COMPRESSOR ROTOR APPARATUS WITH REDUCED END STAGE TEMPERATURE GRADIENT
Publication Date: 2017-Jan-24
The IP.com Prior Art Database
In a gas turbine engine, a compressor disk is washed with heated air. This reduces temperature gradients in the disk and increases fatigue life.
- 1 -
COMPRESSOR ROTOR APPARATUS WITH REDUCED END STAGE
 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 disclosure relates generally to gas turbine engines and more particularly to
the compressor rotors in such engines.
 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.
 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.
 Described herein is a configuration for a compressor rotor having an improved
BRIEF DESCRIPTION OF THE DRAWINGS
 The concept may be best understood by reference to the following description
taken in conjunction with the accompanying drawing figures, in which:
 FIG. 1 is a schematic cross-sectional view of a portion of a conventional
- 2 -
compressor rotor assembly;
 FIG. 2 is a schematic cross-sectional view of a modified compressor rotor
assembly incorporating a plenum therein; and
 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
 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.
 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.
 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.
 A conical aft arm...