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US 2011/0299979 A1Published application
Method for Improving the Stall Margin of an Axial Flow Compressor Using a Casing Treatment
AssigneeSiemens Energy, Inc.
InventorMatthew D. Montgomery
Life of this patent
Estimating the term…
Abstract
A method for determining a preferred circumferential groove arrangement for a casing treatment of an axial flow compressor is disclosed. The method includes using the results from a three dimensional steady state computational fluid dynamic analysis to generate a flow field between a blade tip of a rotating blade and a compressor casing to determine the preferred circumferential groove arrangement. A stall margin for the axial flow compressor will be increased with the method.
Claims20
A method of improving a stall margin of an axial flow compressor in a gas turbine engine, comprising the steps of:
- (a) analytically calculating a baseline performance from a baseline performance analysis for at least one row of rotating compressor blades;
- (b) analytically determining a flow field from the baseline performance analysis for the at least one row of rotating compressor blades at a blade tip region and at an off-design point;
- (c) analytically modeling at least one circumferential groove in a smooth wall of a compressor casing with a groove placement and a groove geometry determined using a set of results obtained in step (b), wherein operating the compressor with the circumferential groove in a casing wall increases a stall margin of the at least one row of rotating compressor blades;
- (d) performing a subsequent analytical performance calculation of the at least one row of rotating compressor blades with the at least one groove analytically modeled in a smooth wall of the compressor casing;
- (e) comparing a subsequent performance determined from the subsequent analytical performance calculation to the baseline performance, and a subsequent stall margin determined from the subsequent analytical performance calculation to a baseline stall margin determined from the baseline performance analysis;
- (f) determining if a change in the subsequent performance when compared to the baseline performance and a change in the subsequent stall margin when compared to the baseline stall margin satisfies an acceptance criteria;
- and
- (g) adjusting at least one of a plurality of groove parameters of the at least one groove and repeating steps (c) through (f) until the change in the subsequent performance to the baseline performance and the change in the subsequent stall margin to the baseline stall margin satisfies the acceptance criteria.
5 more elements
A gas turbine engine having an improved stall margin, comprising:
- an axial flow compressor element having an improved stall margin, comprising;
- a plurality of axially spaced circumferential grooves arranged above at least one rotating row of compressor blades, the plurality of axially spaced circumferential grooves in a casing wall, wherein the plurality of axially spaced circumferential grooves are sized according to a method comprising the steps of;
- (a) performing a first three dimensional steady state computational fluid dynamic analysis including viscous effects at a design operating point for the at least one single row of rotating compressor blades in a multistage compressor and performing a second three dimensional steady state computational fluid dynamic analysis including viscous effects for the at least one row of rotating compressor blades at an off-design operating point, the compressor having a compressor casing having a smooth wall;
- (b) calculating an aerodynamic blade performance for the at least one row of compressor blades using a set of results from the first three dimensional steady state computational fluid dynamic analysis and calculating a stall margin for the at least one row of compressor blades using a set of results from the second three dimensional steady state computational fluid dynamic analysis;
- (c) generating a flow field between a blade tip of the at least one row of rotating compressor blades and the smooth wall of the compressor casing from a set of results from the second three dimensional steady state computational fluid dynamic analysis;
- (d) determining a region in the flow field having a high pressure ratio at the blade tip, the region extending from a blade tip leading edge to a blade tip trailing edge and between adjacent blades in the at least one row of rotating compressor blades;
- (e) analytically modeling at least one circumferential groove in the smooth wall of the compressor casing proximate the region, wherein the groove channels the fluid having the high pressure ratio at the blade tip to flow in the circumferential direction with the flow exiting from the trailing edge blade tip and increasing the stall margin;
- (f) performing a subsequent three dimensional steady state computational fluid dynamic analysis for the single row of rotating compressor blades at the design operating point and the off-design operating point with the at least one groove analytically modeled in the smooth wall of the compressor casing;
- (g) calculating an aerodynamic blade performance of the single row of rotating compressor blades at the design point and a stall margin at the off-design point from the subsequent three dimensional steady state computational fluid dynamic analysis with the grooves analytically modeled in the casing and comparing to the aerodynamic blade performance and stall margin calculated in step (b);
- (h) repeating steps (e)-(g), varying at least one of a plurality of groove parameters until a change in stall margin of the at least one row of rotating compressor blades satisfies an acceptance criteria;
- a combustion element;
- and
- a turbine element.
10 more elements
Dependencies are inferred from the claim text.