A New Streamline Curvature Throughflow Method for Radial Turbomachinery

2008 ◽  
Author(s):  
Michael Casey ◽  
Chris Robinson
Keyword(s):  
Pressure Ratio ◽  
Compressible Fluids ◽  
Preliminary Design ◽  
Streamline Curvature ◽  
Trailing Edges ◽  
Blade Row ◽  
Curved Walls ◽  
Excellent Tool ◽  
Simple Numerical Model ◽  
General Method

This paper describes a newly developed streamline curvature throughflow method for the analysis of radial or mixed flow machines. The code includes curved walls, curved leading and trailing edges, and internal blade row calculating stations. A general method of specifying the empirical data provides separate treatment of blockage, losses, and deviation. Incompressible and compressible fluids are allowed, including real gases and supersonic relative flow in blade rows. The paper describes some new aspects of the code. In particular, a relatively simple numerical model for spanwise mixing is derived, the calculation method for prescribed pressure ratio in compressor bladed rows is described, and the method used to redistribute the flow across the span due to choking is given. Examples are given of the use and validation of the code for many types of radial turbomachinery and these show it is an excellent tool for preliminary design.

A New Streamline Curvature Throughflow Method for Radial Turbomachinery

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Michael Casey ◽  
Chris Robinson
Keyword(s):  
Pressure Ratio ◽  
Compressible Fluids ◽  
Preliminary Design ◽  
Streamline Curvature ◽  
Trailing Edges ◽  
Blade Row ◽  
Curved Walls ◽  
Excellent Tool ◽  
Simple Numerical Model ◽  
General Method

This paper describes a newly developed streamline curvature throughflow method for the analysis of radial or mixed flow machines. The code includes curved walls, curved leading and trailing edges, and internal blade row calculating stations. A general method of specifying the empirical data provides separate treatment of blockage, losses, and deviation. Incompressible and compressible fluids are allowed, including real gases and supersonic relative flow in blade rows. The paper describes some new aspects of the code. In particular, a relatively simple numerical model for spanwise mixing is derived; the calculation method for prescribed pressure ratio in compressor blade rows is described; and the method used to redistribute the flow across the span due to choking is given. Examples are given of the use and validation of the code for many types of radial turbomachinery, and these show that it is an excellent tool for preliminary design.

The Effect of Inlet Conditions on Turbine Endwall Loss

2021 ◽  
Author(s):  
John D. Coull ◽  
Christopher J. Clark
Keyword(s):  
Boundary Layers ◽  
Computational Study ◽  
Preliminary Design ◽  
Flow Conditions ◽  
Blade Row ◽  
Inlet Conditions ◽  
Axial Turbines ◽  
Design Optimum ◽  
Inlet Boundary Condition ◽  
General Method

Abstract There can be significant variation and uncertainty in the flow conditions entering a blade row. This paper explores how this variability can affect endwall loss in axial turbines. A computational study of three cascades with collinear inlet boundary layers is conducted. Endwall loss varies by more than a factor of 3 depending on the inlet conditions. This variation is caused by dissipation of Secondary Kinetic Energy (SKE). The results can be understood by observing that the inlet conditions predominantly control how secondary vorticity is distributed within the blade passage. Modestly-thick inlet boundary layers with high shape factor tend to displace vorticity towards the center of the passage. This displacement reduces vorticity cancellation, increasing secondary velocities and SKE. A general method is formulated to estimate SKE in preliminary design. Optimum aspect ratio is shown to depend on the inlet boundary condition. Strategies to reduce endwall loss and minimize sensitivity to inlet conditions are then highlighted.

New Guidelines for the Preliminary Design and Performance Prediction of Axial-Flow Turbines

1987 ◽  
Vol 201 (4) ◽  
pp. 279-290 ◽  
Author(s):  
D G Wilson
Keyword(s):  
Pressure Ratio ◽  
Axial Flow ◽  
Flow Coefficient ◽  
Preliminary Design ◽  
Outlet Angle ◽  
Blade Row ◽  
Final Design ◽  
And Performance ◽  
New Guidelines ◽  
Design Freedom

Performance charts are included that enable appropriate initial choices (that is, before final design by methods of computational fluid dynamics) to be made of reaction and flow coefficient (or blade row outlet angle) for specified stage loadings in axial-flow turbines. The radial variation in vector diagrams that leads to acceptable radial distributions of reaction at various hub-shroud diameter ratios and loadings are discussed and plotted. From these, the ranges of design freedom to choose the annulus cross-section (for example constant o.d. or i.d.) are plotted as a function of turbine pressure ratio. Also developed is a method of correlating and adding losses that should be more consistent and thermodynamically more rigorous than some previously used methods.

scholarly journals Through-Flow Analysis for Axial-Stage Design Including Streamline-Slope Effects

1993 ◽  
Author(s):  
Theodosios Korakianitis ◽  
Dequan Zou
Keyword(s):  
Mass Flow ◽  
Three Dimensional ◽  
Momentum Equation ◽  
Total Enthalpy ◽  
Streamline Curvature ◽  
Flow Balance ◽  
Through Flow ◽  
Blade Row ◽  
Iteration Loop ◽  
Predictor Corrector

This paper presents a new method to design (or analyze) subsonic or supersonic axial compressor and turbine stages and their three-dimensional velocity diagrams from hub to tip by solving the three-dimensional radial-momentum equation. Some previous methods (matrix through-flow based on the streamfunction approach) can not handle locally supersonic flows, and they are computationally intensive when they require the inversion of large matrices. Other previous methods (streamline curvature) require two nested iteration loops to provide a converged solution: an outside iteration loop for the mass-flow balance; and an inside iteration loop to solve the radial momentum equation at each flow station. The present method is of the streamline-curvature category. It still requires the iteration loop for the mass-flow balance, but the radial momentum equation at each flow station is solved using a one-pass numerical predictor-corrector technique, thus reducing the computational effort substantially. The method takes into account the axial slope of the streamlines. Main design characteristics such as the mass-flow rate, total properties at component inlet, hub-to-tip ratio at component inlet, total enthalpy change for each stage, and the expected efficiency of each streamline at each stage are inputs to the method. Other inputs are the radial position and axial velocity component at one surface of revolution through the axial stages. These can be provided for either the hub, or the mean, or the tip location of the blading. In addition the user specifies the azimuthal deflection of the flow from the axial direction at each radius (or as a function of radius) at each blade row inlet and outlet. By construction the method eliminates radial variations of total enthalpy (work) and entropy at each blade row inlet and outlet. In an alternative formulation enthalpy variations across radial positions at each axial station are included in the analysis. The remaining three-dimensional velocity diagrams from hub to tip, and the radial location of the remaining streamlines, are obtained by solving the momentum equation using a predictor-corrector method. Examples for one turbine and one compressor design are included.

A Rapid Viscous-Inviscid Interaction Method for the Preliminary Design of S-Shaped Transition Ducts

2021 ◽  
Author(s):  
A. Veyrat ◽  
J. F. Carrotte ◽  
A. D. Walker ◽  
C. Hall ◽  
H. Simpson
Keyword(s):  
Shear Stress ◽  
Rapid Assessment ◽  
Layer Growth ◽  
Preliminary Design ◽  
Interaction Method ◽  
Complex Flow ◽  
Potential Core ◽  
Law Of The Wall ◽  
Streamline Curvature ◽  
Use Of Time

Abstract For preliminary design of compressor transition ducts, knowledge-based tools for the rapid assessment of aerodynamic performance of S-shaped ducts are not currently available in the open literature. This is due to the highly complex flow developing under the combined influence of pressure gradients and streamline curvature. This paper presents a new approach enabling an agile design process avoiding premature use of time-consuming high-fidelity CFD calculations. The features of a 2D axisymmetric incompressible steady flow field are captured with a semi-analytical viscous inviscid interaction method. A potential core, based on streamline curvature and implicit velocity profile by parametric spline reconstruction, is coupled to an integral method predicting the turbulent boundary layer growth up to separation. The shear stress distribution is generated by a modified mixing length model for strongly curved flows and wall shear stress closure is performed by inverse calculation of a composite law-of-the-wall. When compared to CFD, the aerodynamic loading is generally predicted to within ±3% but convergence is achieved 20 times faster.

scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Vol 140 (12) ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Margin ◽  
Point Reduction ◽  
Angle Deviation ◽  
Blade Row

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Design Study of an Advanced Concept Simple Gas Turbine for Possible Use in Low Emission Automobiles

1972 ◽  
Author(s):  
H. C. Eatock ◽  
M. D. Stoten
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Simple Cycle ◽  
Turbine Engines ◽  
Nox Emissions ◽  
Preliminary Design ◽  
Gas Turbine Engines ◽  
High Pressure Ratio

United Aircraft Corporation studied the potential costs of various possible gas turbine engines which might be used to reduce automobile exhaust emissions. As part of that study, United Aircraft of Canada undertook the preliminary design and performance analysis of high-pressure-ratio nonregenerated (simple cycle) gas turbine engines. For the first time, high levels of single-stage component efficiency are available extending from a pressure ratio less than 4 up to 10 or 12 to 1. As a result, the study showed that the simple-cycle engine may provide satisfactory running costs with significantly lower manufacturing costs and NOx emissions than a regenerated engine. In this paper some features of the preliminary design of both single-shaft and a free power turbine version of this engine are examined. The major component technology assumptions, in particular the high pressure ratio centrifugal compressor, employed for performance extrapolation are explained and compared with current technology. The potential low NOx emissions of the simple-cycle gas turbine compared to regenerative or recuperative gas turbines is discussed. Finally, some of the problems which might be encountered in using this totally different power plant for the conventional automobile are identified.

1997 Best Paper Award—Education Committee: A Three-Dimensional Turbine Engine Analysis Compressor Code (TEACC) for Steady-State Inlet Distortion

1998 ◽  
Vol 120 (3) ◽  
pp. 422-430 ◽  
Author(s):  
A. Hale ◽  
W. O’Brien
Keyword(s):  
Three Dimensional ◽  
Simulation Tool ◽  
Source Terms ◽  
Turbine Engine ◽  
Streamline Curvature ◽  
Inlet Distortion ◽  
Blade Row ◽  
Practical Design ◽  
Distorted Region ◽  
Computationally Intensive

The direct approach of modeling the flow between all blade passages for each blade row in the compressor is too computationally intensive for practical design and analysis investigations with inlet distortion. Therefore a new simulation tool called the Turbine Engine Analysis Compressor Code (TEACC) has been developed. TEACC solves the compressible, time-dependent, three-dimensional Euler equations modified to include turbomachinery source terms, which represent the effect of the blades. The source terms are calculated for each blade row by the application of a streamline curvature code. TEACC was validated against experimental data from the transonic NASA rotor, Rotor 1B, for a clean inlet and for an inlet distortion produced by a 90-deg, one-per-revolution distortion screen. TEACC revealed that strong swirl produced by the rotor caused the compressor to increase in loading in the direction of rotor rotation through the distorted region and decrease in loading circumferentially away from the distorted region.

Improvements to the Ainley-Mathieson Method of Turbine Performance Prediction

1970 ◽  
Vol 92 (3) ◽  
pp. 252-256 ◽  
Author(s):  
J. Dunham ◽  
P. M. Came
Keyword(s):  
Performance Prediction ◽  
Tip Clearance ◽  
Project Work ◽  
Preliminary Design ◽  
Design Work ◽  
Turbine Performance ◽  
Wide Range ◽  
Blade Row ◽  
Loss Prediction ◽  
Made In

In 1951 Ainley and Mathieson published a method of predicting the design and off-design performance of an axial turbine (British ARC, R & M 2974). The flow and hence the losses were calculated at a single “reference diameter” for each blade row. This method has been widely used ever since. A critical review of the method has been made, based on detailed comparisons between the measured and predicted performance of a wide range of modern turbines. As a result, improvements have been made in the formulas for secondary loss and tip clearance loss prediction. The accuracy of the improved method has been assessed. Despite its relatively simple approach, it is believed that it will remain of great value in project work and preliminary design work.

Health Monitoring of Variable Geometry Gas Turbines for the Canadian Navy

1989 ◽  
Vol 111 (2) ◽  
pp. 244-250 ◽  
Author(s):  
D. E. Muir ◽  
H. I. H. Saravanamuttoo ◽  
D. J. Marshall
Keyword(s):  
Health Monitoring ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Engine Performance ◽  
Variable Geometry ◽  
Axial Compressors ◽  
High Pressure Ratio ◽  
Department Of National Defence ◽  
Canadian Navy ◽  
General Method

The Canadian Department of National Defence has identified a need for improved Engine Health Monitoring procedures for the new Canadian Patrol Frigate (CPF). The CPF propulsion system includes two General Electric LM2500 gas turbines, a high-pressure-ratio engine with multiple stages of compressor variable geometry. A general method for predicting the thermodynamic performance of variable geometry axial compressors has been developed. The new modeling technique is based on a meanline stage-stacking analysis and relies only on the limited performance data typically made available by engine manufacturers. The method has been applied to the LM2500-30 marine gas turbine and the variations in engine performance that can result from a malfunction of the variable geometry system in service have been estimated.

Sign in / Sign up

Export Citation Format

Share Document