Stall inception mechanisms in a contra-rotating fan operating at different speed combinations

2019 ◽  
Vol 234 (8) ◽  
pp. 1041-1052 ◽  
Author(s):  
MP Manas ◽  
AM Pradeep
Keyword(s):  
Pressure Ratio ◽  
Pressure Sensors ◽  
Stability Limit ◽  
Leading Edge ◽  
Axial Direction ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Unsteady Pressure ◽  
Stall Cells ◽  
The Stability

Contra-rotating fan is a concept that can possibly replace the present-day conventional fans due to its several aerodynamic advantages. It has the potential to improve the stability limit and can achieve a higher pressure ratio per stage. One of the advantages of a contra-rotating fan is its capability to operate both the rotors at different speeds. In the present study, experiments are carried out at different speed combinations of the rotors and the stall inception phenomenon is captured using high-response unsteady pressure sensors placed on the casing upstream of the leading edge of rotor-1. The unsteady pressure data are investigated using wavelet and Fourier analysis techniques. It is observed that the mechanism of stall inception is different for different speed combinations. The pre-stall disturbances fall in different frequency ranges for different speed combinations. For the range of speed combinations investigated, the frequency of appearance of stall cells of rotor-1 does not depend on the speed of rotor-2. A higher speed of rotation of rotor-1 leads to a higher frequency of appearance of stall cells and a lower speed of rotation of rotor-1 leads to a lower frequency of appearance of stall cells. For all the speed combinations, there is a range of frequency where no disturbance is observed and this range is termed as the ‘no-disturbance zone’. Disturbances are observed at lower frequencies and at frequencies close to the blade passing frequency. In order to understand the flow physics in detail, computational analysis is carried out for different speed combinations of the rotors. For a higher speed of rotor-2, it is observed that the suction effect of rotor-2 is significant enough to pull the tip-leakage flow towards the axial direction. Thus, the suction effect of rotor-2 plays a significant role in determining the stall of the stage.

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.

Investigation on the Stall Inception Circumferential Position and Stall Process Behavior in a Centrifugal Compressor With Volute

2018 ◽  
Author(s):  
Hanzhi Zhang ◽  
Ce Yang ◽  
Dengfeng Yang ◽  
Wenli Wang ◽  
Changmao Yang ◽  
...  
Keyword(s):  
Centrifugal Compressor ◽  
Static Pressure ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Tip Clearance ◽  
Positive Pressure ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Pressure Distributions ◽  
Impeller Outlet

The present paper numerically and experimentally investigates the stall inception mechanisms in a centrifugal compressor with volute. Current studies about stall inception pay more attention on the axial compressors than the centrifugal compressors; especially, the circumferential position of stall inception onset and the stall process in the centrifugal compressor with asymmetric volute structure have not been studied sufficiently yet. In this work, the compressor performance experiment was conducted and the casing wall static pressure distributions were obtained by seventy-two static pressure sensors firstly. Then, the full annular unsteady simulations were carried out at different stable operating points, and the time-averaged static pressure distributions were compared with the experimental results. Finally, the stall process of the compressor was investigated by unsteady simulations in detail. Results show that the stall inception onset is determined by the impeller leading edge spillage flow, and the occurrence time of trailing edge backflow is prior to the leading edge spillage. The non-uniform static pressure circumferential distribution at impeller outlet induced by volute tongue causes the two stall inception regions locating at certain circumferential positions, which are 120° and 300° circumferential positions at impeller leading edge, corresponding to the circumferential static pressure peak and bulge regions at impeller outlet, respectively. In detail, at rotor revolution 2.86, a small disturbance that the incoming/tip clearance flow interface is perpendicular to axial direction occurs at 120° position, but this disturbance did not cause the compressor stall. Then at revolution 7, the first stall inception zone (spillage region) occurs at 120° position, causing the compressor stall with positive pressure ratio performance. At approximately revolution 23, the second stall inception zone occurs at about 300° position; however, both the intensity and size of this stall inception zone are smaller than those of the first stall inception zone. These two stall inception zones are not moving along circumferential direction because the stall inception circumferential position is dominated by the impeller outlet static pressure distribution. Even that, the obvious low frequency signals appear after the spillage crossing two blade leading edges; because at this moment, the spillage vortex caused by the tip leakage flow begins to shed. However, due to the asymmetric structure limitation, this vortex cannot move across full annular. Furthermore, the spillage vortexes cause the local low static pressure zone ahead of blade leading edge in the centrifugal compressor with volute, suggesting that the spillage can be predicted by the steady casing wall static pressure measuring. The development of blockage zones at impeller leading edge is also investigated quantitatively by analyzing the stall blockage effect.

scholarly journals Axial Flow Compressor Stability Enhancement: Circumferential T-Shape Grooves Performance Investigation

Aerospace ◽  
2021 ◽  
Vol 8 (1) ◽  
pp. 12
Author(s):  
Marco Porro ◽  
Richard Jefferson-Loveday ◽  
Ernesto Benini
Keyword(s):  
Vortex Breakdown ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Suction Side ◽  
Treatment Technique ◽  
Tip Leakage Flow ◽  
Stall Inception ◽  
Stall Margin ◽  
Casing Treatment ◽  
Stall Margin Improvement

This work focuses its attention on possibilities to enhance the stability of an axial compressor using a casing treatment technique. Circumferential grooves machined into the case are considered and their performances evaluated using three-dimensional steady state computational simulations. The effects of rectangular and new T-shape grooves on NASA Rotor 37 performances are investigated, resolving in detail the flow field near the blade tip in order to understand the stall inception delay mechanism produced by the casing treatment. First, a validation of the computational model was carried out analysing a smooth wall case without grooves. The comparisons of the total pressure ratio, total temperature ratio and adiabatic efficiency profiles with experimental data highlighted the accuracy and validity of the model. Then, the results for a rectangular groove chosen as the baseline case demonstrated that the groove interacts with the tip leakage flow, weakening the vortex breakdown and reducing the separation at the blade suction side. These effects delay stall inception, improving compressor stability. New T-shape grooves were designed keeping the volume as a constant parameter and their performances were evaluated in terms of stall margin improvement and efficiency variation. All the configurations showed a common efficiency loss near the peak condition and some of them revealed a stall margin improvement with respect to the baseline. Due to their reduced depth, these new configurations are interesting because they enable the use of a thinner light-weight compressor case as is desirable in aerospace applications.

A Study of Influences of Inlet Total Pressure Distortions On Clearance Flow in an Axial Compressor

2021 ◽  
Author(s):  
Guoming Zhu ◽  
Xiaolan Liu ◽  
Bo Yang ◽  
Moru Song
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Stall Margin ◽  
Flow Cells ◽  
Static Pressure Distribution ◽  
Speed Flow ◽  
Clearance Flow ◽  
The Stability

Abstract The rotating distortion generated by upstream wakes or low speed flow cells is a kind of phenomenon in the inlet of middle and rear stages of an axial compressor. Highly complex inflow can obviously affect the performance and the stability of these stages, and is needed to be considered during compressor design. In this paper, a series of unsteady computational fluid dynamics (CFD) simulations is conducted based on a model of an 1-1/2 stage axial compressor to investigate the effects of the distorted inflows near the casing on the compressor performance and the clearance flow. Detailed analysis of the flow field has been performed and interesting results are concluded. The distortions, such as total pressure distortion in circumferential and radial directions, can block the tip region so that the separation loss and the mixing loss in this area are increased, and the efficiency and the total pressure ratio are dropped correspondingly. Besides, the distortions can change the static pressure distribution near the leading edge of the rotor, and make the clearance flow spill out of the rotor edge more easily under near stall condition, especially in the cases with co-rotating distortions. This phenomenon can be used to explain why the stall margin is deteriorated with nonuniform inflows.

Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe
Keyword(s):  
Pressure Ratio ◽  
Research Work ◽  
Vortex Generator ◽  
Leading Edge ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor

Abstract In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Deficiencies in Turbulence Modelling for the Prediction of the Stability Boundary in Highly Loaded Compressors

2019 ◽  
Author(s):  
Jose Moreno ◽  
John Dodds ◽  
Mehdi Vahdati ◽  
Sina Stapelfeldt
Keyword(s):  
High Speed ◽  
Stability Boundary ◽  
Turbulence Models ◽  
Stability Limit ◽  
Major Effect ◽  
Navier Stokes ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Multi Stage ◽  
The Stability

Abstract Reynolds-averaged Navier-Stokes (RANS) equations are employed for aerodynamic and aeroelastic modelling in axial compressors. Their solutions are highly dependent on the turbulence models for closure. The main objective of this work is to assess the widely used Spalart-Allmaras model’s suitability for compressor flows. For this purpose, an extensive investigation of the sources of uncertainties in a high-speed multi-stage compressor rig was carried out. The grid resolution near the casing end wall, which affects the tip leakage flow and casing boundary layer, was found to have a major effect on the stability limit prediction. Refinements in this region led to a stall margin loss prediction. It was found that this loss was exclusively due to the destruction term in the SA model.

Breakdown of Tip Leakage Vortices in Compressors at Flow Conditions Close to Stall

1997 ◽  
Author(s):  
Stefan Schlechtriem ◽  
Michael Lötzerich
Keyword(s):  
Vortex Breakdown ◽  
Stability Limit ◽  
Flow Conditions ◽  
Mixed Flow ◽  
Stall Inception ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Flow Compressor ◽  
The Stability ◽  
Operating Points

The breakdown of tip leakage vortices at operating points close to the stability limit of transonic compressor rotors has been detected. The aerodynamic phenomenon is considered to have a major impact on stall inception. Computations have been carried out and a detailed visualization of the phenomenon is given. In addition the connection of vortex breakdown to rotating instabilities and stall is discussed. Furthermore the tip flow field of the axial rotor is compared to the results for a centrifugal and a mixed flow compressor operating at similar tip speeds.

Study of the Tip Leakage Flow in a Compressor Cascade With Incidence Exceeding the Stability Limit

2019 ◽  
Author(s):  
Markus W. Leitner ◽  
Stephan Staudacher ◽  
Martin G. Rose
Keyword(s):  
Water Table ◽  
Vortex Breakdown ◽  
Stability Limit ◽  
Leakage Flow ◽  
Angle Of Incidence ◽  
Compressor Cascade ◽  
Tip Leakage Flow ◽  
Unstable Flow ◽  
Tip Leakage ◽  
The Stability

Abstract In axial compressors, tip leakage flow is disadvantageous to efficiency and mass flow stability. We analyzed the tip leakage flow in a compressor cascade on a water table at various angles of incidence. When the angle of incidence is systematically increased, the flow rate is decreased and, finally, the stability limit is exceeded. To study the flow structures and vortex behavior, we installed Particle Tracking Velocimetry (PTV) on the water table. 3D-trajectories of the stable and unstable flow reveal significant effects. Increasing incidence generates a significant change in the nature of the flow. The tip leakage flow fluctuates and features unstable flow phenomena. A large blockage of the flow passage occurs, probably due vortex breakdown. Such a serious disturbance of the incoming flow may induce stall.

Investigation of Tip-Flow Based Stall Criteria Using Rotor Casing Visualization

2008 ◽  
Author(s):  
Matthew A. Bennington ◽  
Joshua D. Cameron ◽  
Scott C. Morris ◽  
Camille Legault ◽  
Sean T. Barrows ◽  
...  
Keyword(s):  
Mass Flow ◽  
Flow Direction ◽  
Leading Edge ◽  
Axial Direction ◽  
Tip Clearance ◽  
Flow Coefficient ◽  
Momentum Balance ◽  
Reverse Direction ◽  
Axial Component ◽  
Stall Inception

Recent stall inception investigations have indicated that short-length scale stall initiates when the interface between the tip gap flow and the approach flow spills forward of the leading edge of the adjacent blade. This hypothesis was investigated in the present work using both numerical and experimental results from a range of compressor geometries and speed. First, full annulus unsteady computations of R35 were used to generate contours of entropy at the casing. It was found that a large gradient in entropy, which marked the leakage fluid, was aligned with the leading edge plane at the stalling mass flow. It was also observed that the flow direction in the region of increased entropy was in the reverse axial direction. The interface between the approach fluid and the reverse-direction leakage flow was related to a region in which the axial component of the wall shear stress was zero. The axial location of this line was measured experimentally using a surface streaking method using two separate facilities. It was found that the location of this line is determined by a momentum balance between the approach fluid and the tip leakage fluid. Measurements were acquired with varied tip clearance, radial distortion, and centerline offset to support these conclusions. In all cases the zero axial shear line was found to move upstream with decreased flow coefficient, and was in close proximity to the rotor leading edge at the stalling mass flow.

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