Effect of Inflow Circumferential Distortion on a Transonic Axial Compressor

2017 ◽  
Author(s):  
Yuyun Li ◽  
Zhiheng Wang ◽  
Guang Xi
Keyword(s):  
Total Pressure ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Leading Edge ◽  
Inlet Distortion ◽  
Compressor Rotor ◽  
Structure Changes ◽  
Transonic Axial Compressor

The Inlet distortion, which may lead to the stability reduction or structure failure, is often non-ignorable in an axial compressor. In the paper, the three-dimensional unsteady numerical simulations on the flow in NASA rotor 67 are carried out to investigate the effect of inlet distortion on the performance and flow structure in a transonic axial compressor rotor. A sinusoidal circumferential total pressure distortion with eleven periods per revolution is adopted to study the interaction between the transonic rotor and inlet circumferential distortion. Concerning the computational expense, the flow in two rotor blade passages is calculated. Various intensities of the total pressure distortion are discussed, and the detailed flow structures under different rotating speeds near the peak efficiency condition are analyzed. It is found that the distortion has a positive effect on the flow near the hub. Even though there is no apparent decrease in the rotor efficiency or total pressure ratio, an obvious periodic loading exists over the whole blade. The blade loadings are concentrated in the region near the leading edge of the rotor blade or regions affected by the oscillating shocks near the pressure side. The time averaged location of shock structure changes little with the distortion, and the motion of shocks and the interactions between the shock and the boundary layer make a great contribution to the instability of the blade structure.

A Comparison between Measured and Computed Flow Fields in a Transonic Compressor Rotor

1980 ◽  
Vol 102 (4) ◽  
pp. 883-889 ◽  
Author(s):  
P. W. McDonald ◽  
C. R. Bolt ◽  
R. J. Dunker ◽  
H. B. Weyer
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Computational Procedure ◽  
Operating Conditions ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Pressure Profiles

The flow field within the rotor of a transonic axial compressor has been computed and compared to measurements obtained with an advanced laser velocimeter. The compressor was designed for a total pressure ratio of 1.51 at a relative tip Mach number of 1.4. The comparisons are made at 100 percent design speed (20,260 RPM) with pressure ratios corresponding to peak efficiency, near surge, and wide open discharge operating conditions. The computational procedure iterates between a blade-to-blade calculation and an intrablade through flow calculation. Calculated Mach number contours, surface pressure distributions, and exit total pressure profiles are in agreement with the experimental data demonstrating the usefulness of quasi three-dimensional calculations in compressor design.

Effect of Trailing Edge Crenulation on the Performance and Stall Margin of a Transonic Axial Compressor

2013 ◽  
Author(s):  
S. Subbaramu ◽  
Quamber H. Nagpurwala ◽  
A. T. Sriram
Keyword(s):  
Beneficial Effect ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compressor Cascade ◽  
Trailing Edge ◽  
Stall Margin ◽  
Positive Effects ◽  
Compressor Rotor ◽  
Transonic Axial Compressor

This paper deals with the numerical investigations on the effect of trailing edge crenulation on the performance of a transonic axial compressor rotor. Crenulation is broadly considered as a series of small notches or slots at the edge of a thin object, like a plate. Incorporating such notches at the trailing edge of a compressor cascade has shown beneficial effect in terms of reduction in total pressure loss due to enhanced mixing in the wake region. These notches act as vortex generators to produce counter rotating vortices, which increase intermixing between the free stream flow and the low momentum wake fluid. Considering the positive effects of crenulation in a cascade, it was hypothesized that the same technique would work in a rotating compressor to enhance its performance and stall margin. However, the present CFD simulations on a transonic compressor rotor have given mixed results. Whereas the peak total pressure ratio in the presence of trailing edge crenulation reduced, the stall margin improved by 2.97% compared to the rotor with straight edge blades. The vortex generation at the crenulated trailing edge was not as strong as reported in case of linear compressor cascade, but it was able to influence the flow field in the rotor tip region so as to energize the low momentum end-wall flow in the aft part of the blade passage. This beneficial effect delayed flow separation and allowed the mass flow rate to be reduced to still lower levels resulting in improved stall margin. The reduction in pressure ratio with crenulation was surprising and might be due to increased mixing losses downstream of the blade.

Numerical Studies on the Effect of Gurney Flap on Aerodynamic Performance and Stall Margin of a Transonic Axial Compressor Rotor

2014 ◽  
Author(s):  
Mudassir Ahmed M. Rafeeq ◽  
Quamber H. Nagpurwala ◽  
Subbaramu Shivaramaiah
Keyword(s):  
Total Pressure ◽  
Rotor Blade ◽  
Peak Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Numerical Studies ◽  
Gurney Flap ◽  
Total Pressure Ratio

Numerical studies have been carried out on the effectiveness of trailing edge Gurney flap on a transonic axial compressor rotor. The baseline geometry of the rotor blade was modified at the trailing edge by introducing Gurney flaps of varying depth and span-wise length, viz. 1 mm, 2 mm and 3 mm depth with 20% span length of Gurney flap from tip (designated as GF1-20, GF2-20 and GF3-20 respectively), and 1 mm depth with 50% and 100% span length (designated as GF1-50 and GF1-100 respectively). Geometric models of the compressor rotor without and with Gurney flaps were generated using CATIA V5 software and CFD simulations at 100% design rotor speed were carried out using ANSYS CFX software. Results have shown that the compressor total pressure ratio increased with increase in both depth and spanwise length of Gurney flap. Peak pressure ratio increased from 1.51 for baseline case to 1.58 for rotor GF1-100. However, the peak isentropic efficiency remained almost constant for various Gurney flap configurations, except for GF1-100 which showed a tendency for improvement in efficiency. The stall margin reduced with the introduction of Gurney flap and was lowest for configuration GF1-100 which gave highest peak pressure ratio. Higher blade loading with Gurney flap was responsible for lowering the stall margin. Analysis of the flow through the blade passages has shown clear formation of trailing end vortex structure in the presence of Gurney flap that resulted in bending of the streamlines towards suction surface of the rotor blade, with consequent reduction in flow deviation and increased flow deflection, and hence increased total pressure ratio.

Experimental Study of the Flow Field Within a Transonic Axial Compressor Rotor by Laser Velocimetry and Comparison With Through-Flow Calculations

1978 ◽  
Vol 100 (2) ◽  
pp. 279-286 ◽  
Author(s):  
R. J. Dunker ◽  
P. E. Strinning ◽  
H. B. Weyer
Keyword(s):  
Flow Field ◽  
Pressure Ratio ◽  
Analytical Data ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Compressor Rotor ◽  
Through Flow ◽  
Equivalent Mass ◽  
Transonic Axial Compressor ◽  
Design Speed

The flow field ahead, within, and behind the rotor of a transonic axial compressor designed for a total pressure ratio of 1.51 at a relative tip Mach number of 1.4 has been studied in detail using an advanced laser velocimeter. The tests were carried out at 70 and 100 percent design speed (20,260 rpm) and equivalent mass flows corresponding to the point of maximum isentropic efficiency. The tests yielded quite complete data on the span- and gap-wise velocity profiles, on the three-dimensional shock waves in and outside of the rotor blade channels, and on the blade wakes. Some of the experimental results will be submitted, discussed, and compared to corresponding analytical data of a through-flow calculation. The comparison reveals considerable discrepancies inside the blade row between the two-dimensional calculation and the experiments primarily due to the loss and deviation correlations used, as well as to the distribution of losses and flow angles inside the blade channels.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

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.

An Explanation for Flow Features of Spike-Type Stall Inception in an Axial Compressor Rotor

2012 ◽  
Author(s):  
Kazutoyo Yamada ◽  
Hiroaki Kikuta ◽  
Ken-ichiro Iwakiri ◽  
Masato Furukawa ◽  
Satoshi Gunjishima
Keyword(s):  
Flow Structure ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Leading Edge ◽  
Low Pressure ◽  
Stall Inception ◽  
Compressor Rotor ◽  
Pressure Region ◽  
Casing Pressure ◽  
Edge Separation

The unsteady behavior and three-dimensional flow structure of spike-type stall inception in an axial compressor rotor have been investigated by experimental and numerical analyses. Previous studies have revealed that the test compressor falls into a mild stall after emergence of a spike, in which multiple stall cells, each consisting of a tornado-like vortex, are rotating. However, the flow mechanism from the spike onset to the mild stall remains unexplained. The purpose of this study is to describe the flow mechanism of a spike stall inception in a compressor. In order to capture the transient phenomena of spike-type stall inception experimentally, an instantaneous casing pressure field measurement technique was developed, in which 30 pressure transducers measure an instantaneous casing pressure distribution inside the passage for one blade pitch at a rate of 25 samplings per blade passing period. This technique was applied to obtain the unsteady and transient pressure fields on the casing wall during the inception process of the spike stall. In addition, the details of the three-dimensional flow structure at the spike stall inception have been analyzed by a numerical approach using the detached-eddy simulation (DES). The instantaneous casing pressure field measurement results at the stall inception show that a low-pressure region starts traveling near the leading edge in the circumferential direction just after the spiky wave was detected in the casing wall pressure trace measured near the rotor leading edge. The DES results reveal the vortical flow structure behind the low-pressure region on the casing wall at the stall inception, showing that the low-pressure region is caused by a tornado-like separation vortex resulting from a leading-edge separation near the rotor tip. A leading-edge separation occurs near the tip at the onset of the spike stall and grows to form the tornado-like vortex connecting the blade suction surface and the casing wall. The casing-side leg of the tornado-like vortex generating the low-pressure region circumferentially moves around the leading-edge line. When the vortex grows large enough to interact with the leading edge of the next blade, the leading-edge separation begins to propagate, and then, the compressor falls into a stall with decreasing performance.

Design and Test of a Transonic Axial Splittered Rotor

2015 ◽  
Author(s):  
Garth V. Hobson ◽  
Anthony J. Gannon ◽  
Scott Drayton
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Procedure ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Transonic Axial Compressor ◽  
Commercial Off The Shelf

A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected in a Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

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