Details of Shrouded Stator Hub Cavity Flow in a Multi-Stage Axial Compressor Part 1: Interactions With the Primary Flow

2021 ◽  
Author(s):  
Nitya Kamdar ◽  
Fangyuan Lou ◽  
Nicole L. Key
Keyword(s):  
Axial Compressor ◽  
Cavity Flow ◽  
Leakage Flow ◽  
Clearance Ratio ◽  
Primary Flow ◽  
Flow Parameters ◽  
Multi Stage ◽  
Flow Interactions ◽  
Compressor Performance ◽  
The Impact

Abstract The flow in shrouded stator cavities can be quite complex with axial, radial, and circumferential variations. As the leakage flow recirculates and is re-injected into the main flow path upstream of the stator, it deteriorates the near-hub flow field and, thus, degrades the overall aerodynamic performance of the compressor. In addition, the windage heating in the cavity can raise thermal-mechanical concerns. Fully understanding the details of the shrouded-hub cavity flow in a multi-stage environment can enable better hub cavity designs. Since the majority of the open literature presents limited details about the structure of compressor cavity flows in the stator wells and how the cavity wells affect the leakage flow, there is a lack of wholistic knowledge of how these flow parameters are interdependent. To shed light on this topic, a coupled CFD model with inclusion of the stator cavity wells for the Purdue 3-Stage (P3S) Axial Compressor Research Facility using the PAX100 configuration was developed and validated against experimental data. Such a model not only quantifies the impact of cavity leakage flow on compressor performance, but it also provides the capability to investigate the flow structure details including the path of the fluid into and out of the cavity. With the model in place, in this part 1 paper, the influence of the hub leakage flow on compressor performance and its interactions with the primary flow were investigated by varying the clearance ratio of a single stator. The understanding of the primary-hub-leakage flow interactions can offer insights leading to better designs of hub cavities.

Details of Shrouded Stator Hub Cavity Flow in a Multi-Stage Axial Compressor Part 2: Leakage Flow Characteristics in Stator Wells

2021 ◽  
Author(s):  
Nitya Kamdar ◽  
Fangyuan Lou ◽  
Nicole L. Key
Keyword(s):  
Flow Characteristics ◽  
Axial Compressor ◽  
Cavity Flow ◽  
Flow Path ◽  
Circumferential Velocity ◽  
Leakage Flow ◽  
Labyrinth Seals ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
The Impact

Abstract The flow in shrouded stator cavities can be quite complex with axial, radial, and circumferential variations. As the leakage flow recirculates and is re-injected into the main flow path upstream of the stator, it deteriorates the near-hub flow field and, thus, degrades the overall aerodynamic performance of the compressor. In addition, the windage heating in the cavity can raise thermal-mechanical concerns. Fully understanding the details of the shrouded-hub cavity flow in a multi-stage environment can enable better hub cavity designs. In the first part of the paper, the influence of the hub leakage flow on compressor performance and its interactions with the primary flow were investigated. While the impact of hub leakage flow on the primary passage is readily available in the open literature, details inside the cavity geometry are scarce due to the difficulties in instrumenting that region for an experiment or modeling the full cavity geometry. To shed light on this topic, the flow physics in the stator cavity inlet and outlet wells are investigated in the present paper using a coupled CFD model with inclusion of the stator cavity wells for the Purdue 3-Stage (P3S) Axial Compressor, which is representative of the rear stages of a high-pressure-compressor in core engines. At the inlet cavity, the presence of at least one pair of vortices influences the trajectory of the cavity leakage flow. The amount of leakage flow also determines the size of the vortical structures, with larger clearances creating a smaller vortex and vice versa. After passing through the labyrinth seals, the leakage flow travels along the stator landing first and then transitions to the rotor drum. In general, a flow path closer to the rotor drum achieves higher circumferential velocity but also exhibits significant temperature rise. A rise in circumferential velocity directly corresponds to a rise in temperature. In addition, the windage heating increases with increasing seal clearance. Furthermore, the inlet well contributes the most to overall windage, nearly 50% of the total windage heating, while the labyrinth seals and outlet well account for very little.

Details of Shrouded Stator HUB Cavity Flow in a Multi-Stage Axial Compressor Part 2: Leakage Flow Characteristics in Stator Wells

2021 ◽  
Author(s):  
Nitya Kamdar ◽  
Fangyuan Lou ◽  
Nicole L. Key
Keyword(s):  
Flow Characteristics ◽  
Axial Compressor ◽  
Leakage Flow ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Present Part ◽  
Compressor Performance ◽  
The Impact ◽  
Shed Light ◽  
Pressure Compressor

Abstract In the first part of the paper, the influence of the hub leakage flow on compressor performance and its interactions with the primary flow were investigated. While the impact of hub leakage flow on the primary passage is readily available in the open literature, details inside the cavity geometry are scarce due to the difficulties in instrumenting that region for an experiment or modeling the full cavity geometry. To shed light on this topic, the flow physics in the stator cavity inlet and outlet wells were investigated in the present part of the paper to understand the flow path of the leakage fluid and windage heating within the cavity using a coupled CFD model with inclusion of the stator cavity wells for the Purdue 3-Stage (P3S) Axial Compressor, which is representative of the rear stages of a high-pressure-compressor in core engines.

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.

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.

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.

Interaction of Shroud Leakage Flow and Main Flow in a Three-Stage LP Turbine

2003 ◽  
Vol 127 (4) ◽  
pp. 649-658 ◽  
Author(s):  
Jochen Gier ◽  
Bertram Stubert ◽  
Bernard Brouillet ◽  
Laurent de Vito
Keyword(s):  
Main Flow ◽  
Secondary Flows ◽  
Test Facility ◽  
Navier Stokes ◽  
Leakage Flow ◽  
Jet Engines ◽  
Flow Interactions ◽  
Lp Turbine ◽  
The Impact ◽  
The Ideal

Endwall losses significantly contribute to the overall losses in modern turbomachinery, especially when aerodynamic airfoil load and pressure ratios are increased. In turbines with shrouded airfoils a large portion of these losses are generated by the leakage flow across the shroud clearance. Generally the related losses can be grouped into losses of the leakage flow itself and losses caused by the interaction with the main flow in subsequent airfoil rows. In order to reduce the impact of the leakage flow and shroud design related losses a thorough understanding of the leakage losses and especially of the losses connected to enhancing secondary flows and other main flow interactions has to be understood. Therefore, a three stage LP turbine typical for jet engines is being investigated. For the three-stage test turbine 3D Navier-Stokes computations are performed simulating the turbine including the entire shroud cavity geometry in comparison with computations in the ideal flow path. Numerical results compare favorably against measurements carried out at the high altitude test facility at Stuttgart University. The differences of the simulations with and without shroud cavities are analyzed for several points of operation and a very detailed quantitative loss breakdown is presented.

Prediction of Tip Leakage Vortex Dynamics in an Axial Compressor Cascade Using RANS Analyses

2020 ◽  
Author(s):  
Martina Ricci ◽  
Roberto Pacciani ◽  
Michele Marconcini ◽  
Andrea Arnone
Keyword(s):  
Vortex Dynamics ◽  
Eddy Viscosity ◽  
Axial Compressor ◽  
Viscosity Model ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Eddy Viscosity Model ◽  
The Impact

Abstract The tip leakage flow in turbine and compressor blade rows is responsible for a relevant fraction of the total loss. It contributes to unsteadiness, and have an important impact on the operability range of compressor stages. Experimental investigations and, more recently, scale-resolving CFD approaches have helped in clarifying the flow mechanism determining the dynamics of the tip leakage vortex. Due to their continuing fundamental role in design verifications, it is important to establish whether RANS/URANS approaches are able to reproduce the effects of such a flow feature, in order to correctly drive the design of the next generation of turbomachinery. Base studies are needed in order to accomplish this goal. In the present work the tip leakage flow in axial compressor rotor blade cascade have been studied. The cascade was tested experimentally in Virginia Tech Low Speed Cascade Wind Tunnel in both stationary and moving endwall configurations. Numerical analyses were performed using the TRAF code, a state-of-the-art in-house-developed 3D RANS/URANS flow solver. The impact of the numerical framework was investigated selecting different advection schemes including a central scheme with artificial dissipation and a high-resolution upwind strategy. In addition, two turbulence models have been used, the Wilcox linear k–ω model and a non-linear eddy viscosity model (Realizable Quadratic Eddy Viscosity Model), which accounts for turbulence anisotropy. The numerical results are scrutinized using the available measurements. A detailed discussion of the vortex evolution inside the blade passage and downstream of the blade trailing edge is presented in terms of streamwise velocity, streamwise vorticity, and turbulent kinetic energy contours. The purpose is to identify guidelines for obtaining the best representation of the vortex dynamics, with the methodologies usually employed in routine design iterations and, at the same time, evidence their weak aspects that need further modelling efforts.

The Stacking of Compressor Stage Characteristics to Give an Overall Compressor Performance Map

1962 ◽  
Vol 13 (4) ◽  
pp. 349-367 ◽  
Author(s):  
M. D. C. Doyle ◽  
S. L. Dixon
Keyword(s):  
Axial Compressor ◽  
Compressor Stage ◽  
Individual Stage ◽  
Optimum Performance ◽  
Method Of Calculation ◽  
Multi Stage ◽  
Performance Map ◽  
Compressor Performance ◽  
Overall Performance ◽  
The Individual

SummaryA method of calculation is developed to compute the overall performance of a multi-stage axial compressor, from a knowledge of the individual stage characteristics, by a “stacking” technique. Compressor models are designed and their overall performance calculated. These results are compared to show, qualitatively, the effect of alterations in design and stage performance on overall performance and to find how compressors should be designed for optimum performance.

scholarly journals A Navier-Stokes Analysis of the Effect of Tip Clearance on Compressor Stall Margin

1995 ◽  
Author(s):  
Robert P. Dring ◽  
William D. Sprout ◽  
Harris D. Weingold
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Stall Margin ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Practical Tool ◽  
The Impact ◽  
Pressure Compressor

A three-dimensional Navier-Stokes calculation was used to analyze the impact of rotor tip clearance on the stall margin of a multi-stage axial compressor. This paper presents a summary of: (1) a study of the sensitivity of the results to grid refinement, (2) an assessment of the calculation’s ability to predict stall margin when the stalling row was the first rotor in a multi-stage rig environment, (3) an analysis of the impact of including the effects of the downstream stator through body force effects on the upstream rotor, and (4) the ability of the calculation to predict the impact of tip clearance on stall margin through a calculation of the rear seven airfoil rows of an eleven stage high pressure compressor rig. The result of these studies was that a practical tool is available which can predict stall margin, and the impact of tip clearance, with reasonable accuracy.

Influence of the Leakage Flow Tangential Velocity on the Loss Generation and Leakage Flow Kinematics in Shrouded Axial Compressor Cascades

2006 ◽  
Author(s):  
D. W. Sohn ◽  
T. Kim ◽  
S. J. Song
Keyword(s):  
Tangential Velocity ◽  
Axial Compressor ◽  
Suction Side ◽  
Velocity Variation ◽  
Leakage Flow ◽  
Compressor Blades ◽  
Blade Passage ◽  
Leakage Flows ◽  
And Behavior ◽  
The Impact

Although compressor blades have long been shrouded for aerodynamic and structural reasons, the impact of the leakage flow in the shroud cavities on passage flows has only recently been investigated. Furthermore, the tangential velocity of the leakage flow, set by the blading and the relative motion between rotating and stationary surfaces, has a strong influence on the passage flow. Yet the influence of the tangential velocity variation on the kinematics and dynamics (loss) of the leakage flow (from its ingress to egress) in the shrouded cavity and main flow in the blade passage are unknown. Therefore, this paper reports on an experimental investigation of the axial evolution of loss generation in the blade passage and behavior of the leakage flow in the seal cavity in shrouded axial compressor cascades subject to the variation of leakage tangential velocity. The newly found results are as follows. First, increasing tangential velocity of the leakage flow reduces loss at 10% and 50% chordwise locations in the passage. However, most of the blockage and loss reductions occurs in the aft half chord and downstream of the blade passage. Second, the increasing tangential velocity spreads the loss core, which is originally concentrated in the suction side hub corner, in the pitchwise direction. Thus, the loss core becomes more two-dimensional, and the region’s radial extent is reduced. Third, increasing tangential velocity of the leakage flow makes the near hub passage flow more radially uniform. Consequently, the shear and resultant mixing loss between the passage and leakage flows are reduced near the hub, reducing the overall loss. Finally, the leakage flow is ingested through the downstream cavity and makes an abrupt turn at the seal tooth. Thus, two distinct flow regions — downstream and upstream of the single-tooth seal — are found. Before the leakage flow rejoins the mainstream via the upstream cavity trench, the leakage flow circumferentially migrates in the direction of rotation. The magnitude of the circumferential shift depends strongly on the leakage tangential velocity.

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