Rotor Casing Contouring in High Pressure Stages of Heavy Duty Gas Turbine Compressors With Large Tip Clearance Heights

2009 ◽  
Author(s):  
Georg Kro¨ger ◽  
Christian Cornelius ◽  
Eberhard Nicke
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Performance Improvement ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Pressure Stage ◽  
Design Point ◽  
Pressure Loss Coefficient ◽  
Exit Pressure ◽  
High Pressure Stage

Clearance leakage losses of axial compressor rotors and stators have a major impact on the overall compressor performance. The clearance heights in the last stages (high pressure stages) of a gas turbine compressor are very large in comparison to the low pressure stages due to mechanical constraints and small blade heights. The reduction of clearance leakage losses in a high pressure stage still holds an important potential for the overall performance improvement at design point conditions. In the following work, a method for tip clearance loss reduction by circumferential casing contouring above a high pressure stage rotor with a constant clearance height is presented. The subsonic compressor blade provides Siemens HPA-Family [1, 2, 3] airfoils. Starting over with a 3D-Optimization of the mentioned rotor casing the work additionally refers to the aerodynamic effects and the off design performance of the optimized geometry. It has been found that an optimized casing and blade tip contour lead to a smaller overall clearance mass flow and lower pressure loss coefficient of the clearance flow so that the endwall blockage is reduced and the stage performance is improved by about 0.35% at design point conditions. Furthermore it was found that the performance improvement drops with increasing exit pressure to about 0.1% close to stall conditions. At lower exit pressure values the optimized geometry provides an additional performance improvement in comparison to the baseline configuration.

scholarly journals Analysis and Improvement of a Two-Stage Centrifugal Compressor Used in an MW-Level Gas Turbine

2018 ◽  
Vol 8 (8) ◽  
pp. 1347 ◽  
Author(s):  
Wei Zhu ◽  
Xiao-Dong Ren ◽  
Xue-Song Li ◽  
Chun-Wei Gu
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Low Pressure ◽  
Two Stage ◽  
Vaned Diffuser ◽  
Pressure Stage ◽  
Optimization System ◽  
High Pressure Stage

The performance of a low/high-pressure-stage centrifugal compressor in a land-use MW-level gas turbine with a pressure ratio of approximately 11 is analyzed and optimized with a 1D aerodynamic design and modeling optimization system. 1D optimization results indicate that the diameter ratio of the low-pressure-stage centrifugal compressor with a vane-less diffuser, and the divergent angle of the high-pressure-stage centrifugal compressor with a vaned diffuser, are extremely large and result in low efficiency. Through modeling design and optimization system analysis, a tandem vaned diffuser is used in the low-pressure stage, and a tandem vaned diffuser with splitter vanes is adopted in the high-pressure stage. Computational fluid dynamics (CFD) results show that the pressure ratio and efficiency of the optimized low/high-pressure-stage centrifugal compressor are significantly improved. Coupling calculations of the low/high-pressure stage of the original and optimized designs are conducted based on the results of MW-level gas turbine cycles. CFD results show that the pressure ratio and efficiency of the optimized two-stage centrifugal compressor increase by approximately 8% and 4%, respectively, under three typical load conditions of 100%, 90%, and 60%.

Design and Analysis of a High Pressure Turbine Using Computational Methods for Small Gas Turbine Application

2013 ◽  
Author(s):  
Kishor Kumar ◽  
R. Prathapanayaka ◽  
S. V. Ramana Murthy ◽  
S. Kishore Kumar ◽  
T. M. Ajay Krishna
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Computational Methods ◽  
Gas Turbine Engine ◽  
Flow Coefficient ◽  
Design Parameters ◽  
Convergence Criterion ◽  
Turbine Engine ◽  
Design Point ◽  
Ansys Cfx

This paper describes the aerodynamic design and analysis of a high-pressure, single-stage axial flow turbine suitable for small gas turbine engine application using computational methods. The specifications of turbine were based on the need of a typical high-pressure compressor and geometric restrictions of small gas turbine engine. Baseline design parameters such as flow coefficient, stage loading coefficient are close to 0.23 and 1.22 respectively with maximum flow expansion in the NGV rows. In the preliminary design mode, the meanline approach is used to generate the turbine flow path and the design point performance is achieved by considering three blade sections at hub, mean and tip using the AMDC+KO+MK+BSM loss models to meet the design constraints. An average exit swirl angle of less than 5 degrees is achieved leading to minimum losses in the stage. Also, NGV and rotor blade numbers were chosen based on the optimum blade solidity. Blade profile is redesigned using the results from blade-to-blade analysis and through-flow analysis based on an enhanced Dawes BTOB3D flow solver. Using PbCFD (Pushbutton CFD) and commercially available CFD software ANSYS-CFX, aero-thermodynamic parameters like pressure ratios, aerodynamic power, and efficiencies are computed and these results are compared with one another. The boundary conditions, convergence criterion, and turbulence model used in CFD computations are set uniform for comparison with 8 per cent turbulence intensity. Grid independence study is performed at design point to optimize the grid density for off-design performance predictions. ANSYS-CFX and PbCFD have predicted higher efficiency of 3.4% and 1.2% respectively with respect to targeted efficiency of 89 per cent.

A Comparison Between the Design Point and Near-Stall Performance of an Axial Compressor

1990 ◽  
Vol 112 (1) ◽  
pp. 109-115 ◽  
Author(s):  
N. M. McDougall
Keyword(s):  
Rotor Blade ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Design Point ◽  
Blade Row ◽  
Clearance Flow ◽  
Stator Blade

Detailed measurements have been made within an axial compressor operating both at design point and near stall. Rotor tip clearance was found to control the performance of the machine by influencing the flow within the rotor blade passages. This was not found to be the case in the stator blade row, where hub clearance was introduced beneath the blade tips. Although the passage flow was observed to be altered dramatically, no significant changes were apparent in the overall pressure rise or stall point. Small tip clearances in the rotor blade row resulted in the formation of corner separations at the hub, where the blade loading was highest. More representative clearances resulted in blockage at the tip due to the increased tip clearance flow. The effects that have been observed emphasize both the three-dimensional nature of the flow within compressor blade passages, and the importance of the flow in the endwall regions in determining the overall compressor performance.

Effect of Rotor Tip Gap Variation at the Rear Stages of an Axial Flow Compressor

2012 ◽  
Vol 225 ◽  
pp. 233-238
Author(s):  
A.M. Pradeep ◽  
R.N. Chiranthan ◽  
Debarshi Dutta ◽  
Bhaskar Roy
Keyword(s):  
Rotor Blade ◽  
Cross Flow ◽  
Axial Compressor ◽  
Axial Flow ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Suction Surface ◽  
Design Point ◽  
Compressor Rotor ◽  
Multi Stage

In this paper, detailed analysis of the tip flow of an axial compressor rotor blade has been carried out using the commercial CFD package ANSYS CFX. The rotor blade was designed such that it is reminiscent of the rear stages of a multi-stage axial compressor. The effects of varying tip gaps are studied using CFD simulations for overall pressure rise and flow physics of the tip flow at the design point and near the peak pressure point. Rig tests of a low speed research compressor rotor with 3% tip clearance provided characteristics plots for validation of the CFD results. With increase in clearance from 1% to 4%, the rotor pressure rise at the design point was observed to decrease linearly. Increase in the clearance increases the cross flow across the tip; however, the magnitude of the average jet velocity crossing the tip decreases. The tip leakage vortex was observed to stay close to the suction surface with increase in clearance.

Brush Seals for the No. 3 Bearing of a Model 7EA Gas Turbine

2000 ◽  
Author(s):  
Steve Ingistov ◽  
Gary Meredith ◽  
Erik Sulda
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Elevated Temperatures ◽  
Axial Compressor ◽  
Lubricating Oil ◽  
Additional Element ◽  
The Third ◽  
Air Stream ◽  
Brush Seals

Gas Turbines in power generation are frequently of the single rotor type. The rotor is directly connected to the electrical generator. The rotor may be supported by two journal bearings or in some cases there is an additional journal bearing situated between the axial compressor discharge and the gas turbine intake. This third bearing serves to provide the rotor with additional support required to reduce rotor dynamic instabilities. The third bearing is, therefore, inside the machine housing and a significant amount of maintenance work is necessary to inspect it. The third bearing is also exposed to elevated temperatures by, essentially, being surrounded by compressor discharge air. A certain amount of compressor discharge air leaks through the seals into the cylindrical space around the third bearing housing and from there, due to significant pressure gradients, into the third bearing. Labyrinth seals are provided to impede air leakage from the pressurized cylindrical space into the bearing cavity. The air that leaks into the bearing housing mixes with a buffer air stream. This buffer air stream serves to cool the bearing cavity and to prevent leakage of hot, high-pressure air into the bearing cavity. Two dry air streams are then routed into the atmosphere via the coaxial space formed by two cylindrical surfaces. The portion of the buffer air stream contacting the bearing lubricating oil is de-misted in a special de-mister vessel. The de-misted air is exhausted into the atmosphere and the separated oil is returned to the gas turbine lubricating oil reservoir. This Paper discusses the introduction of brush seals into the No. 3 bearing housing as an additional element in retarding the high pressure, high temperature air infiltration into the No. 3 bearing housing.

Sign in / Sign up

Export Citation Format

Share Document