scholarly journals Reduction of Unsteady Blade Loading by Beneficial Use of Vortical and Potential Disturbances in an Axial Compressor With Rotor Clocking

1998 ◽  
Vol 120 (4) ◽  
pp. 705-713 ◽  
Author(s):  
S. T. Hsu ◽  
A. M. Wo
Keyword(s):  
Large Scale ◽  
Axial Compressor ◽  
Leading Edge ◽  
Forced Response ◽  
Suction Surface ◽  
Blade Loading ◽  
Beneficial Use ◽  
Unsteady Loading ◽  
Vortical Disturbance ◽  
First Time

This paper demonstrates reduction of stator unsteady loading due to forced response in a large-scale, low-speed, rotor/stator/rotor axial compressor rig by clocking the downstream rotor. Data from the rotor/stator configuration showed that the stator response due to the upstream vortical disturbance reaches a maximum when the wake impinges against the suction surface immediately downstream of the leading edge. Results from the stator/rotor configuration revealed that the stator response due to the downstream potential disturbance reaches a minimum with a slight time delay after the rotor sweeps pass the stator trailing edge. For the rotor/stator/rotor configuration, with Gap1 = 10 percent chord and Gap2 = 30 percent chord, results showed a 60 percent reduction in the stator force amplitude by clocking the downstream rotor so that the time occurrence of the maximum force due to the upstream vortical disturbance coincides with that of the minimum force due to the downstream potential disturbance. This is the first time, the authors believe, that beneficial use of flow unsteadiness is definitively demonstrated to reduce the blade unsteady loading.

scholarly journals Reduction of Unsteady Blade Loading by Beneficial Use of Vortical and Potential Disturbances in an Axial Compressor With Rotor Clocking

1997 ◽  
Author(s):  
S. T. Hsu ◽  
Andrew M. Wo
Keyword(s):  
Large Scale ◽  
Axial Compressor ◽  
Leading Edge ◽  
Forced Response ◽  
Suction Surface ◽  
Blade Loading ◽  
Beneficial Use ◽  
Unsteady Loading ◽  
Vortical Disturbance ◽  
First Time

This paper demonstrates reduction of stator unsteady loading due to forced response in a large-scale, low-speed, rotor/stator/rotor axial compressor rig by clocking the downstream rotor. Data from the rotor/stator configuration showed that the stator response due to the upstream vortical disturbance reaches a maximum when the wake impinges against the suction surface immediately downstream of the leading edge. Results from the stator/rotor configuration revealed that the stator response due to the downstream potential disturbance reaches a minimum with a slight time delay after the rotor sweeps pass the stator trailing edge. For the rotor/stator/rotor configuration, with Gap1= 10% chord and Gap2= 30% chord, results showed a 60% reduction in the stator force amplitude by clocking the downstream rotor so that the time occurrence of the maximum force due to the upstream vortical disturbance coincides with that of the minimum force due to the downstream potential disturbance. This is the first time, the authors believe, that beneficial use of flow unsteadiness is definitively demonstrated to reduce the blade unsteady loading.

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.

Method for Direct Parametric Analysis of Nonlinear Forced Response of Bladed Discs With Friction Contact Interfaces

2004 ◽  
Author(s):  
E. P. Petrov
Keyword(s):  
Parametric Analysis ◽  
Large Scale ◽  
Normal Load ◽  
Forced Response ◽  
Finite Element Models ◽  
Contact Interface ◽  
Friction Contact ◽  
Interface Parameters ◽  
First Time ◽  
Contact Parameters

An effective method for direct parametric analysis of periodic nonlinear forced response of bladed discs with friction contact interfaces has been developed. The method allows, for the first time, forced response levels to be calculated directly as a function of contact interface parameters such as the friction coefficient, contact surface stiffness (normal and tangential coefficients), clearances, interferences, and the normal stresses at the contact interfaces. The method is based on exact expressions for sensitivities of the multiharmonic interaction forces with respect to variation of all parameters of the friction contact interfaces. These novel expressions are derived in the paper for a friction contact model, accounting for the normal load variation and the possibility of separation-contact transitions. Numerical analysis of effects of the contact parameters on forced response levels has been performed using large-scale finite element models of a practical bladed turbine disc with underplatform dampers and with shroud contacts.

Three-Dimensional Turbulent Flow in the Tip Region of an Axial Compressor Rotor Passage at a Near Stall Condition

2001 ◽  
Author(s):  
Hongwei Ma ◽  
Haokang Jiang
Keyword(s):  
Turbulent Flow ◽  
Large Scale ◽  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Suction Surface ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Compressor Rotor

This paper presents an experimental study of the three-dimensional turbulent flow field in the tip region of an axial flow compressor rotor passage at a near stall condition. The investigation was conducted in a low-speed large-scale compressor using a 3-component Laser Doppler Velocimetry and a high frequency pressure transducer. The measurement results indicate that a tip leakage vortex is produced very close to the leading edge, and becomes the strongest at about 10% axial chord from the leading edge. Breakdown of the vortex periodically occurs at about 1/3 chord, causing very strong turbulence in the radial direction. Flow separation happens on the tip suction surface at about half chord, prompting the corner vortex migrating toward the pressure side. Tangential migration of the low-energy fluids results in substantial flow blockage and turbulence in the rear of a rotor passage. Unsteady interactions among the tip leakage vortex, the separated vortex and the corner flow should contribute to the inception of the rotating stall in a compressor.

scholarly journals The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

1995 ◽  
Vol 117 (4) ◽  
pp. 491-505 ◽  
Author(s):  
K. L. Suder ◽  
R. V. Chima ◽  
A. J. Strazisar ◽  
W. B. Roberts
Keyword(s):  
Surface Roughness ◽  
Surface Finish ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Blade Surface ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Performance Deterioration ◽  
Design Speed

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

scholarly journals Wake Vorticity Decay and Blade Response in an Axial Compressor With Varying Axial Gap

1999 ◽  
Author(s):  
A. M. Wo ◽  
M. H. Chung ◽  
S. J. Chang ◽  
S. F. Lee
Keyword(s):  
Large Scale ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Rotor Wake ◽  
Blade Loading ◽  
Wake Interaction ◽  
Blade Row ◽  
Unsteady Force ◽  
Gap Spacing ◽  
Blade Row Interaction

This paper addresses the decay of rotor wake vorticity for a rotor/stator axial compressor, with the axial gap between blade rows being 10, 20 and 30 percent chord, and at both design and high loading levels. Experiments were conducted in a large-scale, low-speed axial compressor. Navier-Stokes calculations were also executed. Both data and Navier-Stokes results reveal that the decay of rotor wake vorticity increases substantially as the axial gap decreases; the decay for 10 percent gap is about twice that of 30 percent. Increased time-mean blade loading causes the vorticity decay to also increase, with this effect more pronounced for large axial gap than small. At the stator inlet mid-pitch location, the wake maximum vorticity for 10 and 30 percent chord gap cases being nearly the same (differ by 3.8%) at design loading. The corresponding stator unsteady force agrees within 5.2%. Variation of vorticity decay with axial gap is directly linked to the change in potential disturbance by the downstream stator on the rotor wake due to the change in gap spacing. This suggests that the stator potential disturbance causes the upstream rotor wake to decay at an increased rate which, in turns, results in a lowered level of stator response compared to that without this stator/wake interaction effect. Thus, in this context, blade row interaction is considered beneficial.

scholarly journals Navier–Stokes and Potential Calculations of Axial Spacing Effect on Vortical and Potential Disturbances and Gust Response in an Axial Compressor

1997 ◽  
Vol 119 (3) ◽  
pp. 472-481 ◽  
Author(s):  
M.-H. Chung ◽  
A. M. Wo
Keyword(s):  
Spacing Effect ◽  
Axial Compressor ◽  
Leading Edge ◽  
Moving Frame ◽  
Navier Stokes ◽  
Two Dimensional ◽  
Blade Row ◽  
Vortical Disturbance ◽  
Blade Row Interaction ◽  
Gust Response

The effect of blade row axial spacing on vortical and potential disturbances and gust response is studied for a compressor stator/rotor configuration near design and at high loadings using two-dimensional incompressible Navier–Stokes and potential codes, both written for multistage calculations. First, vortical and potential disturbances downstream of the isolated stator in the moving frame are defined; these disturbances exclude blade row interaction effects. Then, vortical and potential disturbances for the stator/rotor configuration are calculated for axial gaps of 10, 20, and 30 percent chord. Results show that the potential disturbance is uncoupled locally; the potential disturbance calculated from the isolated stator configuration is a good approximation for that from the stator/rotor configuration upstream of the rotor leading edge at the locations studied. The vortical disturbance depends strongly on blade row interactions. Low-order modes of vortical disturbance are of substantial magnitude and decay much more slowly downstream than do those of potential disturbance. Vortical disturbance decays linearly with increasing mode except very close to the stator trailing edge. For a small axial gap, e.g., 10 percent chord, both vortical and potential disturbances must be included to determine the rotor gust response.

An Experimental Study of Heat Transfer in a Large-Scale Turbine Rotor Passage

1992 ◽  
Author(s):  
Michael F. Blair
Keyword(s):  
Heat Transfer ◽  
Experimental Study ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Turbine Rotor ◽  
Secondary Flows ◽  
Suction Surface ◽  
Tip Leakage ◽  
High Heat Transfer

An experimental study of the heat transfer distribution in a turbine rotor passage was conducted in a large–scale, ambient temperature, rotating turbine model. Meat transfer was measured for both the full–span suction and pressure surfaces of the airfoil as well as for the hub endwall surface. The objective of this program was to document the effects of flow three–dimensionality on the heat transfer in a rotating blade row (vs. a stationary cascade). Of particular interest were the effects of the hub and tip secondary flows, tip leakage and the leading–edge horseshoe vortexsystem. The effect of surface roughness on the passage heat transfer was also investigated. Midspan results are compared with both smooth–wall and rough–wall finite–difference two dimensional heat transfer predictions. Contour maps of Stanton number for both the rotor airfoil and endwall surfaces revealed numerous regions of high heat transfer produced by the three dimensional flows within the rotor passage. Of particular importance are regions of local enhancement (as much as 100% over midspan values) produced on the airfoil suction surface by the secondary flows and tip–leakage vortices and on the hub endwall by the leading–edge horseshoe vortex system.

Effect of Surface Roughness Location and Reynolds Number on Compressor Cascade Performance

2010 ◽  
Author(s):  
Seung Chul Back ◽  
Garth V. Hobson ◽  
Seung Jin Song ◽  
Knox T. Millsaps
Keyword(s):  
Surface Roughness ◽  
Reynolds Number ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Compressor Cascade ◽  
Suction Surface ◽  
Trailing Edge ◽  
Blade Loading ◽  
Linear Compressor Cascade

An experimental investigation has been conducted to characterize the influence of surface roughness location and Reynolds number on compressor cascade performance. Flow field surveys have been conducted in a low-speed, linear compressor cascade. Pressure, velocity, and flow angles have been measured via a 5-hole probe, pitot probe, and pressure taps on the blades. In addition to the entirely smooth and entirely rough blade cases, blades with roughness covering the leading edge; pressure side; and 5%, 20%, 35%, 50%, and 100% of suction side from the leading edge have been studied. All of the tests have been done for Reynolds number ranging from 300,000 to 640,000.Cascade performance (i.e. blade loading, loss, and deviation) is more sensitive to roughness on the suction side than pressure side. Roughness near the trailing edge of suction side increases loss more than that near the leading edge. When the suction side roughness is located closer to the trailing edge, the deviation and loss increase more rapidly with Reynolds number. For a given roughness location, there exists a Reynolds number at which loss begins to visibly increase. Finally, increasing the area of rough suction surface from the leading edge reduces the Reynolds number at which the loss coefficient begins to increase.

Influence of Surface Roughness on Three-Dimensional Separation in Axial Compressors

2004 ◽  
Vol 126 (4) ◽  
pp. 455-463 ◽  
Author(s):  
Semiu A. Gbadebo ◽  
Tom P. Hynes ◽  
Nicholas A. Cumpsty
Keyword(s):  
Surface Roughness ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Leading Edge ◽  
Major Effect ◽  
Surface Flow ◽  
Design Flow ◽  
Flow Coefficient ◽  
Navier Stokes ◽  
Suction Surface

Surface roughness on a stator blade was found to have a major effect on the three-dimensional (3D) separation at the hub of a single-stage low-speed axial compressor. The change in the separation with roughness worsened performance of the stage. A preliminary study was carried out to ascertain which part of the stator suction surface and at what operating condition the flow is most sensitive to roughness. The results show that stage performance is extremely sensitive to surface roughness around the leading edge and peak-suction regions, particularly for flow rates corresponding to design and lower values. Surface flow visualization and exit loss measurements show that the size of the separation, in terms of spanwise and chordwise extent, is increased with roughness present. Roughness produced the large 3D separation at design flow coefficient that is found for smooth blades nearer to stall. A simple model to simulate the effect of roughness was developed and, when included in a 3D Navier–Stokes calculation method, was shown to give good qualitative agreement with measurements.

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