Compressor rotor with anti-vortex fins

Patent metadata

Field	Value
Publication number	US-10012240-B2
Application number	US-201715492281-A
Country	US
Kind code	B2
Filing date	Apr 20, 2017
Priority date	Aug 29, 2014
Publication date	Jul 3, 2018
Grant date	Jul 3, 2018

Abstract

A compressor rotor of a gas turbine engine includes a rotor body having a face adapted to face an adjacent rotor. The rotor body extends radially between an outer peripheral rim surface and an inner rim surface. The inner rim surface defines a bore of the rotor body. A plurality of blades extends radially from the outer peripheral rim surface. A plurality of anti-vortex fins extends axially from the face of the rotor body facing the adjacent rotor. The plurality of anti-vortex fins forms a plurality of open radial passageways. The plurality of anti-vortex fins extends axially to a predetermined thickness such that, when assembled with the second rotor, axial extremities of the plurality of anti-vortex fins being in close proximity with the adjacent rotor and the adjacent rotor closes the radial passageways. A method of providing a first rotor for assembly with a second facing rotor of a compressor rotor assembly is also presented.

First claim

The invention claimed is:

1. A compressor rotor of a gas turbine engine, the compressor rotor comprising:
a rotor body having a face adapted to face an adjacent rotor, the rotor body extending radially between an outer peripheral rim surface and an inner rim surface, the inner rim surface defining a central bore of the rotor body;
a plurality of blades extending radially outwardly from the outer peripheral rim surface into a compressor gaspath;
a plurality of anti-vortex fins extending axially from the face of the rotor body facing the adjacent rotor, the plurality of anti-vortex fins extending radially outwardly from the inner rim surface towards the outer peripheral rim surface along a major portion of a radial distance between the outer peripheral rim surface and the central bore, the plurality of anti-vortex fins forming a plurality of open radial passageways configured to direct compressor bleed air radially inwardly from the compressor gaspath to the central bore, the plurality of anti-vortex fins extending axially to a predetermined thickness such that, when assembled with the second rotor, axial extremities of the plurality of anti-vortex fins being in close proximity with the adjacent rotor and the adjacent rotor closes the radial passageways; and
an intermediate rim surface on the face of the rotor body facing the adjacent rotor, the intermediate rim surface including a plurality of openings in fluid flow communication with respective ones of the radial passageways.

2. The compressor rotor of claim 1 , wherein the intermediate rim surface is disposed between the outer peripheral rim surface and the inner rim surface and connects the plurality of anti-vortex fins, and wherein the openings are in a one-to-one relationship with the plurality of radial passageways.

3. The compressor rotor of claim 1 , wherein the radial passageways are disposed circumferentially spaced-apart.

4. The compressor rotor of claim 1 , wherein the radial passageways are curved in a direction of rotation of the compressor rotor.

5. The compressor rotor of claim 1 , wherein the anti-vortex fins are integral to the rotor body.

6. The compressor rotor of claim 1 , wherein the radial passageways are tapered toward the inner peripheral rim surface.

7. A compressor rotor assembly of gas turbine engine, the compressor rotor assembly comprising:
first and second adjacent rotors, the first rotor including:
a rotor body, the rotor body having axially opposed faces, the rotor body extending radially between an outer peripheral rim surface and an inner rim surface, the inner rim surface defining a central bore of the body;
a plurality of blades extending radially from the outer peripheral rim surface into a compressor gaspath;
a plurality of anti-vortex fins extending axially from the face of the first rotor facing the second rotor, the plurality of anti-vortex fins extending radially outwardly from the inner rim surface towards the outer peripheral rim surface, the plurality of anti-vortex fins forming a plurality of radial passageways closed by the second rotor so that the radial passageways are fluidly independent from each other, the radial passageways being configured to individually direct compressor bleed air radially inwardly from the compressor gaspath to the central bore, and an intermediate rim surface on the face of the first rotor facing the second rotor, the intermediate rim surface including a plurality of openings connected in fluid flow communication to respective ones of the radial passageways.

8. The gas turbine engine of claim 7 , further comprising an intermediate rim surface disposed on the face of the first rotor facing the second rotor, the intermediate rim surface connecting the plurality of anti-vortex fins.

9. The gas turbine engine of claim 7 , wherein the radial passageways are disposed circumferentially spaced-apart.

10. The gas turbine engine of claim 7 , wherein the radial passageways are curved in a direction of rotation of the rotor.

11. The gas turbine engine of claim 7 , wherein the second rotor is an impeller.

12. The gas turbine engine of claim 7 , wherein the radial passageways are tapered toward the inner peripheral rim surface.

Assignees

• Pratt & Whitney Canada

Inventors

• Urac Tibor

Classifications

• F02C6/08

  the gas being bled from the gas-turbine compressor · CPC title

• F04D29/667 Primary

  by influencing the flow pattern, e.g. suppression of turbulence · CPC title

• F04D29/321 Primary

  for axial flow compressors · CPC title

• F04D29/284

  for compressors · CPC title

• F01D5/046

  Heating, heat insulation or cooling means · CPC title