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Turbulated combustor aft-end liner assembly and related cooling method

A liner, turbulator technology, used in the field of internal cooling, which can solve problems such as thermal gradients and pressure losses

Inactive Publication Date: 2010-12-15
GENERAL ELECTRIC CO
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  • Summary
  • Abstract
  • Description
  • Claims
  • Application Information

AI Technical Summary

Problems solved by technology

These various known techniques enhance heat transfer but have variable effects on thermal gradients and pressure losses

Method used

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  • Turbulated combustor aft-end liner assembly and related cooling method
  • Turbulated combustor aft-end liner assembly and related cooling method
  • Turbulated combustor aft-end liner assembly and related cooling method

Examples

Experimental program
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Embodiment Construction

[0019] FIG. 1 schematically depicts the interface region between the aft end of the combustor liner and the forward end of the transition piece in an annular gas turbine combustor 10 . As can be seen in this example, the transition piece 12 includes a radially inner transition piece body 14 and a radially outer transition piece impingement sleeve 16 spaced from the transition piece body 14 . Upstream of the transition piece 12 is a combustion liner 18 and a combustor flow sleeve 20 defined in surrounding relationship with the liner.

[0020] Flow from a gas turbine compressor (not shown) enters the casing 24 . About 50% of the compressor discharge air passes through orifices (not shown in detail) formed along and around the transition piece impingement sleeve 16 to transition between the transition piece body 14 and the radially outer Flow in the annular region or annular chamber 26 between the impingement sleeves 16 . The remaining approximately 50% of the compressor exhaus...

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PUM

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Abstract

In a combustor for a turbine a cover sleeve (140) is disposed between the aft end portion of the combustor liner (118) and a resilient seal structure (38) to define an air flow passage therebetween. The cover sleeve (140) has at a forward end thereof a plurality of air inlet feed holes (146) for directing cooling air into the air flow passage. A radially outer surface of the combustor liner aft end portion defining the air flow passage includes a plurality of turbulators (52) projecting towards but spaced from the cover sleeve (140) and a plurality of supports (144) extending to and engaging the cover sleeve to space the cover sleeve from the turbulators to define the air flow passage.

Description

technical field [0001] This invention relates to internal cooling within gas turbine engines; and more particularly, to methods for providing better and more uniform cooling in the transition region between the combustor liner and the transition duct that directs the combustion gases to the first stage of the turbine. Components and methods of cooling. Background technique [0002] Conventional gas turbine combustors use diffusion (ie, non-premixed) combustion, where fuel and air enter the combustion chamber separately. The process of mixing and burning produces flame temperatures in excess of 3900°F. Since conventional combustor liners and / or transition pieces are generally able to withstand a maximum temperature of only about 1500°F (about 820°C) for about ten thousand hours (10,000 hours), for durability, creep resistance and seal integrity To ensure reliability, measures must be taken to protect the combustor liner and / or transition duct, as well as the seal configurat...

Claims

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Application Information

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Patent Type & Authority Applications(China)
IPC IPC(8): F23D14/46F23D14/00F23D14/70F23D14/78
CPCF23R3/04F01D9/023F05D2260/22141F23R2900/03045F23R2900/00012
Inventor J·D·布朗M·伯克曼S·K·富尔赫尔A·史密斯
Owner GENERAL ELECTRIC CO