Thermal shielding in a gas turbine

Patent metadata

Field	Value
Publication number	US-10443402-B2
Application number	US-201615258701-A
Country	US
Kind code	B2
Filing date	Sep 7, 2016
Priority date	Sep 21, 2015
Publication date	Oct 15, 2019
Grant date	Oct 15, 2019

Abstract

A turbine blade includes a labyrinth of internal channels for the circulation of coolant received through an inlet formed in a terminal portion of the blade root and leading to a duct-defined wall. A first passage intersects the duct and extends through the blade towards a tip. An end of the first passage is arranged to capture incoming coolant flow. A second passage intersects the duct at a position downstream of the first passage intersection. The duct and/or the passage intersections are configured to create a pressure drop in the duct in the direction from the inlet to the second passage intersection. In an axial direction, a duct wall terminates at a position between the inlet and the second passage intersection.

First claim

The invention claimed is:

1. A turbine blade having a body enclosing a labyrinth of internal channels for a circulation of coolant received through an inlet formed in a terminal portion of the blade root, the labyrinth comprising;
the inlet arranged on an axially upstream face of the terminal portion leading to a duct defined by a duct wall;
in use, a clearance space bounded by an external surface of the duct wall and a bucket groove of a disc hub in which the blade is carried, the clearance space creating a leakage path for air directed to the inlet;
a first passage intersecting the duct at a first passage intersection and extending through the blade body towards a tip of the blade, a proximal end of the first passage being arranged, in use, to capture incoming coolant flow; and
a second passage intersecting the duct at a second passage intersection at a position downstream of the first passage intersection, wherein:
the duct and/or the passage intersections are configured to create a pressure drop in the duct in the direction from the inlet to the second passage intersection; and
in an axial direction, the duct wall terminates at a position between the inlet and the second passage intersection so as to balance the pressure of coolant in the duct with the pressure of coolant in the leakage path thereby reducing a mass flow of coolant entering the leakage path in the clearance space.

2. The turbine blade as claimed in claim 1 wherein a second passage inlet to the second passage is provided at the second passage intersection, the second passage inlet having a cross section which is less than that of the second passage intersection.

3. The turbine as claimed in claim 1 wherein the first passage is a leading edge passage.

4. The turbine blade as claimed in claim 1 wherein the second passage is a trailing edge passage.

5. The turbine blade as claimed in claim 1 wherein a third passage joins the second passage to form two inlet routes to a multipass passage extending through a-mid-portion to a trailing edge portion of the blade body.

6. The turbine blade as claimed in claim 1 wherein the wall terminates to an upstream side of the second passage intersection.

7. The turbine blade as claimed in claim 1 wherein the wall terminates partway along the second passage intersection.

8. The turbine blade as claimed in claim 1 wherein the wall extends axially to a position which is from 50% to 85% of the axial length of the bucket groove.

9. The turbine blade as claimed in claim 6 wherein the wall extends to a position which is from 40% to 60% of the axial length of the bucket groove.

10. The turbine blade as claimed in claim 7 wherein the wall extends to a position which is from 70% to 90% of the axial length of the bucket groove.

11. The turbine blade as claimed in claim 1 wherein the duct and the inlet are formed integrally with the blade in a single casting process.

12. The turbine blade as claimed in claim 1 wherein the duct wall is defined by two or more components which are subsequently joined or fastened together.

13. The turbine blade as claimed in claim 12 wherein the duct wall is manufactured using an additive layer manufacturing method and is subsequently friction welded to a cast blade portion which defines the remainder of the duct wall.

14. The turbine blade as claimed in claim 12 wherein the duct and inlet are provided integrally with a lock plate secured to the blade and/or a disc having a bucket groove which, in use, carries the blade.

15. The turbine blade as claimed in claim 12 wherein the duct and inlet are provided in the form of an insert positioned in an assembly of the blade and a disc having a bucket groove which, in use, carries the blade, after the blade is received in the bucket groove.

16. The turbine blade as claimed in claim 12 wherein the duct and inlet are provided integrally with a seal plate secured to the disc and or a disc having a bucket groove which, in use, carries the blade.

17. A gas turbine engine comprising one or more discs having bucket grooves into which is located the turbine blade having the configuration according to claim 1 .

Assignees

• Rolls Royce Plc

Inventors

• Burford Peter C
• Dawson John
• Shah Parth

Classifications

• F01D5/3015 Primary

  with side plates · CPC title

• F05D2230/31

  Layer deposition · CPC title

• F01D5/082

  on the side of the rotor disc · CPC title

• F01D5/187 Primary

  Convection cooling · CPC title

• F05D2230/21

  by casting · CPC title