% speed to comparable test data. Results of this study indicated that the frozen rotor approach provided just as accurate results as the sliding mesh but with a greatly reduced cycle time. Once the CFD approach was calibrated, the same techniques were used to determine compressor performance and operability when a full range of swirl distortion patterns were generated by upstream swirl generators. The compressor speed line shift due to co-rotating and counter-rotating bulk swirl resulted in a predictable performance and operability shift. Of particular importance is the compressor performance and operability resulting from an exposure to a set of paired swirl distortions. The CFD generated speed lines follow similar trends to those produced by parallel compressor analysis.

scholarly journals Effects of Endwall Suction and Blowing on Compressor Stability Enhancement

1989 ◽  
Author(s):  
N. K. W. Lee ◽  
E. M. Greitzer
Keyword(s):  
Flow Injection ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Primary Objective ◽  
Stall Margin ◽  
Casing Treatment ◽  
Blade Row ◽  
Stall Margin Improvement ◽  
Stability Enhancement ◽  
Suction And Blowing

An experimental investigation was carried out to examine the effects on stall margin of flow injection into, and flow removal out of, the endwall region of an axial compressor blade row. A primary objective of the investigation was clarification of the mechanism by which casing treatment (which involves both removal and injection) suppresses stall in turbomachines. To simulate the relative motion between blade and treatment, the injection and removal took place through a slotted hub rotating beneath a cantilevered stator row. Overall performance data and detailed (time-averaged) flowfield measurements were obtained. Flow injection and removal both increased the stalling pressure rise, but neither was as effective as the wall treatment. Removal of high blockage flow is thus not the sole reason for the observed stall margin improvement in casing or hub treatment, as injection can also contribute significantly to stall suppression. The results also indicate that the increase in stall pressure rise with injection is linked to the streamwise momentum of the injected flow, and it is suggested that this should be the focus of further studies.

Compressor Performance 2D Modelling at Reverse Flow Conditions

2010 ◽  
Author(s):  
L. Gallar ◽  
I. Tzagarakis ◽  
V. Pachidis ◽  
R. Singh
Keyword(s):  
Experimental Data ◽  
Reverse Flow ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Engine Performance ◽  
Two Dimensional ◽  
Flow Conditions ◽  
Compressor Surge ◽  
Compressor Performance ◽  
Shaft Failure

After a shaft failure the compression system of a gas turbine is likely to surge due to the heavy vibrations induced on the engine after the breakage. Unlike at any other conditions of operation, compressor surge during a shaft over-speed event is regarded as desirable as it limits the air flow across the engine and hence the power available to accelerate the free turbine. It is for this reason that the proper prediction of the engine performance during a shaft over-speed event claims for an accurate modelling of the compressor operation at reverse flow conditions. The present study investigates the ability of the existent two dimensional algorithms to simulate the compressor performance in backflow conditions. Results for a three stage axial compressor at reverse flow were produced and compared against stage by stage experimental data published by Gamache. The research shows that due to the strong radial fluxes present over the blades, two dimensional approaches are inadequate to provide satisfactory results. Three dimensional effects and inaccuracies are accounted for by the introduction of a correction parameter that is a measure of the pressure loss across the blades. Such parameter is tailored for rotors and stators and enables the satisfactory agreement between calculations and experiments in a stage by stage basis. The paper concludes with the comparison of the numerical results with the experimental data supplied by Day on a four stage axial compressor.

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