Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor—Part I: Unsteady Profile Pressures and the Effect of Clocking

2004 ◽  
Vol 126 (4) ◽  
pp. 507-518 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Pressure Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Experimental Investigations ◽  
Low Speed ◽  
Unsteady Aerodynamic ◽  
Pressure Distributions ◽  
Blade Row ◽  
Pressure Changes

This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage Low-Speed Research Compressor of Dresden. In part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore, the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are calculated from the measured profile pressure distributions. The unsteady pressure distributions were analyzed in the first, a middle and the last compressor stage both on the rotor and stator blades. The measurements were carried out on pressure side and suction side at midspan. Several operating points were investigated. A complex behavior of the unsteady profile pressures can be observed, resulting from the superimposed influences of the wakes and the potential effects of several up- and downstream blade rows of the four-stage compressor. The profile pressure changes nearly simultaneously along the blade chord if a disturbance arrives at the leading edge or the trailing edge of the blade. Thus the unsteady profile pressure distribution is nearly independent of the convective wake propagation within the blade passage. A phase shift of the reaction of the blade to the disturbance on the pressure and suction side is observed. In addition, clocking investigations were carried out to distinguish between the different periodic influences from the surrounding blade rows. For this reason the unsteady profile pressure distribution on rotor 3 was measured, while stators 1–4 were separately traversed stepwise in the circumferential direction. Thus the wake and potential effects of the up- and downstream blade rows on the unsteady profile pressure could clearly be distinguished and quantified.

Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor: Part I — Unsteady Profile Pressures and the Effect of Clocking

2004 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Pressure Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Experimental Investigations ◽  
Complex Behaviour ◽  
Low Speed ◽  
Unsteady Aerodynamic ◽  
Pressure Distributions ◽  
Blade Row

This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage Low-Speed Research Compressor of Dresden. In part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are calculated from the measured profile pressure distributions. The unsteady pressure distributions were analysed in the first, a middle and the last compressor stage both on the rotor and stator blades. The measurements were carried out on pressure side and suction side at midspan. Several operating points were investigated. A complex behaviour of the unsteady profile pressures can be observed, resulting from the superimposed influences of the wakes and the potential effects of several up- and downstream blade rows of the four-stage compressor. The profile pressure changes nearly simultaneously along the blade chord if a disturbance arrives at the leading edge or the trailing edge of the blade. Thus the unsteady profile pressure distribution is nearly independent of the convective wake propagation within the blade passage. A phase shift of the reaction of the blade to the disturbance on the pressure and suction side is observed. In addition clocking investigations were carried out to distinguish between the different periodic influences from the surrounding blade rows. For this reason the unsteady profile pressure distribution on rotor 3 was measured, while stator 1–4 were separately traversed stepwise in the circumferential direction. Thus the wake and potential effects of the up- and downstream blade rows on the unsteady profile pressure could clearly be distinguished and quantified.

Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor: Part II — Unsteady Aerodynamic Forces of Rotor and Stator Blades

2004 ◽  
Author(s):  
Ronald Mailach ◽  
Lutz Mu¨ller ◽  
Konrad Vogeler
Keyword(s):  
Aerodynamic Force ◽  
Axial Compressor ◽  
Stability Limit ◽  
Experimental Investigations ◽  
Low Speed ◽  
Unsteady Aerodynamic ◽  
Pressure Distributions ◽  
Third Stage ◽  
Blade Row ◽  
Stator Blade

This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage Low-Speed Research Compressor of Dresden. In part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are determined from the measured profile pressure distributions. A method to calculate the aerodynamic blade forces on the basis of the experimental data is presented. The resulting aerodynamic blade forces are discussed for the rotor and stator blade rows of the first stage and the third stage of the compressor. Different operating points between design point and stability limit of the compressor were chosen to investigate the influence of loading on the aerodynamic force excitation. The time traces and the frequency contents of the unsteady aerodynamic blade force are discussed. Strong periodic influences of the incoming wakes and of potential effects of downstream blade rows can be observed. The amplitude and shape of the unsteady aerodynamic blade force depend on the interaction of the superimposed influences of the blade rows.

Rotor-Stator Interactions in a Four-Stage Low-Speed Axial Compressor—Part II: Unsteady Aerodynamic Forces of Rotor and Stator Blades

2004 ◽  
Vol 126 (4) ◽  
pp. 519-526 ◽  
Author(s):  
Ronald Mailach ◽  
Lutz Mu¨ller ◽  
Konrad Vogeler
Keyword(s):  
Aerodynamic Force ◽  
Axial Compressor ◽  
Stability Limit ◽  
Experimental Investigations ◽  
Low Speed ◽  
Unsteady Aerodynamic ◽  
Pressure Distributions ◽  
Third Stage ◽  
Blade Row ◽  
Stator Blade

This two-part paper presents detailed experimental investigations of unsteady aerodynamic blade row interactions in the four-stage low-speed research compressor of Dresden. In Part I of the paper the unsteady profile pressure distributions for the nominal setup of the compressor are discussed. Furthermore the effect of blade row clocking on the unsteady profile pressures is investigated. Part II deals with the unsteady aerodynamic blade forces, which are determined from the measured profile pressure distributions. A method to calculate the aerodynamic blade forces on the basis of the experimental data is presented. The resulting aerodynamic blade forces are discussed for the rotor and stator blade rows of the first stage and the third stage of the compressor. Different operating points between design point and stability limit of the compressor were chosen to investigate the influence of loading on the aerodynamic force excitation. The time traces and the frequency contents of the unsteady aerodynamic blade force are discussed. Strong periodic influences of the incoming wakes and of potential effects of downstream blade rows can be observed. The amplitude and shape of the unsteady aerodynamic blade force depend on the interaction of the superimposed influences of the blade rows.

Aerodynamic Blade Row Interactions in an Axial Compressor—Part II: Unsteady Profile Pressure Distribution and Blade Forces

2004 ◽  
Vol 126 (1) ◽  
pp. 45-51 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Boundary Layer ◽  
Pressure Distribution ◽  
Pressure Sensors ◽  
Stability Limit ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Experimental Investigations ◽  
Unsteady Pressure ◽  
Blade Row

This two-part paper presents experimental investigations of unsteady aerodynamic blade row interactions in the first stage of the four-stage low-speed research compressor of Dresden. Both the unsteady boundary layer development and the unsteady pressure distribution of the stator blades are investigated for several operating points. The measurements were carried out on pressure side and suction side at midspan. In Part II of the paper the investigations of the unsteady pressure distribution on the stator blades are presented. The experiments were carried out using piezoresistive miniature pressure sensors, which are embedded into the pressure and suction side surface of a single blade. The unsteady pressure distribution on the blade is analyzed for the design point and an operating point near the stability limit. The investigations show that it is strongly influenced by both the incoming wakes and the potential flow field of the downstream rotor blade row. If a disturbance arrives the leading edge or the trailing edge of the blade the pressure changes nearly simultaneously along the blade chord. Thus the unsteady profile pressure distribution is independent of the wake propagation within the blade passage. A phase shift of the reaction on pressure and suction side is observed. The unsteady response of the boundary layer and the profile pressure distribution is compared. Based on the unsteady pressure distribution the unsteady pressure forces of the blades are calculated and discussed.

Aerodynamic Blade Row Interactions in an Axial Compressor: Part II — Unsteady Profile Pressure Distribution and Blade Forces

2003 ◽  
Author(s):  
Ronald Mailach ◽  
Konrad Vogeler
Keyword(s):  
Boundary Layer ◽  
Pressure Distribution ◽  
Pressure Sensors ◽  
Stability Limit ◽  
Leading Edge ◽  
Side Surface ◽  
Suction Side ◽  
Experimental Investigations ◽  
Unsteady Pressure ◽  
Blade Row

This two-part paper presents experimental investigations of unsteady aerodynamic blade row interactions in the first stage of the four-stage Low-Speed Research Compressor of Dresden. Both the unsteady boundary layer development and the unsteady pressure distribution of the stator blades are investigated for several operating points. The measurements were carried out on pressure side and suction side at midspan. In part II of the paper the investigations of the unsteady pressure distribution on the stator blades are presented. The experiments were carried out using piezoresistive miniature pressure sensors, which are embedded into the pressure and suction side surface of a single blade. The unsteady pressure distribution on the blade is analysed for the design point and an operating point near the stability limit. The investigations show that it is strongly influenced by both the incoming wakes and the potential flow field of the downstream rotor blade row. If a disturbance arrives the leading edge or the trailing edge of the blade the pressure changes nearly simultaneously along the blade chord. Thus the unsteady profile pressure distribution is independent of the wake propagation within the blade passage. A phase shift of the reaction on pressure and suction side is observed. The unsteady response of the boundary layer and the profile pressure distribution is compared. Based on the unsteady pressure distribution the unsteady pressure forces of the blades are calculated and discussed.

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.

scholarly journals Aerodynamic Detuning Analysis of an Unstalled Supersonic Turbofan Cascade

1986 ◽  
Vol 108 (1) ◽  
pp. 60-67 ◽  
Author(s):  
D. Hoyniak ◽  
S. Fleeter
Keyword(s):  
Flow Field ◽  
Leading Edge ◽  
Rotor Blades ◽  
Driving Mechanism ◽  
Aeroelastic Stability ◽  
Aerodynamic Forces ◽  
Unsteady Aerodynamic ◽  
Influence Coefficients ◽  
Blade Row ◽  
The Stability

A new, and as yet unexplored, approach to passive flutter control is aerodynamic detuning, defined as designed passage-to-passage differences in the unsteady aerodynamic flow field of a rotor blade row. Thus, aerodynamic detuning directly affects the fundamental driving mechanism for flutter, i.e., the unsteady aerodynamic forces and moments acting on individual rotor blades. In this paper, a model to demonstrate the enhanced supersonic unstalled aeroelastic stability associated with aerodynamic detuning is developed. The stability of an aerodynamically detuned cascade operating in a supersonic inlet flow field with a subsonic leading edge locus is analyzed, with the aerodynamic detuning accomplished by means of nonuniform circumferential spacing of adjacent rotor blades. The unsteady aerodynamic forces and moments on the blading are defined in terms of influence coefficients in a manner that permits the stability of both a conventional uniformly spaced rotor configuration as well as the detuned nonuniform circumferentially spaced rotor to be determined. With Verdon’s uniformly spaced Cascade B as a baseline, this analysis is then utilized to demonstrate the potential enhanced aeroelastic stability associated with this particular type of aerodynamic detuning.

Investigation of Unsteady Flow Field in a Low-Speed One and a Half Stage Axial Compressor: Effects of Tip Gap Size on the Tip Clearance Flow Structure at Near Stall Operation

2014 ◽  
Author(s):  
Chunill Hah ◽  
Michael Hathaway ◽  
Joseph Katz
Keyword(s):  
Unsteady Flow ◽  
Flow Structure ◽  
Flow Instability ◽  
Axial Compressor ◽  
Leading Edge ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Low Speed ◽  
Clearance Flow ◽  
Tip Clearance Vortex

The primary focus of this paper is to investigate the effect of rotor tip gap size on how the rotor unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor at near stall operation (for example, where maximum pressure rise is obtained). A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at this flow condition with both a small and a large tip gaps. The numerically obtained flow fields at the small clearance matches fairly well with the available initial measurements obtained at the Johns Hopkins University with 3-D unsteady PIV in an index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. The numerical results are also compared with previously published measurements in a low speed single stage compressor (Maerz et al. [2002]). The current study shows that, with the smaller rotor tip gap, the tip clearance vortex moves to the leading edge plane at near stall operating condition, creating a nearly circumferentially aligned vortex that persists around the entire rotor. On the other hand, with a large tip gap, the clearance vortex stays inside the blade passage at near stall operation. With the large tip gap, flow instability and related large pressure fluctuation at the leading edge are observed in this one and a half stage compressor. Detailed examination of the unsteady flow structure in this compressor stage reveals that the flow instability is due to shed vortices near the leading edge, and not due to a three-dimensional separation vortex originating from the suction side of the blade, which is commonly referred to during a spike-type stall inception. The entire tip clearance flow is highly unsteady. Many vortex structures in the tip clearance flow, including the sheet vortex system near the casing, interact with each other. The core tip clearance vortex, which is formed with the rotor tip gap flows near the leading edge, is also highly unsteady or intermittent due to pressure oscillations near the leading edge and varies from passage to passage. For the current compressor stage, the evidence does not seem to support that a classical vortex breakup occurs in any organized way, even with the large tip gap. Although wakes from the IGV influence the tip clearance flow in the rotor, the major characteristics of rotor tip clearance flows in isolated or single stage rotors are observed in this one and a half stage axial compressor.

Numerical Simulation of Effects of Tip Geometry on the Performance of an Axial Compressor

2012 ◽  
Author(s):  
Hongwei Ma ◽  
Jun Zhang
Keyword(s):  
Mass Flow ◽  
Tangential Force ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Base Line ◽  
Compressor Rotor ◽  
Blade Tip

The purpose of this paper is to investigate numerically the effects of the tip geometry on the performance of an axial compressor rotor. There are three case studies which are compared with the base line tip geometry. 1) baseline (flat tip); 2) Cavity (tip with a cavity); 3) SSQA (suction side squealer tip) and 4) SSQB (modified suction side squealer tip). The case of SSQB is a combination of suction side squealer tip and the cavity tip. From leading edge to 10% chord, the tip has a cavity. From 10% chord to trailing edge, the tip has a suction side squealer. The numerical results of 2) show that the cavity tip leads to lower leakage mass flow and greater loss in tip gap and the rotor passage. The loading near the blade tip is lower than the baseline, thus the tangential force of the blade is lower. It leads to lower pressure rise than the baseline. The performance of the compressor for the tip with cavity is worse than the baseline. The results of 3) show that the higher curvature of the suction side squealer increases the loading of the blade and the tangential blade force. With the suction side squealer tip, the leakage flow experiences two vena contractor thus the mass of the leakage flow is reduced which is benefit for the performance of the compressor. The loss in the tip gap is lower than baseline. The performance is better than the baseline with greater pressure rise of the rotor, smaller leakage mass flow and lower averaged loss. For the case the SSQB, the leakage mass flow is lower than the SSQA and the loss in the tip gap and the rotor passage is greater than SSQA. The performance of the case of the SSQB is worse than the case of SSQA.

Visualizations of Flow Structures in the Rotor Passage of an Axial Compressor at the Onset of Stall

2017 ◽  
Vol 139 (4) ◽  
Author(s):  
Huang Chen ◽  
Yuanchao Li ◽  
David Tan ◽  
Joseph Katz
Keyword(s):  
High Speed ◽  
Large Scale ◽  
Incidence Angle ◽  
Axial Compressor ◽  
Leading Edge ◽  
Suction Side ◽  
Stereoscopic Particle Image Velocimetry ◽  
High Speed Imaging ◽  
Vortical Structures ◽  
Flow Phenomena

Experiments preformed in the JHU refractive index matched facility examine flow phenomena developing in the rotor passage of an axial compressor at the onset of stall. High-speed imaging of cavitation performed at low pressures qualitatively visualizes vortical structures. Stereoscopic particle image velocimetry (SPIV) measurements provide detailed snapshots and ensemble statistics of the flow in a series of meridional planes. At prestall condition, the tip leakage vortex (TLV) breaks up into widely distributed intermittent vortical structures shortly after rollup. The most prominent instability involves periodic formation of large-scale backflow vortices (BFVs) that extend diagonally upstream, from the suction side (SS) of one blade at midchord to the pressure side (PS) near the leading edge of the next blade. The 3D vorticity distributions obtained from data recorded in closely spaced planes show that the BFVs originate form at the transition between the high circumferential velocity region below the TLV center and the main passage flow radially inward from it. When the BFVs penetrate to the next passage across the tip gap or by circumventing the leading edge, they trigger a similar phenomenon there, sustaining the process. Further reduction in flow rate into the stall range increases the number and size of the backflow vortices, and they regularly propagate upstream of the leading edge of the next blade, where they increase the incidence angle in the tip corner. As this process proliferates circumferentially, the BFVs rotate with the blades, indicating that there is very little through flow across the tip region.

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