The Use of 3D Viscous Flow Calculations in the Design and Analysis of Industrial Centrifugal Compressors

1992 ◽  
Vol 114 (1) ◽  
pp. 27-37 ◽  
Author(s):  
M. V. Casey ◽  
P. Dalbert ◽  
P. Roth
Keyword(s):  
Flow Field ◽  
Surface Flow ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Centrifugal Impeller ◽  
Suction Surface ◽  
Meridional Velocity ◽  
Thin Region ◽  
And Performance ◽  
Operating Points

The 3D viscous codes of Denton and Dawes have been used to predict the flow field and performance of a back-swept industrial centrifugal impeller without an inducer (the Eckardt impeller “B”). The calculated flow field and performance agree very well with measurements at several operating points from surge to choke. Both codes predict that the suction surface flow near the impeller inlet is on the verge of separation at the design point. Calculations with tip clearance using the Dawes code predict a thin region of backflow in the clearance jet near the casing wall. The secondary flows arising from the meridional curvature in the impeller redistribute the losses generated on the blades and give rise to a spanwise stratification of the circumferentially mass-averaged losses. The resulting spanwise entropy gradient can be included in a throughflow calculation to improve the prediction of the meridional velocity distribution at impeller exit.

scholarly journals Laser Anemometer Measurements of the Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller

1997 ◽  
Author(s):  
G. J. Skoch ◽  
P. S. Prahst ◽  
M. P. Wernet ◽  
J. R. Wood ◽  
A. J. Strazisar
Keyword(s):  
Flow Field ◽  
Pressure Ratio ◽  
Tip Clearance ◽  
Centrifugal Impeller ◽  
Vaneless Diffuser ◽  
Suction Surface ◽  
Tip Clearance Flow ◽  
Through Flow ◽  
Clearance Flow ◽  
Flow Velocities

A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code. The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.

scholarly journals Prediction of 3-D Viscous Flow of Centrifugal Impeller With Tip-Clearance: Part 1 — High Speed Compressor

1997 ◽  
Author(s):  
E. Y. K. Ng
Keyword(s):  
Flow Field ◽  
Viscous Flow ◽  
High Speed ◽  
Flow Patterns ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Centrifugal Impeller ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Entire Flow

This paper described a 3-D, finite-volume procedure for simulating the entire viscous flow field within the impellers of a high subsonic radial compressor with uniform tip-clearance. The aim of the calculation is to verify the compressible Navier-Stokes solvers by comparing the computation against the experimental data of Krain backswept compressor at choke condition. The time-marching implicit technique used is a modified form of the basic Beam-Warming algorithm with a mixing length turbulence model in the blade-relative frame using rotating cylindrical coordinate system and appropriate source terms. An outline of the scheme is addressed and the present capabilities of the solver are assessed. The numerical calculations of the well-known jet/wake-type flow patterns agree well with measurement for the compressor rotor, and detailed comparisons indicate that the tip-clearance effect must be accurately modeled to compute the impeller flow patterns properly. In the case of larger separation region near suction surface of Krain compressor, it is due to the larger tip gap used in the model which causes the thickening of the casing boundary layer followed by the deterioration of the entire flow field.

Measurements of Secondary Losses in a Turbine Cascade With the Implementation of Nonaxisymmetric Endwall Contouring

2009 ◽  
Vol 132 (1) ◽  
Author(s):  
D. C. Knezevici ◽  
S. A. Sjolander ◽  
T. J. Praisner ◽  
E. Allen-Bradley ◽  
E. A. Grover
Keyword(s):  
Kinetic Energy ◽  
Axial Flow ◽  
Surface Flow ◽  
Secondary Flows ◽  
Test Facility ◽  
Suction Surface ◽  
Blade Passage ◽  
Low Pressure Turbine ◽  
Endwall Contouring ◽  
Pressure And Velocity Measurements

An approach to endwall contouring has been developed with the goal of reducing secondary losses in highly loaded axial flow turbines. The present paper describes an experimental assessment of the performance of the contouring approach implemented in a low-speed linear cascade test facility. The study examines the secondary flows of a cascade composed of Pratt & Whitney PAKB airfoils. This airfoil has been used extensively in low-pressure turbine research, and the present work adds intrapassage pressure and velocity measurements to the existing database. The cascade was tested at design incidence and at an inlet Reynolds number of 126,000 based on inlet midspan velocity and axial chord. Quantitative results include seven-hole pneumatic probe pressure measurements downstream of the cascade to assess blade row losses and detailed seven-hole probe measurements within the blade passage to track the progression of flow structures. Qualitative results take the form of oil surface flow visualization on the endwall and blade suction surface. The application of endwall contouring resulted in lower secondary losses and a reduction in secondary kinetic energy associated with pitchwise flow near the endwall and spanwise flow up the suction surface within the blade passage. The mechanism of loss reduction is discussed in regard to the reduction in secondary kinetic energy.

Effects of Circumferential Nonuniform Tip Clearance on Flow Field and Performance of a Transonic Turbine

2021 ◽  
Author(s):  
Yun Zheng ◽  
Xiubo Jin ◽  
Hui Yang ◽  
Qingzhe Gao ◽  
Kang Xu
Keyword(s):  
Flow Field ◽  
Tip Clearance ◽  
Transonic Turbine ◽  
And Performance

Effects of Circumferential Nonuniform Tip Clearance on Flow Field and Performance of a Transonic Turbine

2020 ◽  
Author(s):  
Yun Zheng ◽  
Xiubo Jin ◽  
Hui Yang ◽  
Qingzhe Gao ◽  
Kang Xu
Keyword(s):  
Flow Field ◽  
Numerical Study ◽  
Circumferential Direction ◽  
Tip Clearance ◽  
Computational Domain ◽  
Aerodynamic Efficiency ◽  
Aerodynamic Loss ◽  
The Difference ◽  
And Performance ◽  
Downstream Flow

Abstract The numerical study is performed by means of an in-house CFD code to investigate the effect of circumferential nonuniform tip clearance due to the casing ovalization on flow field and performance of a turbine stage. A method called fast-moving mesh is used to synchronize the non-circular computational domain with the rotation of the rotor row. Four different layouts of the circumferential nonuniform clearance are calculated and evaluated in this paper. The results show that, the circumferential nonuniform clearance could reduce the aerodynamic performance of the turbine. When the circumferential nonuniformity δ reaches 0.4, the aerodynamic efficiency decreases by 0.58 percentage points. Through the analysis of the flow field, it is found that the casing ovalization leads to the difference of the size of the tip clearance in the circumferential direction, and the aerodynamic loss of the position of large tip clearance is greater than that of small tip clearance, which is related to the scale of leakage vortex. In addition, the flow field will become nonuniform in the circumferential direction, especially at the rotor exit, which will adversely affect the downstream flow field.

Experimental Investigations of Tip Clearance Flow and its Influence on Secondary Flows in a 1-1/2 Stage Axial Turbine

2000 ◽  
Author(s):  
B. Stephan ◽  
H. E. Gallus ◽  
R. Niehuis
Keyword(s):  
Flow Field ◽  
Flow Behavior ◽  
Secondary Flows ◽  
Tip Clearance ◽  
Radial Clearance ◽  
Experimental Investigations ◽  
Axial Turbine ◽  
Tip Clearance Flow ◽  
Clearance Flow ◽  
Flow Phenomena

A multistage turbomachine has inherently unsteady flow fields due to the relative motion between rotor and stator airfoils, which lead to viscous and inviscid interactions between the blade rows. Additionally, the radial clearance between casing and rotor strongly influences the 3D flow field and the loss generation in turbomachines. The objective of the presented study is to investigate the effects of tip clearance on secondary flow phenomena and, in consequence, on the performance of a 1-1/2 stage axial turbine. The low aspect ratio of the blades and their prismatic design leads to a high degree of secondary flows and three-dimensionality. Extended measurements of the flow field behind each blade row with pneumatic and hotwire probes have been conducted for three different tip clearances. Experimental results reveal significant change of flow behavior and turbine performance with increasing tip clearance.

Influence of Swirl Distribution on Blade Shape and Performance of a Centrifugal Impeller

2020 ◽  
Author(s):  
Qin Cui ◽  
Guoliang Qin ◽  
Lei Li ◽  
Cheng Jia ◽  
Yi Wang
Keyword(s):  
Flow Field ◽  
Centrifugal Compressor ◽  
Flow Direction ◽  
Input Parameter ◽  
Vortex Method ◽  
Centrifugal Impeller ◽  
Streamline Curvature ◽  
Design Program ◽  
Blade Shape ◽  
And Performance

Abstract All-over-controlled vortex method is an effective tool to inversely design the 3D impeller of a centrifugal compressor. In this method, swirl distribution is treated as a significant input parameter to control the blade shape, impeller flow field, and compressor performance. It is acknowledged that swirl distribution is prescribed by designers mostly relying on the personal experience at the beginning of design. So how to specify the swirl distribution is still a big challenge for impeller designers. Of the most interest in this paper is to provide a theoretical technique that can be readily applied to specify swirl distribution and reduce the dependence on the designers experience. A judgement criterion rCθ – ωr2 is proposed to design the swirl distribution. Based on the streamline curvature method, a 3D centrifugal impeller design program is developed to design centrifugal impeller. The scale and uniformity of rCθ – ωr2 along flow direction are discussed theoretically to conduct the specifying of swirl distribution. The theoretical analysis is verified by a specific centrifugal compressor case. Then commercial CFD software is used to predict the flow field and the performance of the impeller. The results demonstrate that the scale and distribution uniformity of rCθ – ωr2 have a significant effect on the blade shape and the flow field within the impeller, and possible loss can be reduced. For the new designer, it is possible to preliminarily recognize and eliminate the infeasible swirl distribution, and adjust the unsatisfactory swirl distribution using rCθ – ωr2. Proper blade shape and good impeller performance can be achieved with the help of the judgement criterion rCθ – ωr2.

Fluid Dynamics and Performance of Partially and Fully Shrouded Axial Turbines

2004 ◽  
Author(s):  
L. Porreca ◽  
T. Behr ◽  
J. Schlienger ◽  
A. I. Kalfas ◽  
R. S. Abhari ◽  
...  
Keyword(s):  
Flow Field ◽  
Detailed Comparison ◽  
Labyrinth Seal ◽  
Tip Clearance ◽  
Field Analysis ◽  
Test Cases ◽  
Flow Measurements ◽  
And Performance ◽  
Partial Shroud ◽  
Stage Efficiency

A unique comparative experimental and numerical investigation carried out on two test cases with shroud configurations differing only in the labyrinth seal path, is presented in this paper. The blade geometry and tip clearance is identical in the two test cases. The geometries under investigation are representative of an axial turbine with a full and partial shroud, respectively. Global performance and flow field data were acquired and analyzed. Computational simulations were carried out to complement the investigation and to facilitate the analysis of the steady and unsteady flow measurements. A detailed comparison between the two test cases is presented in terms of flow field analysis and performance evaluation. The analysis focuses on the flow effects reflected on the overall performance in a multi-stage environment. Strong interaction between the cavity flow and the blade tip region of the rotor blades is observed up to the blade mid span. A marked effect of this interaction can be seen in the downstream second stator where different vortex structures are observed. Moreover, in the partial shroud test case, a strong tip leakage vortex is developed from the first rotor and transported through the downstream blade row. A measurable change in the second stage efficiency was observed between the two test cases. In low aspect ratio blades within a multistage environment, small changes in the cavity geometry can have a significant effect on the mainstream flow. The present analysis has shown that an integrated and matched blade-shroud aerodynamic design has to be adopted to reach optimal performances. The additional losses resulting from small variations of the sealing geometry could result in a gain of up to one point in the overall stage efficiency.

Investigations on a Radial Compressor Tandem-Rotor Stage With Adjustable Geometry

1992 ◽  
Author(s):  
B. Josuhn-Kadner ◽  
B. Hoffmann
Keyword(s):  
Flow Field ◽  
Navier Stokes ◽  
Compressor Stage ◽  
Radial Compressor ◽  
Design Point ◽  
Small Influence ◽  
Point Condition ◽  
And Performance ◽  
Operating Points

A radial compressor stage has been investigated mainly experimentally for aerodynamic stage optimization. The rotor (πt = 3.9) consists of a profiled axial inducer and a conventionally designed radial impeller. Inducer and impeller can be locked at different circumferential positions relative to each other thus, forming a tandem wheel with adjustable geometry. Conventional and Laser-2-Focus system measurements for the tandem-rotor and the stage were performed at different operating points to study the influence of the circumferential clearance geometry between inducer and impeller with respect to compressor characteristics and performance. Furthermore, 3-D Navier-Stokes calculations are being developed at design point condition to analyse the flow field. Small influence of the inducer adjustment on the rotor characteristics is observed. The maximum rotor efficiency of 93.5 % varies in a range of less than 1 % depending on the different inducer positions.

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