Seal-plate anti-rotation in a stage of a gas turbine engine

Patent metadata

Field	Value
Publication number	US-10352175-B2
Application number	US-201615258721-A
Country	US
Kind code	B2
Filing date	Sep 7, 2016
Priority date	Sep 21, 2015
Publication date	Jul 16, 2019
Grant date	Jul 16, 2019

Abstract

A turbine stage assembly includes a disc carrying a cascade of blades and an annular seal-plate that is secured to the disc by a first connection. One or more of the blades include a root portion configured to be received in a complementarily shaped radially extending slot in the disc such that a face of the root portion faces the plate. A terminal portion of the root portion is cut away adjacent the face to present an open space between a radially inner wall of the slot and a wall of the cutaway root portion. A first part of a second connector extends radially inwardly from the wall of the cutaway root portion. The first part of the second connector is configured to engage with a complementing second part of the second connector provided on the seal-plate.

First claim

The invention claimed is:

1. A turbine stage assembly comprising:
a disc carrying a cascade of blades and an annular seal-plate, the annular seal-plate being secured to the disc by a first connection means;
one or more of the blades comprising a main body and a root portion configured to be received in a complementarily shaped radially extending recess in the disc such that a face of the root portion faces the annular seal-plate, the root portion defining a wall of a duct, the wall of the duct providing, in use, a heat shield for the base of the radially extending recess into which the root portion is received, the duct being arranged to receive cooling air and being in fluid communication with cooling passages extending through the root portion and into the main body of the blade;
a terminal portion of the root portion being cut away axially adjacent the wall of the duct to the face of the root portion to present an open space between a radially inner wall of the radially extending recess and a radially opposing wall of the cutaway root portion that includes an orifice that opens into a cooling passage inside the main body of the blade for delivery of coolant to the cooling passage; and
a first part of a second connector means extending radially inwardly from the radially opposing wall of the cutaway root portion towards the radially inner wall of the radially extending recess and partly across the open space, the first part of the second connector means comprising a pair of tangs defining a slot into which is received a protrusion forming a complementing second part of the second connector means extending from the seal-plate.

2. The turbine stage assembly as claimed in claim 1 wherein the tangs of the first part of the second connector means are configured to follow walls of the radially extending recess and define a straight sided slot between the tangs.

3. The turbine stage assembly as claimed in claim 2 wherein the cutaway root portion is configured to result in an inclined and/or curved face on the radially opposing wall of the cutaway root portion.

4. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 3 .

5. The turbine stage assembly as claimed in claim 2 wherein the first part of the second connector means is integrally cast into the blade.

6. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 5 .

7. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 2 .

8. The turbine stage assembly as claimed in claim 1 wherein the cutaway root portion is configured to result in an inclined and/or curved face on the radially opposing wall of the cutaway root portion.

9. The turbine stage assembly as claimed in claim 8 wherein the first part of the second connector means is integrally cast into the blade.

10. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 9 .

11. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 8 .

12. The turbine stage assembly as claimed in claim 1 wherein the first part of the second connector means is integrally cast into the blade.

13. The gas turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 12 .

14. The turbine engine comprising one or more turbine stage assemblies, the turbine stage assemblies having a configuration according to claim 1 .

Assignees

• Rolls Royce Plc

Inventors

• Dawson John
• Burford Peter C

Classifications

• F05D2230/31

  Layer deposition · CPC title

• F05D2260/20

  Heat transfer, e.g. cooling · CPC title

• F01D5/3007

  of axial insertion type · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2230/21

  by casting · CPC title