Axial Flow Compressor and Turbine Loss Coefficients: A Comparison of Several Parameters

1972 ◽  
Vol 94 (3) ◽  
pp. 193-201 ◽  
Author(s):  
L. E. Brown
Keyword(s):  
Mach Number ◽  
Test Data ◽  
Axial Flow ◽  
Loss Coefficient ◽  
Low Mach Number ◽  
Loss Parameter ◽  
Loss Coefficients ◽  
Flow Compressor ◽  
Stage Efficiency ◽  
Better Than

Many loss parameters are used in the turbomachinery field for correlating the effects on losses of numerous geometric and aerodynamic variables associated with blade rows. The parameter most common to these correlations is the ratio of a loss parameter to a velocity parameter, here called the loss coefficient. Such loss coefficients of different forms used for compressors by Howell and the NACA and those used for turbines by Ainley and Soderberg, plus an additional one, are compared explicitly for possible use in both compressors and turbines. Over a range of Mach numbers, loss coefficient values are compared with loss levels fixed and for representative blading cascade test data, and pressure recoveries and stage efficiencies are compared with loss coefficient values fixed. It is shown that for a low Mach number the different parameters are equal and interchangeable; however, as the Mach number increases, differences appear and grow larger, so that a given combination of loss coefficient value and Mach number implies different entropy-rise values depending upon which parameter is being used. The criteria used here for comparing the different parameters are that one loss coefficient is better than another a– if its loss coefficient values corresponding to test data vary less over a significant range of Mach number, and b– if the stage efficiency implied by a fixed loss coefficient value varies in a more realistic way over a range of Mach number. The Soderberg parameter was found to be better for both compressors and turbines than the other loss coefficients investigated.

Design and Development of a 12:1 Pressure Ratio Compressor for the Ruston 6-MW Gas Turbine

1982 ◽  
Author(s):  
F. Carchedi ◽  
G. R. Wood
Keyword(s):  
Gas Turbine ◽  
Test Data ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Aerodynamic Design ◽  
Design And Development ◽  
Surge Line ◽  
Flow Compressor ◽  
Industrial Gas Turbine ◽  
Spanwise Flow

This paper describes the design and development of a 15-stage axial flow compressor for a −6MW industrial gas turbine. Detailed aspects of the aerodynamic design are presented together with rig test data for the complete characteristic including stage data. Predictions of spanwise flow distributions are compared with measured values for the front stages of the compressor. Variable stagger stator blading is used to control the position of the low speed surge line and the effects of the stagger changes are discussed.

Some Blade Designs for an Axial Flow Compressor Stage

1954 ◽  
Vol 58 (517) ◽  
pp. 61-64
Author(s):  
R. G. Taylor
Keyword(s):  
Mach Number ◽  
Axial Flow ◽  
Compressor Stage ◽  
Axial Flow Compressor ◽  
Design Specifications ◽  
Reaction Condition ◽  
Forced Vortex ◽  
Vortex Solutions ◽  
Flow Compressor ◽  
Constant Mach Number

Two design conditions for an axial flow compressor stage are proposed and examined. These are, the constant reaction condition (incorporating I “ radial equilibrium ”), and the condition that the Mach number at inlet to the rotor shall be invariant with radius. In addition, the combination of these two properties in one stage is considered. It is found, with further assumptions regarding the nature of the flow, that a forced vortex type of flow will satisfy both design specifications. The forced vortex solutions for the various cases are presented, and for constant Mach number at inlet to the rotor, more general solutions are given.

Effects of Endwall Profiling on Axial Flow Compressor Stage

2009 ◽  
Author(s):  
Jialing Lu ◽  
Wuli Chu ◽  
Yanhui Wu
Keyword(s):  
Flow Field ◽  
Engineering Design ◽  
Axial Flow ◽  
Design Tool ◽  
Operating Conditions ◽  
Compressor Stage ◽  
Corner Separation ◽  
Axial Flow Compressor ◽  
Flow Compressor ◽  
Stage Efficiency

In recent years endwall profiling has been well validated as a major new engineering design tool for the reduction of secondary loss in turbines. However, its application on compressors have been rarely performed and reported. This paper documents the findings of the analysis for diminishing compressor stator corner separation using endwall profiling; In the study, novel profiled endwalls were designed and numerically studied on a subsonic axial-flow compressor stage. The compressor stator endwalls were profiled on both axial and azimuthal directions. The results showed, the stator corner separation was significantly suppressed under all the operating conditions by implementing this profiled endwall. Significant improvements on stage pressure ratios and stage efficiency were observed. Detailed flow field changes, as well as endwall profiling methods are provided in the paper, so that the results of this research can be referenced to other compressor designs.

Endwall Effects at Two Tip Clearances in a Multistage Axial Flow Compressor With Controlled Diffusion Blading

1994 ◽  
Vol 116 (4) ◽  
pp. 635-645 ◽  
Author(s):  
M. A. Howard ◽  
P. C. Ivey ◽  
J. P. Barton ◽  
K. F. Young
Keyword(s):  
Unsteady Flow ◽  
Axial Flow ◽  
Tip Clearance ◽  
Exit Angle ◽  
Controlled Diffusion ◽  
The Third ◽  
Flow Phenomena ◽  
Multistage Compressor ◽  
Flow Compressor ◽  
Stage Efficiency

Effects of tip clearance, secondary flow, skew, and corner stall on the performance of a multistage compressor with controlled diffusion blading have been studied experimentally. Measurements between 1 and 99 percent annulus height were carried out in both the first and the third stages of a four-stage low-speed compressor with repeating-stage blading. Measurements were obtained at a datum rotor tip clearance and at a more aerodynamically desirable lower clearance. The consequences of the modified rotor tip clearance on both rotor and stator performance are examined in terms of loss coefficient and gas exit angle. Stator losses close to the casing are found to increase significantly when the clearance of an upstream rotor is increased. These increased stator losses cause 30 percent of the stage efficiency reduction that arises with increased rotor tip clearance. The deviation angles due to tip clearance from the multistage measurements are found to be similar to data from single-stage machines with conventional blading, which suggests that the unsteady flow phenomena associated with the multistage environment do not dominate the physics of the flow.

scholarly journals Aerodynamic Modeling of Multistage Compressor Flowfields: Part 2 — Modeling Deterministic Stresses

1997 ◽  
Author(s):  
Edward J. Hall
Keyword(s):  
Test Data ◽  
Axial Flow ◽  
Time Dependent ◽  
State University ◽  
Turbine Engine ◽  
Prediction Scheme ◽  
Proposed Model ◽  
Multistage Compressors ◽  
Multistage Compressor ◽  
Flow Compressor

The primary purpose of this study was to investigate improved numerical techniques for predicting flows through multistage compressors. The vehicle chosen for this study was the Pennsylvania State University Research Compressor (PSRC). The PSRC facility consists of a 3-1/2 stage axial flow compressor which shares design features which are consistent with embedded stages of modern gas turbine engine axial flow compressors. In Part 2 of this two part paper, time-dependent predictions of rotor/stator/rotor aerodynamic interactions were employed to quantify the levels and distribution of deterministic stresses resulting from the average-passage flowfield description. Details of the spanwise and blade-to-blade distributions of the velocity correlations are examined and compared with results based on physical deterministic flow structures such as blade wakes and clearance flows. The predicted “apparent” wake profile decay resulting from the interaction of the wake through a downstream blade row is presented and compared with test data. This “apparent” wake profile decay is employed to define a simplified model for deterministic stress correlations in a steady state flowfield prediction scheme which retains the “mixing plane” methodology. Calculations based on this proposed model are described and predicted results are compared with both time-dependent predictions and test data. The resulting prediction strategy is both computational efficient and contains sufficient physical realism to permit its use in design studies.

Pressure loss of gaseous flow at a micro-tube outlet

2010 ◽  
Vol 225 (3) ◽  
pp. 649-657
Author(s):  
Y Horii ◽  
Y Asako ◽  
C Hong ◽  
J Lee
Keyword(s):  
Mach Number ◽  
Atmospheric Pressure ◽  
Pressure Loss ◽  
Stagnation Pressure ◽  
Loss Coefficient ◽  
Arbitrary Lagrangian Eulerian ◽  
Gaseous Flow ◽  
Pressure Loss Coefficient ◽  
Loss Coefficients ◽  
Energy Equations

The pressure loss of gaseous flow at a micro-tube outlet was investigated numerically. The numerical methodology is based on the arbitrary Lagrangian—Eulerian (ALE) method. Axis-symmetric compressible momentum and energy equations are solved to obtain the pressure loss coefficient of gaseous flow at a micro-tube outlet. Computed tube diameters are 50, 100, and 150μm. The stagnation pressure of upper stream of the tube is chosen in such a way that the Mach number at the tube outlet ranges from 0.1 to 1.2. The ambient (back) pressure is fixed at the atmospheric pressure. The pressure loss coefficients are compared with available experimental data for a conventionally sized tube. The effects of the Mach number and the tube diameter on the pressure loss coefficient are discussed and a correlation for the pressure loss coefficient is proposed.

Assessment of the Impact of a Heterogeneous Stator on the Noise of an Axial-Flow Low Mach-number Stage

2019 ◽  
Author(s):  
Miguel Pestana ◽  
Marlene Sanjose ◽  
Stephane Moreau ◽  
Michel Roger ◽  
Mathieu Gruber
Keyword(s):  
Mach Number ◽  
Axial Flow ◽  
Low Mach Number ◽  
The Impact

Performance of a jaws inlet under off-design conditions

2016 ◽  
Vol 231 (2) ◽  
pp. 294-305 ◽  
Author(s):  
Tianlai Gu ◽  
Shuai Zhang ◽  
Yao Zheng
Keyword(s):  
Mach Number ◽  
Flow Fields ◽  
Back Pressure ◽  
Low Mach Number ◽  
Mass Capture ◽  
Constant Area ◽  
Flow Capture ◽  
Mach Numbers ◽  
High Mach ◽  
Better Than

Numerical analysis was conducted of a jaws inlet under different working conditions, including angles of attack of 0° and 3°, varying Mach number, and varying back pressure with a constant-area isolator, to investigate its performance and flow fields of starting and unstarting states. Results reveal that the jaws inlet has an enhanced flow capture capability in starting states, with the mass capture ratio higher than 0.96, but relatively reduced working range of inflow Mach numbers. Its performance at a low Mach number is better than that at a high Mach number. Non-uniform flow fields are observed in unstarting cases at low Mach numbers and high back pressures, while separation structures are confined in the pitching compression section. Further increase in Mach number or decrease in back pressure does not result in significant changes in the separation structures. In the unstarting case under a high back pressure, it is hard to achieve restarting through reductions in the back pressure.

A Numerical Investigation of Loss Coefficient Variation in Various Incidence Angles in Tandem Blades Cascade

2014 ◽  
Author(s):  
Arash Soltani Dehkharqani ◽  
Masoud Boroomand ◽  
Hamzeh Eshraghi
Keyword(s):  
Pressure Ratio ◽  
Axial Flow ◽  
Suction Side ◽  
Loss Coefficient ◽  
Flow Coefficient ◽  
Flow Angle ◽  
Multi Stage ◽  
Wide Range ◽  
And Performance ◽  
Flow Compressor

There is a severe tendency to reduce weight and increase power of gas turbine. Such a requirement is fulfilled by higher pressure ratio of compressor stages. Employing tandem blades in multi-stage axial flow compressors is a promising methodology to control separation on suction sides of blades and simultaneously implement higher turning angle to achieve higher pressure ratio. The present study takes into account the high flow deflection capabilities of the tandem blades consisting of NACA-65 airfoil with fixed percent pitch and axial overlap at various flow incidence angles. In this regard, a two-dimensional cascade model of tandem blades is constructed in a numerical environment. The inlet flow angle is varied in a wide range and overall loss coefficient and deviation angles are computed. Moreover, the flow phenomena between the blades and performance of both forward and afterward blades are investigated. At the end, the aerodynamic flow coefficient of tandem blades are also computed with equivalent single blades to evaluate the performance of such blades in both design and off-design domain of operations. The results show that tandem blades are quite capable of providing higher deflection with lower loss in a wide range of operation and the base profile can be successfully used in design of axial flow compressor. In comparison to equivalent single blades, tandem blades have less dissipation because the momentum exerted on suction side of tandem blades confines the size of separation zone near trailing edges of blades.

Design and Development of a 12:1 Pressure Ratio Compressor for the Ruston 6-MW Gas Turbine

1982 ◽  
Vol 104 (4) ◽  
pp. 823-831 ◽  
Author(s):  
F. Carchedi ◽  
G. R. Wood
Keyword(s):  
Gas Turbine ◽  
Test Data ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Aerodynamic Design ◽  
Design And Development ◽  
Surge Line ◽  
Flow Compressor ◽  
Industrial Gas Turbine ◽  
Spanwise Flow

The paper describes the design and development of a 15 stage axial flow compressor for a 6-MW industrial gas turbine. Detailed aspects of the aerodynamic design are presented together with rig test data for the complete characteristic including stage data. Predictions of spanwise flow distributions are compared with measured values for the front stages of the compressor. Variable stagger stator blading is used to control the position of the low-speed surge line and the effects of the stagger changes are discussed.

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