scholarly journals Separated Flow Topology in Compressors

2019 ◽  
Vol 141 (9) ◽  
Author(s):  
James V. Taylor
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Separated Flow ◽  
Cell Behavior ◽  
Blade Design ◽  
Flow Topology ◽  
Design Point ◽  
Topological Reasoning ◽  
Blade Row ◽  
Different Response

Abstract When a multistage high-speed compressor is operated away from its design point, extreme incidence is caused in some blade rows. This results in large, localized separations that are three dimensional in nature. In this paper, topological reasoning is used to describe the behavior of these three-dimensional separations. It is shown that two classes of separation exist: one in which the flow progresses from attached to separate in a smooth way and another where there is a discontinuity in the response of the flow topology. It is shown that the global structure of the flow depends on the type of topological response that occurs. When the response is discontinuous, nonaxisymmetric cells of separated blades are formed. When the response is smooth, the resultant separated flow is axisymmetric. The paper is split into two broad sections: The first section presents examples of the two different classes of topological response that can occur in a single blade row, and it also shows how an engineer can achieve a different response by altering the blade design. The second section covers the analysis of a multistage high-speed compressor. The compressor initially presents the discontinuous behavior with rotating cells of separations. It is then redesigned to reduce the severity of the cell behavior or remove it entirely.

Experiments on the stability of sinusoidal flow over a circular cylinder

2002 ◽  
Vol 457 ◽  
pp. 157-180 ◽  
Author(s):  
TURGUT SARPKAYA
Keyword(s):  
Stokes Flow ◽  
Coherent Structures ◽  
High Speed ◽  
Three Dimensional ◽  
Separated Flow ◽  
Circular Cylinders ◽  
Centrifugal Forces ◽  
Complex Interactions ◽  
Sinusoidal Flow ◽  
The Stability

The instabilities in a sinusoidally oscillating non-separated flow over smooth circular cylinders in the range of Keulegan–Carpenter numbers, K, from about 0.02 to 1 and Stokes numbers, β, from about 103 to 1.4 × 106 have been observed from inception to chaos using several high-speed imagers and laser-induced fluorescence. The instabilities ranged from small quasi-coherent structures, as in Stokes flow over a flat wall (Sarpkaya 1993), to three-dimensional spanwise perturbations because of the centrifugal forces induced by the curvature of the boundary layer (Taylor–Görtler instability). These gave rise to streamwise-oriented counter-rotating vortices or mushroom-shaped coherent structures as K approached the Kh values theoretically predicted by Hall (1984). Further increases in K for a given β led first to complex interactions between the coherent structures and then to chaotic motion. The mapping of the observations led to the delineation of four states of flow in the (K, β)-plane: stable, marginal, unstable, and chaotic.

A Comparison Between the Design Point and Near-Stall Performance of an Axial Compressor

1990 ◽  
Vol 112 (1) ◽  
pp. 109-115 ◽  
Author(s):  
N. M. McDougall
Keyword(s):  
Rotor Blade ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Design Point ◽  
Blade Row ◽  
Clearance Flow ◽  
Stator Blade

Detailed measurements have been made within an axial compressor operating both at design point and near stall. Rotor tip clearance was found to control the performance of the machine by influencing the flow within the rotor blade passages. This was not found to be the case in the stator blade row, where hub clearance was introduced beneath the blade tips. Although the passage flow was observed to be altered dramatically, no significant changes were apparent in the overall pressure rise or stall point. Small tip clearances in the rotor blade row resulted in the formation of corner separations at the hub, where the blade loading was highest. More representative clearances resulted in blockage at the tip due to the increased tip clearance flow. The effects that have been observed emphasize both the three-dimensional nature of the flow within compressor blade passages, and the importance of the flow in the endwall regions in determining the overall compressor performance.

A Quasi-Three-Dimensional Turbomachinery Blade Design System: Part I—Throughflow Analysis

1985 ◽  
Vol 107 (2) ◽  
pp. 301-307 ◽  
Author(s):  
I. K. Jennions ◽  
P. Stow
Keyword(s):  
Three Dimensional ◽  
Computing System ◽  
Design System ◽  
Blade Design ◽  
Weighted Averages ◽  
Streamline Curvature ◽  
Blade Row ◽  
Blade Section ◽  
System Part ◽  
Analysis System

The purpose of this work has been to develop a quasi-three-dimensional blade design and analysis system incorporating fully linked throughflow, blade-to-blade and blade section stacking programs. In Part I of the paper, the throughflow analysis is developed. This is based on a rigorous passage averaging technique to derive throughflow equations valid inside a blade row. The advantages of this approach are that the meridional streamsurface does not have to be of a prescribed shape, and by introducing density weighted averages the continuity equation is of an exact form. Included in the equations are the effects of blade blockage, blade forces, blade-to-blade variations and loss. The solution of the equations is developed for the well-known streamline curvature method, and the contributions from these extra effects on the radial equilibrium equation are discussed. Part II of the paper incorporates the analysis into a quasi-three-dimensional computing system and demonstrates its operational feasibility.

Study of Blade Sweep Effects of Compressor

2005 ◽  
Author(s):  
C. Xu ◽  
R. S. Amano
Keyword(s):  
Flow Field ◽  
Turbine Blade ◽  
Experimental Studies ◽  
Three Dimensional ◽  
Experimental Results ◽  
Numerical Analyses ◽  
Blade Design ◽  
3D Flow ◽  
Efficient Manner ◽  
Blade Row

The three dimensional blading had been used for years in the process of turbomachine designs. In need of turbine blade designs in an efficient manner, the current advancement of CFD technologies allows effective 3D predictions of a complex 3D flow field in turbine blade passages, which can improve the turbine blade performances. Since numerous advantages of 3-D CFD usage had been reported in the open literature, many industries already started to use 3D blading in their turbomachines. In addition, a blade lean and a sweep for the blade design had been also implemented to increase the blade row efficiency. Experimental studies have shown some advantages of these lean and sweep features. Most of the experimental results combine many other features together. However, it is difficult to determine what the effects of different features should be. In this study, detailed numerical analyses were developed and these were used to present the results to gain better understanding of different feature of 3D blading for turbine designers and engineers. Throughout this paper performance impacts on different 3D features are presented and the superiority of the present approach is discussed.

A Comparison Between the Design Point and Near Stall Performance of an Axial Compressor

1989 ◽  
Author(s):  
N. M. McDougall
Keyword(s):  
Rotor Blade ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Design Point ◽  
Blade Row ◽  
Clearance Flow ◽  
Stator Blade

Detailed measurements have been made within an axial compressor operating both at design point and near stall. Rotor tip clearance was found to control the performance of the machine by influencing the flow within the rotor blade passages. This was not found to be the case in the stator blade row, where hub clearance was introduced beneath the blade tips. Although the passage flow was observed to be altered dramatically, no significant changes were apparent in the overall pressure rise or stall point. Small tip clearances in the rotor blade row resulted in the formation of corner separations at the hub, where the blade loading was highest. More representative clearances resulted in blockage at the tip due to the increased tip clearance flow. The effects which have been observed emphasize both the three dimensional nature of the flow within compressor blade passages, and the importance of the flow in the endwall regions in determining the overall compressor performance.

Improving Aerodynamic Matching of Axial Compressor Blading Using a Three-Dimensional Multistage Inverse Design Method

2005 ◽  
Vol 129 (1) ◽  
pp. 108-118 ◽  
Author(s):  
M. P. C. van Rooij ◽  
T. Q. Dang ◽  
L. M. Larosiliere
Keyword(s):  
Inverse Method ◽  
Design Method ◽  
Three Dimensional ◽  
Computational Time ◽  
Design Calculation ◽  
Pressure Loading ◽  
Design Intent ◽  
Design Point ◽  
Analysis And Design ◽  
Blade Row

Current turbomachinery design systems increasingly rely on multistage CFD as a means to diagnose designs and assess performance potential. However, design weaknesses attributed to improper stage matching are addressed using often ineffective strategies involving a costly iterative loop between blading modification, revision of design intent, and further evaluation of aerodynamic performance. A scheme is proposed herein which greatly simplifies the design point blade row matching process. It is based on a three-dimensional viscous inverse method that has been extended to allow blading analysis and design in a multi-blade row environment. For computational expediency, blade row coupling is achieved through an averaging-plane approximation. To limit computational time, the inverse method was parallelized. The proposed method allows improvement of design point blade row matching by direct regulation of the circulation capacity of the blading within a multistage environment. During the design calculation, blade shapes are adjusted to account for inflow and outflow conditions while producing a prescribed pressure loading. Thus, it is computationally ensured that the intended pressure-loading distribution is consistent with the derived blading geometry operating in a multiblade row environment that accounts for certain blade row interactions. The viability of the method is demonstrated in design exercises involving the rotors of a 2.5 stage, highly loaded compressor. Individually redesigned rotors display mismatching when run in the 2.5 stage, evident as a deviation from design intent. However, simultaneous redesign of the rotors in their multistage environment produces the design intent, indicating that aerodynamic matching has been achieved.

Study of Steady State and Transient Blade Row CFD Methods in a Moderately Loaded NASA Transonic High-Speed Axial Compressor Stage

2013 ◽  
Author(s):  
Mohammed Abdullah Qizar ◽  
Mahmoud L. Mansour ◽  
Shraman Goswami
Keyword(s):  
Steady State ◽  
High Speed ◽  
Transient Flow ◽  
Three Dimensional ◽  
Operating Conditions ◽  
Computational Time ◽  
Laser Doppler Velocimeter ◽  
Compressor Stage ◽  
Steady Stage ◽  
Blade Row

The effect of blade row interaction and hub leakage flow on the performance of moderately loaded NASA transonic hybrid compressor stage (Rotor 35 / Stator 37) is investigated through three-dimensional steady state and time-accurate, Navier Stokes calculations of the stage using the ANSYS CFX code at peak efficiency and near stall operating conditions. Understanding unsteady flow phenomena in compressor stages requires the use of time-accurate CFD simulations. Due to the inherent differences in blade counts between adjacent blade rows, the flow conditions at any given instant in adjacent blade rows differ. Depending on the blade counts, it may be necessary to model the entire annulus of the stage; however, this requires considerable computational time and memory resources. Several methods for modeling the transient flow in turbo machinery stages which require a minimal number of blade passages per row, and therefore reduced computational demands, have been presented in the literature. Recently, some of these methods have become available in commercial CFD solvers. The paper describes the steady and the unsteady CFD approaches used for investigating the ability to predict the measured performance of the NASA transonic axial stage design known as the hybrid stage, which consists of the axial Rotor35 and the axial stator 37. The steady approach employs the mixing-plane while the unsteady approaches are URANS with one based on full annulus simulation for the stage and the second enables simulations for the stage using reduced computational model, with a single passage from each blade row based on the time-tilting or the time-transformation technique. The above methods are evaluated and compared in terms of computational efficiency and comparison is made to steady stage simulations. Comparisons to overall performance data and two-dimensional Laser Doppler Velocimeter measurements of the velocity field are used to assess the predictive capabilities of the methods. Computed flow features are examined, and compared with reported measurements. This paper presents validation and calibration of methods used for determining blade row interactions and the respective predictive capabilities against the full annulus and the experimental test data.

Improving Aerodynamic Matching of Axial Compressor Blading Using a 3D Multistage Inverse Design Method

2005 ◽  
Author(s):  
M. P. C. van Rooij ◽  
T. Q. Dang ◽  
L. M. Larosiliere
Keyword(s):  
Design Method ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design Calculation ◽  
Pressure Loading ◽  
Design Point ◽  
Analysis And Design ◽  
Blade Row ◽  
Inverse Design Method ◽  
Stage Matching

Current turbomachinery design systems increasingly rely on multistage CFD as a means to diagnose designs and assess performance potential. However, design weaknesses attributed to improper stage matching are addressed using often ineffective strategies involving a costly iterative loop between blading modification, revision of design intent, and further evaluation of aerodynamic performance. A scheme is proposed herein which greatly simplifies the design point blade row matching process. It is based on a three-dimensional viscous inverse method that has been extended to allow blading analysis and design in a multi-blade row environment. For computational expediency, blade row coupling is achieved through an averaging-plane approximation. The proposed method allows improvement of design point blade row matching by direct regulation of the circulation capacity of the blading within a multistage environment. During the design calculation, blade shapes are adjusted to account for inflow and outflow conditions while producing a prescribed pressure loading. Thus, it is computationally ensured that the intended pressure-loading distribution is consistent with the derived blading geometry operating in a multi-blade row environment that accounts for certain blade row interactions. The viability of the method is explored in design exercises on a 2.5-stage, highly loaded compressor.

Analytical Investigation of a Low Pressure Turbine With and Without Flowpath Endwall Gaps, Seals and Clearance Features

2005 ◽  
Author(s):  
David Cherry ◽  
Aspi Wadia ◽  
Rob Beacock ◽  
Mani Subramanian ◽  
Paul Vitt
Keyword(s):  
Performance Prediction ◽  
High Speed ◽  
Pressure Ratio ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Low Pressure ◽  
Analytical Investigation ◽  
Mesh Topology ◽  
Low Pressure Turbine ◽  
Blade Row

Numerical simulations for low pressure turbine (LPT) stages of a high bypass turbofan engine are presented and discussed in this study. A smooth flowpath configuration and a flowpath configuration with endwall features consistent with the actual engine geometry were considered for the numerical analysis to demonstrate the significance of including hub and tip flowpath details for proper performance prediction and design improvement studies. Fully three-dimensional, multistage, mutiblock, viscous flow analysis methodology was applied for first three stages of a moderately loaded LPT to predict aerodynamic performance of individual components, stage and for the overall turbine. Numerical results were obtained first for the smooth endwall configuration that ignores flowpath cavities, gaps and leaks in the numerical model. Following the smooth endwall calculations, a second set of calculations was performed with hub and tip flowpath details to closely represent actual engine geometry and experimental rig hardware. The approach of using smooth endwall contours for multi stage, multi blade row computational analysis is quite common for modeling simplicity. However, as the flow features are expected to be more complex in high pressure ratio, highly loaded turbine stages of next generation aircraft engines, it is imperative that flowpath and endwall geometry details such as gaps, seals, leakage and clearance effects are included in the numerical simulation for improved component design and stage performance prediction. This study addresses this particular issue by including endwall details and quantifies performance differences between the two modeling approaches. An O-H mesh topology was utilized for the blades, wheel space cavities, labyrinth seals and clearances for better flowfield resolution and numerical accuracy. Component performance, secondary flow details of endwall cavities, seal leakage and loss features of each blade row, for individual stage and for the overall turbine stage is presented and discussed for the two sets of calculations. Computed results are compared with experimental data obtained with high speed rig testing for verification and for understanding of the flow physics.

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