Acoustic and Convective Mechanisms Contributing to Non-Synchronous-Vibrations in a Multistage Compressor

2019 ◽  
Author(s):  
Christoph Brandstetter ◽  
Benoit Paoletti ◽  
Xavier Ottavy
Keyword(s):  
High Speed ◽  
Laser Doppler Anemometry ◽  
Leading Edge ◽  
Acoustic Resonance ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Unsteady Pressure ◽  
Mode Decomposition ◽  
Design Speed ◽  
Lock In

Abstract In the present paper we show an analysis of Non-Synchronous-Oscillations in the rear stage of an experimental high-speed 3.5 stage compressor. The machine CREATE is representative of the rear median-stages of a modern civil turbofan engine except for an increased tip clearance on middle and last stage. Using distributed unsteady pressure measurements a comprehensive dataset during transient throttling maneuvers is presented, including temporal and spatial mode decomposition and the derivation of wave propagation speeds. To support the aerodynamic characterization, the steady flow structure at near stall conditions has been investigated using Laser-Doppler-Anemometry and unsteady pressure probes. In repeated experiments the machine has encountered alternating circumferential modes, several hundred revolutions before rotating stall at design speed. It was found that acoustic resonance of an upstream cut-on mode between the compressor stages enforces the development of particular modes. These phenomena have not been observed on the same machine with regular tip clearance. In contrast to front-stage observations reported in literature, where leading edge separations cause strong excitations of the blade eigenmodes, initially weak modulations of the flow in the rotor tip region are observed to be amplified due to acoustic feedback. With rising amplitude lock-in to distinct circumferential mode orders develops. Apart from the final lock-in procedure, signatures of solely aerodynamic fluctuations, discussed in literature as ‘Rotating Instabilities’ or ‘Part Span Stall’, are observed intermittently at different speedlines.

Rotating Stall Observations in a High Speed Compressor—Part II: Numerical Study

2015 ◽  
Vol 137 (5) ◽  
Author(s):  
J. Dodds ◽  
M. Vahdati
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Numerical Predictions ◽  
Stage 1 ◽  
Variable Stator

In this two-part paper the phenomenon of part span rotating stall is studied. The objective is to improve understanding of the physics by which stable and persistent rotating stall occurs within high speed axial flow compressors. This phenomenon is studied both experimentally (Part I) and numerically (Part II). The experimental observations reported in Part I are now explored through the use of 3D unsteady Reynolds-averaged Navier–Stokes (RANS) simulation. The objective is to both validate the computational model and, where possible, explore some physical aspects of the phenomena. Unsteady simulations are presented, performed at a fixed speed with the three rows of variable stator vanes adjusted to deliberately mismatch the front stages and provoke stall. Two families of rotating stall are identified by the model, consistent with experimental observations from Part I. The first family of rotating stall originates from hub corner separations developing on the stage 1 stator vanes. These gradually coalesce into a multicell rotating stall pattern confined to the hub region of the stator and its downstream rotor. The second family originates from regions of blockage associated with tip clearance flow over the stage 1 rotor blade. These also coalesce into a multicell rotating stall pattern of shorter length scale confined to the leading edge tip region. Some features of each of these two patterns are then explored as the variable stator vanes (VSVs) are mismatched further, pushing each region deeper into stall. The numerical predictions show a credible match with the experimental findings of Part I. This suggests that a RANS modeling approach is sufficient to capture some important aspects of part span rotating stall behavior.

scholarly journals Influence of blade vibration on part-span rotating stall

2020 ◽  
Vol 4 ◽  
pp. 285-295
Author(s):  
Fanzhou Zhao ◽  
John Dodds ◽  
Mehdi Vahdati
Keyword(s):  
High Speed ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Blade Vibration ◽  
Natural Form ◽  
Multi Stage ◽  
Sensitivity Studies ◽  
Aeroelastic Simulations ◽  
Pattern Sensitivity ◽  
Lock In

This paper presents the interaction between blade vibration and part-span rotating stall in a multi-stage high speed compressor. Unsteady aerodynamic and aeroelastic simulations were conducted using URANS CFD. Steady state computations showed short length scale disturbances formed local to the tip of a front stage rotor. Using a full annulus model, these disturbances were shown to coalesce into flow structures rotating around the annulus at approximately 76% of the shaft rotational speed. Natural evolution of the rotating stall did not result in a coherent spatial pattern. Sensitivity studies showed that operating point and tip clearance have significant impact on the developed state of rotating stall. Subsequent analyses carried out with prescribed rotor blade vibration showed a spatial ‘lock-in’ event where the circumferential order of the part-span rotating stall shifted to match that induced by the vibration mode. Moreover, in contrast to its natural form in the absence of vibration, the fully developed rotating stall showed a coherent stall signal. More importantly, it was found that numerical boundary conditions such as mixing plane and sliding planes can significantly influence the outcome of prediction.

Rotating Stall Observations in a High Speed Compressor: Part 2 — Numerical Study

2014 ◽  
Author(s):  
J. Dodds ◽  
M. Vahdati
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Numerical Predictions ◽  
Stage 1 ◽  
Variable Stator

In this two part paper the phenomenon of part span rotating stall is studied. The objective is to improve understanding of the physics by which stable and persistent rotating stall occurs within high speed axial flow compressors. This phenomenon is studied both experimentally (part 1) and numerically (part 2). The experimental observations reported in Part 1 are now explored through the use of 3D unsteady RANS simulation. The objective is to both to validate the computational model and, where possible, explore some physical aspects of the phenomena. Unsteady simulations are presented, performed at a fixed speed with the three rows of variable stator stagger vanes adjusted to deliberately mismatch the front stages and provoke stall. Two families of rotating stall are identified by the model, consistent with experimental observations from Part 1. The first family of rotating stall originates from hub corner separations developing on the stage 1 stator vanes. These gradually coalesce into a multi-cell rotating stall pattern confined to the hub region of the stator and its downstream rotor. The second family originates from regions of blockage associated with tip clearance flow over the stage 1 rotor blade. These also coalesce into a multi-cell rotating stall pattern of shorter length scale confined to the leading edge tip region. Some features of each of these two patterns are then explored as the variable stator vanes are mismatched further, pushing each region deeper into stall. The numerical predictions show a credible match with the experimental findings of Part 1. This suggests that a RANS modelling approach is sufficient to capture some important aspects of part span rotating stall behavior.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

Stall Inception in a High Speed Centrifugal Compressor During Speed Transients

2017 ◽  
Author(s):  
Fangyuan Lou ◽  
John C. Fabian ◽  
Nicole L. Key
Keyword(s):  
Heat Transfer ◽  
Centrifugal Compressor ◽  
High Speed ◽  
Flow Instability ◽  
Leading Edge ◽  
Fast Response ◽  
Rotating Stall ◽  
Transient Performance ◽  
Stall Inception ◽  
Pressure Transducers

The inception and evolution of rotating stall in a high-speed centrifugal compressor are characterized during speed transients. Experiments were performed in the Single Stage Centrifugal Compressor (SSCC) facility at Purdue University and include speed transients from sub-idle to full speed at different throttle settings while collecting transient performance data. Results show a substantial difference in the compressor transient performance for accelerations versus decelerations. This difference is associated with the heat transfer between the flow and the hardware. The heat transfer from the hardware to the flow during the decelerations locates the compressor operating condition closer to the surge line and results in a significant reduction in surge margin during decelerations. Additionally, data were acquired from fast-response pressure transducers along the impeller shroud, in the vaneless space, and along the diffuser passages. Two different patterns of flow instabilities, including mild surge and short-length-scale rotating stall, are observed during the decelerations. The instability starts with a small pressure perturbation at the impeller leading edge and quickly develops into a single-lobe rotating stall burst. The stall cell propagates in the direction opposite of impeller rotation at approximately one third of the rotor speed. The rotating stall bursts are observed in both the impeller and diffuser, with the largest magnitudes near the diffuser throat. Furthermore, the flow instability develops into a continuous high frequency stall and remains in the fully developed stall condition.

Experiments on an Axial Fan Stage: Time-Resolved Analysis of Rotating Instability Modes

2015 ◽  
Vol 137 (6) ◽  
Author(s):  
Benjamin Pardowitz ◽  
Ulf Tapken ◽  
Lars Neuhaus ◽  
Lars Enghardt
Keyword(s):  
Leading Edge ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Axial Fan ◽  
Time Resolved ◽  
Instability Waves ◽  
Instability Modes ◽  
Mode Decomposition ◽  
Rotating Instability ◽  
Annular Cascade

Rotating instability (RI) occurs at off-design conditions in axial compressors, predominantly in rotor configurations with large tip clearances. Characteristic spectral signatures with side-by-side peaks below the blade passing frequency (BPF) are typically referred to RI located in the clearance region next to the leading edge (LE). Each peak can be assigned to a dominant circumferential mode. RI is the source of the clearance noise (CN) and an indicator for critical operating conditions. Earlier studies at an annular cascade pointed out that RI modes of different circumferential orders occur stochastically distributed in time and independently from each other, which is contradictory to existing explanations of RI. Purpose of the present study is to verify this generally with regard to axial rotor configurations. Experiments were conducted on a laboratory axial fan stage mainly using unsteady pressure measurements in a sensor ring near the rotor LE. A mode decomposition based on cross spectral matrices was used to analyze the spectral and modal RI patterns upstream of the rotor. Additionally, a time-resolved analysis based on a spatial discrete-Fourier-transform (DFT) was applied to clarify the temporal characteristics of the RI modes and their potential interrelations. The results and a comparison with the previous findings on the annular cascade corroborate a new hypothesis about the basic RI mechanism. This hypothesis implies that instability waves of different wavelengths are generated stochastically in a shear layer resulting from a backflow in the tip clearance region.

scholarly journals Tip Clearance Actuation With Magnetic Bearings for High-Speed Compressor Stall Control

2000 ◽  
Vol 123 (3) ◽  
pp. 464-472 ◽  
Author(s):  
Z. S. Spakovszky ◽  
J. D. Paduano ◽  
R. Larsonneur ◽  
A. Traxler ◽  
M. M. Bright
Keyword(s):  
High Speed ◽  
Design Procedure ◽  
Magnetic Bearing ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Magnetic Bearings ◽  
Test Facility ◽  
Control Analysis ◽  
Servo Actuator ◽  
Stall Control

Magnetic bearings are widely used as active suspension devices in rotating machinery, mainly for active vibration control purposes. The concept of active tip-clearance control suggests a new application of magnetic bearings as servo-actuators to stabilize rotating stall in axial compressors. This paper presents a first-of-a-kind feasibility study of an active stall control experiment with a magnetic bearing servo-actuator in the NASA Glenn high-speed single-stage compressor test facility. Together with CFD and experimental data a two-dimensional, incompressible compressor stability model was used in a stochastic estimation and control analysis to determine the required magnetic bearing performance for compressor stall control. The resulting requirements introduced new challenges to the magnetic bearing actuator design. A magnetic bearing servo-actuator was designed that fulfilled the performance specifications. Control laws were then developed to stabilize the compressor shaft. In a second control loop, a constant gain controller was implemented to stabilize rotating stall. A detailed closed loop simulation at 100 percent corrected design speed resulted in a 2.3 percent reduction of stalling mass flow, which is comparable to results obtained in the same compressor by Weigl et al. (1998. ASME J. Turbomach. 120, 625–636) using unsteady air injection. The design and simulation results presented here establish the viability of magnetic bearings for stall control in aero-engine high-speed compressors. Furthermore, the paper outlines a general design procedure to develop magnetic bearing servo-actuators for high-speed turbomachinery.

Fan Stability With Leading Edge Damage: Blind Prediction and Validation

2021 ◽  
Author(s):  
E. J. Gunn ◽  
T. Brandvik ◽  
M. J. Wilson ◽  
R. Maxwell
Keyword(s):  
Supersonic Flow ◽  
High Speed ◽  
Leading Edge ◽  
Rotating Stall ◽  
Stall Inception ◽  
Blind Prediction ◽  
Edge Damage ◽  
The Impact ◽  
Operating Points ◽  
Relative Flow

Abstract This paper considers the impact of a damaged leading edge on the stall margin and stall inception mechanisms of a transonic, low pressure ratio fan. The damage takes the form of a squared-off leading edge over the upper half of the blade. Full-annulus, unsteady CFD simulations are used to explain the stall inception mechanisms for the fan at low- and high-speed operating points. A combination of steady and unsteady simulations show that the fan is predicted to be sensitive to leading edge damage at low speed, but insensitive at high speed. This blind prediction aligns well with rig test data. The difference in response is shown to be caused by the change between subsonic and supersonic flow regimes at the leading edge. Where the inlet relative flow is subsonic, rotating stall is initiated by growth and propagation of a subsonic leading edge flow separation. This separation is shown to be triggered at higher mass flow rates when the leading edge is damaged, reducing the stable flow range. Where the inlet relative flow is supersonic, the flow undergoes a supersonic expansion around the leading edge, creating a supersonic flow patch terminated by a shock on the suction surface. Rotating stall is triggered by growth of this separation, which is insensitive to leading edge shape. This creates a marked difference in sensitivity to damage at low- and high-speed operating points.

Experimental Investigation of Tip Clearance Flow in a Transonic Compressor With and Without Plasma Actuators

2014 ◽  
Vol 137 (4) ◽  
Author(s):  
S. Saddoughi ◽  
G. Bennett ◽  
M. Boespflug ◽  
S. L. Puterbaugh ◽  
A. R. Wadia
Keyword(s):  
Flow Control ◽  
High Speed ◽  
Plasma Actuator ◽  
Tip Clearance ◽  
Plasma Actuators ◽  
Tip Clearance Flow ◽  
Transonic Compressor ◽  
Low Aspect Ratio ◽  
Clearance Flow ◽  
Design Speed

Blade tip losses represent a major performance penalty in low aspect ratio transonic compressors. This paper reports on the experimental evaluation of the impact of tip clearance with and without plasma actuator flow control on performance of an U.S. Air Force-designed low aspect ratio, high radius ratio single-stage transonic compressor rig. The detailed stage performance measurements without flow control at three clearance levels, classified as small, medium, and large, are presented. At design-speed, increasing the clearance from small to medium resulted in a stage peak efficiency drop of almost six points with another four point drop in efficiency with the large clearance (LC). Comparison of the speed lines at high-speed show significantly lower pressure rise with increasing tip clearance, the compressor losing 8% stall margin (SM) with medium clearance (MC) and an additional 1% with the LC. Comparison of the stage exit radial profiles of total pressure and adiabatic efficiency at both part-speed and design-speed and with throttling are presented. Tip clearance flow-control was investigated using dielectric barrier discharge (DBD) type plasma actuators. The plasma actuators were placed on the casing wall upstream of the rotor leading edge and the compressor mapped from part-speed to high-speed at three clearances with both axial and skewed configurations at six different frequency levels. The plasma actuators did not impact steady state performance. A maximum SM improvement of 4% was recorded in this test series. The LC configuration benefited the most with the plasma actuators. Increased voltage provided more SM improvement. Plasma actuator power requirements were almost halved going from continuous operation to pulsed plasma. Most of the improvement with the plasma actuators is attributed to the reduction in unsteadiness of the tip clearance vortex near-stall resulting in additional reduction in flow prior to stall.

Delay of Rotating Stall in Compressors Using Plasma Actuators

2016 ◽  
Vol 138 (9) ◽  
Author(s):  
Farzad Ashrafi ◽  
Mathias Michaud ◽  
Huu Duc Vo
Keyword(s):  
Centrifugal Compressor ◽  
Leading Edge ◽  
Plasma Actuator ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Compressor Stage ◽  
Dbd Plasma ◽  
Two Stage ◽  
Stall Inception ◽  
Flow Acceleration

Rotating stall is a well-known aerodynamic instability in compressors that limits the operating envelope of aircraft gas turbine engines. An innovative method for delaying the most common form of rotating stall inception using an annular dielectric barrier discharge (DBD) plasma actuator had been proposed. A DBD plasma actuator is a simple solid-state device that converts electricity directly into flow acceleration through partial air ionization. However, the proposed concept had only been preliminarily evaluated with numerical simulations on an isolated axial rotor using a relatively basic CFD code. This paper provides both an experimental and a numerical assessment of this concept for an axial compressor stage as well as a centrifugal compressor stage, with both stages being part of a low-speed two-stage axial-centrifugal compressor test rig. The two configurations studied are the two-stage configuration with a 100 mN/m annular casing plasma actuator placed just upstream of the axial rotor leading edge (LE) and the single-stage centrifugal compressor with the same actuator placed upstream of the impeller LE. The tested configurations were simulated with a commercial RANS CFD code (ansys cfx) in which was implemented the latest engineering DBD plasma model and dynamic throttle boundary condition, using single-passage multiple blade row computational domains. The computational fluid dynamics (CFD) simulations indicate that in both types of compressors, the actuator delays the stall inception by pushing the incoming/tip clearance flow interface downstream into the blade passage. In each case, the predicted reduction in stalling mass flow matches the experimental value reasonably well.

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