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Methods and apparatus for cooling gas turbine engine blade tips

A technology for gas turbines and turbine blades, which is applied in the direction of blade support elements, engine cooling, turbine/propulsion device cooling, etc., and can solve problems such as high cost and limited life

Inactive Publication Date: 2005-01-19
GENERAL ELECTRIC CO
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

This cooling system is costly, uses large volumes of cooling air, and thus creates local life-limiting issues

Method used

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  • Methods and apparatus for cooling gas turbine engine blade tips
  • Methods and apparatus for cooling gas turbine engine blade tips
  • Methods and apparatus for cooling gas turbine engine blade tips

Examples

Experimental program
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Embodiment Construction

[0012] figure 1 is a schematic diagram of a gas turbine 10 including a low pressure compressor 12 , a high pressure compressor 14 and a combustor 16 . The gas turbine 10 also includes a high pressure turbine 18 and a low pressure turbine 20 . Compressor 12 and turbine 20 are connected by a first shaft 22 and compressor 14 and turbine 18 are connected by a second shaft 21 . In one embodiment, the gas turbine 10 is a GE90 gas turbine commercially available from General Electric Aircraft Engines, Inc. of Cincinnati, Ohio.

[0013] During operation, air flows through the low pressure compressor 12 and compressed air is supplied from the low pressure compressor 12 to the high pressure compressor 14 . Highly compressed air enters the combustion chamber 16 . Airflow from combustor 16 drives turbines 18 and 20 and exits gas turbine 10 through nozzle 24 .

[0014] FIG. 2 is a partial cross-sectional view of a rotor assembly 40 including a stator 42 that may be used with the gas tur...

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PUM

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Abstract

A turbine for a gas turbine engine including a turbine nozzle assembly that facilitates reducing an operating temperature of rotor blades in a cost-effective and reliable manner is described. Each rotor blade includes a tip that rotates in close proximity to a shroud that extends circumferentially around the rotor assembly. The turbine nozzle assembly includes a plurality of turbine vane segments that channel combustion gases to downstream rotor blades. Each turbine vane segment extends radially outward from an inner platform and includes a tip, a root, and a body that extends therebetween. The turbine vane segment tip is formed integrally with an outer band that mounts the vane segments within the gas turbine engine. The outer band is in flow communication with a cooling fluid source, and includes at least one opening.

Description

【Technical field】 [0001] The present invention relates to a turbine assembly and, more particularly, to a method and apparatus for cooling gas turbine rotor blade tips. 【Background technique】 [0002] A gas turbine typically includes a compressor, a combustor and at least one turbine. The compressor compresses the air that mixes with the fuel and enters the combustion chamber. The mixture is then ignited and hot combustion gases are produced, which enter a turbine, which extracts energy from the combustion gases to drive a compressor and generate useful work to propel an aircraft in flight or power a load such as a generator. [0003] A turbine includes a rotor assembly and a stator assembly. The rotor assembly includes a plurality of rotor blades extending radially outward from a rotor disk. Specifically, each rotor blade extends radially between the platforms adjacent the rotor disk to a blade tip. Combustion gas flow passages through the rotor assembly are defined rad...

Claims

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Application Information

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IPC IPC(8): F01D9/02F01D5/18F01D9/04F01D25/12F01D25/24F02C7/18
CPCF01D9/041F01D5/188Y02T50/676F05D2260/201F01D25/12F05D2260/205Y02T50/60
Inventor R·J·贝库克O·D·埃尔德曼R·E·阿坦斯H·P·小里克
Owner GENERAL ELECTRIC CO
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