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COMPRESSOR ROTOR APPARATUS WITH REDUCED END STAGE TEMPERATURE GRADIENT

IP.com Disclosure Number: IPCOM000248968D
Publication Date: 2017-Jan-24
Document File: 7 page(s) / 150K

Publishing Venue

The IP.com Prior Art Database

Abstract

In a gas turbine engine, a compressor disk is washed with heated air. This reduces temperature gradients in the disk and increases fatigue life.

This text was extracted from a PDF file.
This is the abbreviated version, containing approximately 49% of the total text.

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COMPRESSOR ROTOR APPARATUS WITH REDUCED END STAGE

TEMPERATURE GRADIENT

ABSTRACT

[0001] In a gas turbine engine, a compressor disk is washed with heated air. This reduces

temperature gradients in the disk and increases fatigue life.

BACKGROUND

[0002] This disclosure relates generally to gas turbine engines and more particularly to

the compressor rotors in such engines.

[0003] A gas turbine engine includes a compressor used to pressurize intake air which

then flows to a downstream combustor and one or more turbines. A typical compressor

includes a series of stages, each stage including a row of stationary stator vanes and a row

of rotating compressor blades.

[0004] With the development of newer gas turbine engines, the Brayton cycle is being

pushed to higher limits resulting in higher temperatures at the aft end of the compressor.

This results in very high temperature gradients from the outer periphery of the

compressor end-stage rotor disk to the disk web. The higher the gradients, the greater the

thermal stresses in the rotor disk and the lower its low cycle fatigue (LCF) life.

[0005] Described herein is a configuration for a compressor rotor having an improved

fatigue life.

BRIEF DESCRIPTION OF THE DRAWINGS

[0006] The concept may be best understood by reference to the following description

taken in conjunction with the accompanying drawing figures, in which:

[0007] FIG. 1 is a schematic cross-sectional view of a portion of a conventional

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compressor rotor assembly;

[0008] FIG. 2 is a schematic cross-sectional view of a modified compressor rotor

assembly incorporating a plenum therein; and

[0009] FIG. 3 is a schematic cross-sectional view of a portion of an alternative

compressor rotor assembly incorporating a plenum.

DETAILED DESCRIPTION OF THE CONCEPT

[0010] Illustrated in FIG. 1 is a schematic representation of a portion of a rotor assembly

10 of a high-pressure compressor ("HPC") of a gas turbine engine. It will be understood

that the compressor includes a number of stages of axial-flow blading; for example a

typical compressor could include 8-10 stages. In operation, the static air pressure is

incrementally increased by each subsequent compressor stage, with the final stage

discharging air at the intended compressor discharge pressure ("CDP") for subsequent

flow into a diffuser and thence into a combustor. The concepts described herein relate to

the configuration at the aft end (exit) of the compressor.

[0011] The rotor assembly 10 includes a final stage rotor disk 12 with a bore 14, a web

16, and a rim 18. The rim 18 is integral with a generally cylindrical compressor spool 20.

The compressor spool 20 incorporates an inter-stage seal 22 disposed immediately

upstream of the rim 18.

[0012] The rim 18 includes a circumferential dovetail slot carrying a row of airfoil-

shaped compressor blades 24. The tips 26 of the compressor blades 24 run in close

proximity to a surrounding annular compressor casing 28.

[0013] A conical aft arm...