Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1987 ◽  
Vol 109 (1) ◽  
pp. 142-148 ◽  
Author(s):  
C. H. Law ◽  
A. J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

scholarly journals Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1986 ◽  
Author(s):  
C. Herbert Law ◽  
Arthur J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performances of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

1987 ◽  
Vol 109 (3) ◽  
pp. 354-361 ◽  
Author(s):  
Y. Dong ◽  
S. J. Gallimore ◽  
H. P. Hodson
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Tracer Gas ◽  
Suction Surface ◽  
Axial Compressors ◽  
Small Clearance ◽  
Oil Flow ◽  
Stator Blade ◽  
Flow Compressor

Measurements have been performed in a low-speed high-reaction single-stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface–hub corner separations, their associated loss mechanisms, and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualizaton indicated that the leakage reduced the extensive suction surface–hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

Three-Dimensional Flows and Loss Reduction in Axial Compressors

1986 ◽  
Author(s):  
Y. Dong ◽  
S. J. Gallimore ◽  
H. P. Hodson
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Axial Flow ◽  
Flow Visualisation ◽  
Tracer Gas ◽  
Suction Surface ◽  
Axial Compressors ◽  
Oil Flow ◽  
Stator Blade ◽  
Flow Compressor

Measurements have been performed in a low speed high reaction single stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface-hub corner separations, their associated loss mechanisms and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualisation indicated that the leakage reduced the extensive suction surface-hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.

Wakes in Axial Compressors

1959 ◽  
Vol 63 (583) ◽  
pp. 415-416 ◽  
Author(s):  
H. Pearson ◽  
A. B. McKenzie
Keyword(s):  
Static Pressure ◽  
Early Stage ◽  
Axial Compressor ◽  
Pressure Gradients ◽  
Blade Vibration ◽  
Axial Compressors ◽  
Pressure Distributions ◽  
The People ◽  
Total Head ◽  
Flow Test

The tendency in the past has been to assume that when wakes or non-uniform total head profiles are fed into an axial compressor then substantially constant static pressure prevails at the entry, the variations in total head appearing as variations in velocity. This variation in velocity causes variation in incidence on the early stage blade rows and thus can give rise to excitation of blade vibration. This assumption is implicit, for instance, in References 1 and 2, but we think has been a common assumption by most of the people working in this field.Where the compressor is fed by a duct of substantially parallel walls for a reasonable length ahead, such an assumption appeared justifiable. Such a duct when given an air flow test with its outlet discharging, for instance, to atmosphere instead of to the compressor, then the distribution assumed would normally be obtained and in fact many surveys of such ducts have been represented in this fashion. The object of this note is to show that, in fact, this distribution will not normally occur when the compressor is present and we may normally expect much more nearly a constant velocity into the compressor with attendant static pressure distributions to match with the total head variations ahead of the intake, with of course, the attendant curved flow to support the static pressure gradients.

Experimental Investigation of Effects of Leading-Edge Tubercles on Compressor Cascade Performance

2014 ◽  
Author(s):  
M. C. Keerthi ◽  
Abhijit Kushari ◽  
Ashoke De ◽  
Arun Kumar
Keyword(s):  
Incidence Angle ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Compressor Cascade ◽  
Compressor Blades ◽  
Stall Margin ◽  
Axial Compressors ◽  
Pressure Loss Coefficient ◽  
Lift And Drag

In the present study, the effectiveness of passive structures called tubercles on an axial compressor blade row is studied experimentally. Tubercles are the modifications to the leading edge of an airfoil in the form of blunt wave-like serrations. Although several studies on the effect of tubercles on isolated blades are available in literature, detailed study of their effect on a cascade of blades, such as in the case of an axial flow turbo-machine is lacking. Such an application in an axial compressor will result in a significant increase in the stall margin. Presently, experiments have been performed on a linear compressor cascade with a blade height of 0.15 m and mean chord of 0.06 m, on a NACA 65209 airfoil profile. The plain and modified blades are fabricated using rapid prototyping to ensure conformity to the required geometry. The cascade is designed in such a way that the incidence (angle of attack) and the stagger can be changed easily. The measurements are taken at the exit plane using a five-hole Pitot probe to obtain three-components of velocity and static pressure data over fine measurement grids. The effect is determined in terms of lift and drag coefficients, lift-to-drag ratio and total pressure loss coefficient. Experiments have been carried out for different pitch and amplitude (serration depth) of tubercles to understand their effect. The stall incidence angle for the best performing blade is found to increase up to 8.6° from that of the unmodified blade of 6.0°. Application of such structures in axial compressor blades may well be adequate to prevent stalling in axial compressors over a wide operating range.

Unsteady and Three-Dimensional Flow Mechanism of Spike-Type Stall Inception in an Axial Flow Compressor Rotor

2011 ◽  
Author(s):  
Hiroaki Kikuta ◽  
Ken-ichiro Iwakiri ◽  
Masato Furukawa ◽  
Kazutoyo Yamada ◽  
Satoshi Gunjishima ◽  
...  
Keyword(s):  
Three Dimensional ◽  
Axial Flow ◽  
Leading Edge ◽  
Three Dimensional Flow ◽  
Stall Inception ◽  
Dimensional Flow ◽  
Compressor Rotor ◽  
Pressure Distributions ◽  
Blade Tip ◽  
Flow Compressor

The unsteady behaviors and three-dimensional flow structure of the spike-type stall inception in an axial flow compressor rotor have been investigated by experimental and numerical analyses. In order to capture the transient phenomena of spike-type stall inception experimentally, “SFMT (Simultaneous Field Measurement Technique)”, by which instantaneous pressure distributions on the casing wall were acquired, was developed. By applying this technique, the unsteady flow pattern on the casing wall was extracted for each phase of development process of the stall inception. The details of three-dimensional flow structure in the stall inception process were revealed by the numerical analysis using a detached-eddy simulation (DES). At the stall inception, the characteristic patterns of the casing wall pressure distributions are observed in the experimental results: the low pressure regions moving in the circumferential direction and the variations of the low pressure regions at the leading edge. Considering the results of DES, these patterns are made by the vortices fragmented from the deformed tip leakage vortex or the tornado-type separation vortex and also are made by the tornado-type separation vortex itself, as well. The vortical flow structures have been elucidated. These vortices actually result from the leading edge separation at the blade tip. Therefore, it has been found that spike-type stall inception is dominated by the leading edge separation at the rotor blade tip.

The Flow Mechanism of How Distorted Flows Deteriorate Stability of an Axial Flow Compressor

2007 ◽  
Author(s):  
J. Zhang ◽  
F. Lin ◽  
J. Chen ◽  
C. Nie
Keyword(s):  
Cfd Simulation ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Resonance Phenomenon ◽  
Flow Mechanism ◽  
Stall Inception ◽  
Blade Passage ◽  
Inlet Distortion ◽  
Flow Compressor

The stalling behavior in a single-stage low-speed axial compressor under inlet distortion is investigated. A blade-passage-scale flow mechanism is proposed to explain the stability deterioration caused by inlet distortion for the tested compressor exhibiting spike stall inception. In contrast to the existing understanding of inlet distortion based on system scale dynamics, the main elements of this flow mechanism are the unsteady behavior of tip leakage vortices (TLV) under inlet distortion; its effect on the initiation of spike flow disturbances, and its interaction with distorted sectors. Rotating inlet distortion (RID) is used as a tool because RID makes it possible to directly compare the flows between distorted and clean flow sectors with fixed measurement stations on the casing, and the fact that the stationary inlet distortion is only a special case of RID makes the results generic. The tests demonstrate that the blade loading in the distorted sector is heavier than that in the non-distorted sector, causing the TLV in the distorted sector move closer to the leading edge of the rotor blade and thus be the first to initiate the spike-like disturbance. The unsteady CFD simulation further confirms that such a disturbance corresponds to a vortex spinning out of the leading edge of the blades. However, the initiation of this spike-like disturbance doesn’t necessarily trigger the full stall immediately. The tracking of the disturbances indicates that most of such spike-like disturbances will be smeared by non-distorted sector and the growth of the spike-like disturbances actually relate closely to how and how often the path of the propagating disturbances come across the path of the rotating distorted sector. The proposed blade-passage-scale flow mechanism also offers an alternative explanation to the “resonance” phenomenon in rotating inlet distortion research, which was explained with excitation-and-response theory for compressors that exhibit modal stall inception.

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.

Numerical Efforts of Aerodynamic Re-Design in a Single-Stage Transonic Axial Compressor: Part 1 — Stator Design

2017 ◽  
Author(s):  
Justin (Jongsik) Oh
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Axial Flow ◽  
Single Stage ◽  
Design Flow ◽  
Design Parameters ◽  
Original Design ◽  
Circular Arc ◽  
Surge Margin ◽  
Flow Compressor

In many aerodynamic design parameters for the axial-flow compressor, three variables of tailored blading, blade lean and sweep were considered in the re-design efforts of a transonic single stage which had been designed in 1960’s NASA public domains. As Part 1, the re-design was limited to the stator vane only. For the original MCA (Multiple Circular Arc) blading, which had been applied at all radii, the CDA (Controlled Diffusion Airfoil) blading was introduced at midspan as the first variant, and the endwalls of hub and casing (or tip) were replaced with the DCA (Double Circular Arc) blading for the second variant. Aerodynamic performance was predicted through a series of CFD analysis at design speed, and the best aerodynamic improvement, in terms of pressure ratio/efficiency and operability, was found in the first variant of tailored blading. It was selected as a baseline for the next design efforts with blade lean, sweep and both combined. Among 12 variants, a case of positively and mildly leaned blades was found the most attractive one, relative to the original design, providing benefits of an 1.0% increase of pressure ratio at design flow, an 1.7% increase of efficiency at design flow, a 10.5% increase of the surge margin and a 32.3% increase of the choke margin.

scholarly journals A Comparison of Calculated Flows in Radial Turbines With Experiment

1972 ◽  
Author(s):  
W. G. Cartwright
Keyword(s):  
Velocity Profile ◽  
Static Pressure ◽  
The Other ◽  
Peak Efficiency ◽  
Exit Velocity ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Angle Of Deviation ◽  
Radial Turbines

The flow in the rotors of three radial turbines, of differing peak efficiency, is analyzed using a streamline curvature method. The turbine of greatest efficiency is analyzed at both on- and off-design conditions; the other two turbines at the design point only. Comparison is made between the predictions of the calculation and the experimental determination of two features of the flow — the shroud static pressure distribution and the outlet velocity profile. Fair agreement with the shroud pressure is obtained at on-design conditions, but correlation with the exit velocity distribution is poor. Some improvement in the calculation of the exit profile is achieved when the analysis is modifed so as to allow for the experimentally observed angle of deviation at the blade trailing edge. Consideration is given to the ability of the analytical method to discriminate between turbines which prove experimentally to have high or low peak efficiency.

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