The Influence of Wake–Wake Interactions on Loss Fluctuations of a Downstream Axial Compressor Blade Row

1998 ◽  
Vol 120 (4) ◽  
pp. 695-704 ◽  
Author(s):  
G. J. Walker ◽  
J. D. Hughes ◽  
I. Ko¨hler ◽  
W. J. Solomon
Keyword(s):  
Free Stream ◽  
Guide Vane ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
Viscous Losses ◽  
Wake Interaction ◽  
Blade Row ◽  
Wake Interactions ◽  
Periodic Fluctuations ◽  
Disturbance Field

The interaction between wakes of an adjacent rotor–stator or stator–rotor blade row pair in an axial turbomachine is known to produce regular spatial variations in both the time-mean and unsteady flow fields in a frame relative to the upstream member of the pair. This paper examines the influence of such changes in the free-stream disturbance field on the viscous losses of a following blade row. Hot-wire measurements are carried out downstream of the outlet stator in a 1.5-stage axial compressor having equal blade numbers in the inlet guide vane (IGV) and stator rows. Clocking of the IGV row is used to vary the disturbance field experienced by the stator blades; the influence on stator wake properties is evaluated. The magnitude of periodic fluctuations in ensemble-averaged stator wake thickness is significantly influenced by IGV wake-rotor wake interaction effects. The changes in time-mean stator losses appear marginal.

scholarly journals The Influence of Wake-Wake Interactions on Loss Fluctuations of a Downstream Axial Compressor Blade Row

1997 ◽  
Author(s):  
G. J. Walker ◽  
J. D. Hughes ◽  
I. Köhler ◽  
W. J. Solomon
Keyword(s):  
Free Stream ◽  
Guide Vane ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
Viscous Losses ◽  
Wake Interaction ◽  
Blade Row ◽  
Wake Interactions ◽  
Periodic Fluctuations ◽  
Disturbance Field

The interaction between wakes of an adjacent rotor-stator or stator-rotor blade row pair in an axial turbomachine is known to produce regular spatial variations in both the time-mean and unsteady flow fields in a frame relative to the upstream member of the pair. This paper examines the influence of such changes in the free-stream disturbance field on the viscous losses of a following blade row. Hot-wire measurements are carried out downstream of the outlet stator in a 1.5-stage axial compressor having equal blade numbers in the inlet guide vane (IGV) and stator rows. Clocking of the IGV row is used to vary the disturbance field experienced by the stator blades: the influence on stator wake properties is evaluated. The magnitude of periodic fluctuations in ensemble-average stator wake thickness is significantly influenced by IGV wake-rotor wake interaction effects. The changes in time-mean stator losses appear marginal.

Steady and Unsteady Flow Field in a Multistage Low-Speed Axial Compressor: A Test Case

2008 ◽  
Author(s):  
Mario Ku¨nzelmann ◽  
Ralf Mu¨ller ◽  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Flow Field ◽  
Unsteady Flow ◽  
Laser Doppler Anemometry ◽  
Guide Vane ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
Test Case ◽  
Data Set ◽  
Low Speed ◽  
Blade Row

This paper introduces a new test case for compressor aerodynamics. The dataset is provided for the Dresden four-stage Low-Speed Research Compressor (LSRC), which was put into operation in 1995. The compressor consists of four identical stages, which are preceded by an inlet guide vane. The data set will be provided for the reference blading of the compressor with cantilevered stator vanes. This blading was developed on the basis of the profiles of a middle stage of a high-pressure compressor of a jet engine. This paper makes available the blading geometry as well as a variety of flow field measurement results. This includes the compressor map, selected pressure distributions and other results of flow field measurements with conventional techniques (e.g. Pitot probes, 5-hole probes). Furthermore different aspects of blade row interactions were addressed in this compressor within recent years. The periodical unsteady flow field within a selected rotor blade row was investigated using Laser-Doppler-Anemometry. Further results on the unsteady profile pressures and profile boundary layers will be provided. Supplementary, numerical results will be compared to the experiments. Results are available for several stages of the compressor and different operating points. With this test case a unique database for the aerodynamics in a multistage axial compressor will be provided that can be used for the validation of numerical codes.

Periodic Transition on an Axial Compressor Stator: Incidence and Clocking Effects: Part I—Experimental Data

1999 ◽  
Vol 121 (3) ◽  
pp. 398-407 ◽  
Author(s):  
G. J. Walker ◽  
J. D. Hughes ◽  
W. J. Solomon
Keyword(s):  
Free Stream ◽  
Guide Vane ◽  
Axial Compressor ◽  
Separated Flow ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Pressure Surface ◽  
Surface Pressure Distribution ◽  
Surface Transition ◽  
Hot Film

Periodic wake-induced transition on the outlet stator of a 1.5-stage axial compressor is examined using hot-film arrays on both the suction and pressure surfaces. The time-mean surface pressure distribution is varied by changing the blade incidence, while the free-stream disturbance field is altered by clocking of the stator relative to an inlet guide vane row. Ensemble-averaged plots of turbulent intermittency and relaxation factor (extent of calmed flow following the passage of a turbulent spot) are presented. These show the strength of periodic wake-induced transition phenomena to be significantly influenced by both incidence and clocking effects. The nature and extent of transition by other modes (natural, bypass, and separated flow transition) are altered accordingly. Leading edge and midchord separation bubbles are affected in a characteristically different manner by changing free-stream periodicity. There are noticeable differences between suction and pressure surface transition behavior, particularly as regards the strength and extent of calming. In Part II of this paper, the transition onset observations from the compressor stator are used to evaluate the quasi-steady application of conventional transition correlations to predict unsteady transition onset on the blading of an embedded axial compressor stage.

Natural Transition Phenomena on an Axial Compressor Blade

2000 ◽  
Author(s):  
J. D. Hughes ◽  
G. J. Walker
Keyword(s):  
Guide Vane ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
Instability Wave ◽  
Blade Loading ◽  
Instability Waves ◽  
Transition Prediction ◽  
Tollmien Schlichting ◽  
Vane Clocking ◽  
Hot Film

Data from a surface hot-film array on the outlet stator of a 1.5 stage axial compressor are analyzed to look for direct evidence of natural transition phenomena. An algorithm is developed to identify instability waves within the Tollmien Schlichting (T-S) frequency range. The algorithm is combined with a turbulent intermittency detection routine to produce space∼time diagrams showing the probability of instability wave occurrence prior to regions of turbulent flow. The paper compares these plots for a range of blade loading, with free-stream conditions corresponding to the maximum and minimum inflow disturbance periodicity produced by inlet guide vane clocking. Extensive regions of amplifying instability waves are identified in nearly all cases. The implications for transition prediction in decelerating flow regions on axial turbomachine blades are discussed.

A New Loss and Deviation Model for Axial Compressor Inlet Guide Vanes

2013 ◽  
Author(s):  
Milan Banjac ◽  
Milan V. Petrovic ◽  
Alexander Wiedermann
Keyword(s):  
Guide Vane ◽  
Thickness Distribution ◽  
Axial Compressor ◽  
Algebraic Model ◽  
Operating Conditions ◽  
Inlet Guide Vane ◽  
Viscous Effects ◽  
Through Flow ◽  
Trailing Vortices ◽  
Compressor Inlet

This paper describes a new universal algebraic model for the estimation of flow deflection and losses in axial compressor inlet guide vane devices. The model deals with nominal flow and far-off-design operating conditions in connection with large stagger angle adjustments. The first part of the model considers deflection and losses in 2D cascades, taking into account the main cascade geometry parameters and operating conditions, such as Mach number and stagger adjustment. The second part of the model deals with additional deviation and losses due to secondary flow caused by the endwall viscous effects and by the trailing vortices. The model is developed for NACA65 airfoils, NACA63-A4K6 airfoils and airfoils having an NACA65 thickness distribution on a circular-arc camber line. It is suitable for application in 1D or 2D through-flow calculations for design and analysis cases. The development of the method is based on systematic CFD flow calculations for various cascade geometries and operating parameters. The comparison of correlation results with experimental data for several test cases shows good agreement.

Wall Boundary Layer Development Near the Tip Region of an IGV of an Axial Flow Compressor

1984 ◽  
Vol 106 (2) ◽  
pp. 337-345
Author(s):  
B. Lakshminarayana ◽  
N. Sitaram
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Inlet Guide Vane ◽  
Wall Boundary Layer ◽  
Wall Boundary ◽  
Boundary Layer Development ◽  
Flow Compressor

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

The Influence of Turbulence on Wake Dispersion and Blade Row Interaction in an Axial Compressor

2005 ◽  
Author(s):  
Alan D. Henderson ◽  
Gregory J. Walker ◽  
Jeremy D. Hughes
Keyword(s):  
Boundary Layer ◽  
Guide Vane ◽  
Axial Compressor ◽  
Boundary Layer Transition ◽  
Inlet Guide Vane ◽  
Grid Turbulence ◽  
Layer Transition ◽  
Free Stream Turbulence ◽  
Hot Film ◽  
Blade Element

The influence of free-stream turbulence on wake dispersion and boundary layer transition processes has been studied in a 1.5-stage axial compressor. An inlet grid was used to produce turbulence characteristics typical of an embedded stage in a multistage machine. The grid turbulence strongly enhanced the dispersion of inlet guide vane (IGV) wakes. This modified the interaction of IGV and rotor wakes, leading to a significant decrease in periodic unsteadiness experienced by the downstream stator. These observations have important implications for the prediction of clocking effects in multistage machines. Boundary layer transition characteristics on the outlet stator were studied with a surface hot-film array. Observations with grid turbulence were compared with those for the natural low turbulence inflow to the machine. The transition behavior under low turbulence inflow conditions with the stator blade element immersed in the dispersed IGV wakes closely resembled the behavior with elevated grid turbulence. It is concluded that with appropriate alignment, the blade element behavior in a 1.5-stage axial machine can reliably indicate the blade element behavior of an embedded row in a multistage machine.

scholarly journals Multi-Objective Optimization in Axial Compressor Design Using a Linked CFD-Solver

2008 ◽  
Author(s):  
Kai Becker ◽  
Martin Lawerenz ◽  
Christian Voß ◽  
Reinhard Mo¨nig
Keyword(s):  
Guide Vane ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Inlet Guide Vane ◽  
Optimization Strategy ◽  
Objective Functions ◽  
Multi Objective Optimization ◽  
Multi Objective ◽  
Flow Information ◽  
Design Variation

In combination with a multi-objective 3D optimization strategy, a linked CFD-solver is presented in this paper, combining 3D-Reynolds-averaged-Navier-Stokes and an inviscid throughflow method. It enables the adjustment of the 3D boundary conditions for any design variation and contains new options for configuring the objective functions. The link is achieved by matching the flow information between both CFD codes in an iterative procedure. Compared to an individual 3D-CFD calculation, the convergence does not take significantly longer. The potential of the linked CFD-solver is demonstrated in a multi-objective optimization for one blade row to be optimized and one operating point at a 3-stage axial compressor with inlet guide vane. Within the optimization, the objective functions are formulated, so that the performance of the axial compressor is enhanced in addition to the improved efficiency of the 3D-cascade.

Influence of Gap Detailing on Calculated Unsteady Non-Adjacent Blade Row Aero-Forcing in a Transonic Compressor Stage

2020 ◽  
Vol 142 (11) ◽  
Author(s):  
Tobias Gezork ◽  
Paul Petrie-Repar
Keyword(s):  
Guide Vane ◽  
Computational Cost ◽  
Life Time ◽  
Forced Response ◽  
Inlet Guide Vane ◽  
Compressor Blades ◽  
Transonic Compressor ◽  
Forcing Function ◽  
Gap Flow ◽  
Blade Row

Abstract Resonant or close to resonant forced response excitation of compressor blades limits component life time and can potentially lead to high-cycle fatigue failure if the exciting forces are large and damping is insufficient. When numerically quantifying the forcing function by means of simulations, simplifications are typically made in the analysis to reduce complexity and computational cost. In this paper, we numerically investigate how the blade forcing function is influenced by the rotor tip gap flow and by flow across gaps in the upstream variable inlet guide vane row. Unsteady simulations are made using a test rig geometry where a forcing crossing with an excitation from a non-adjacent blade row had previously been measured. The effects of the gaps on the forcing function for the first torsion mode are presented for both the non-adjacent blade row excitation (changes compared with a case without gaps indicating a 20% reduction) and an adjacent excitation (changes indicating an 80% increase in terms of forcing function amplitude comparing with a case without gaps).

Full-Annulus Simulations of Airfoil Clocking in a 1-1/2 Stage Axial Compressor

1999 ◽  
Author(s):  
D. J. Dorney ◽  
D. L. Sondak ◽  
P. G. A. Cizmas ◽  
V. E. Saren ◽  
N. M. Savin
Keyword(s):  
Relative Motion ◽  
High Speed ◽  
Guide Vane ◽  
Research Effort ◽  
Axial Compressor ◽  
Inlet Guide Vane ◽  
Axial Compressors ◽  
Computational Research ◽  
Potential Interactions ◽  
Made In

Axial compressors have inherently unsteady flow fields because of relative motion between rotor and stator airfoils. This relative motion leads to viscous and inviscid (potential) interactions between blade rows. As the number of stages increases in a turbomachine, the buildup of convected wakes can lead to progressively more complex wake/wake and wake/airfoil interactions. Variations in the relative circumferential positions of stators or rotors can change these interactions, leading to different unsteady forcing functions on airfoils and different compressor efficiencies. In addition, as the Mach number increases the interaction between blade rows can be intensified due to potential effects. It has been shown, both experimentally and computationally, that airfoil clocking can be used to improve the efficiency and reduce the unsteadiness in multiple-stage axial turbomachines with equal blade counts in alternate blade rows. While previous investigations have provided an improved understanding of the physics associated with airfoil clocking, more research is needed to determine if airfoil clocking is viable for use in modern gas-turbine compressors. This paper presents the results of a combined experimental/computational research effort to study the physics of airfoil clocking in a high-speed axial compressor. Computational simulations have been performed for eight different clocking positions of the stator airfoils in a 1-1/2 stage high-speed compressor. To accurately model the experimental compressor, full-annulus simulations were conducted using 34 IGV, 35 rotor and 34 stator airfoils. It is common practice to modify blade counts to reduce the computational work required to perform turbomachinery simulations, and this approximation has been made in all computational clocking studies performed to date. A simulation was also performed in the present study with 1 inlet guide vane, 1 rotor airfoil, and 1 stator airfoil to model blade rows with 34 airfoils each in order to examine the effects of this approximation. Time-averaged and unsteady data (including performance and boundary layer quantities) were examined. The predicted results indicate that simulating the full annulus gives better qualitative agreement with the experimental data, as well as more accurately modeling the interaction between adjacent blade rows.

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