BLEED DUCT ASSEMBLY FOR A GAS TURBINE ENGINE

• US 20150052908A1
• Filed: 08/08/2014
• Published: 02/26/2015
• Est. Priority Date: 08/21/2013
• Status: Active Grant

• Current Assignee: Rolls-Royce Deutschland Limited & Company KG (Rolls-Royce Holdings Plc)
• Original Assignee: Rolls-Royce Deutschland Limited & Company KG (Rolls-Royce Holdings Plc)
• Inventors: TODOROVIC, Predrag
• Granted Patent: US 9,915,229 B2
• Time in Patent Office: Days
• US Class Current: 60/785
• CPC Class Codes: F01D 25/162 Bearing supports
• CPC Class Codes: F02C 6/08 the gas being bled from the gas-turbine compressor
• CPC Class Codes: F02C 9/18 by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages F02C3/113 takes precedence
• CPC Class Codes: F02K 3/02 in which part of the working fluid by-passes the turbine and combustion chamber
• CPC Class Codes: F02K 3/075 controlling flow ratio between flows
• CPC Class Codes: F05D 2210/30 Flow characteristics
• CPC Class Codes: F05D 2240/12 Fluid guiding means, e.g. vanes
• CPC Class Codes: F05D 2250/38 angled, e.g. sweep angle
• CPC Class Codes: F05D 2250/75 given by its similarity to a letter, e.g. T-shaped

Abstract

A bypass duct has a support unit comprising a pair of aerofoils arranged as an “A” frame. A bleed duct assembly is provided on the radially inner wall of the bypass duct annulus and the aerofoils project from the surface and extend across the annulus between the inner wall and an outer wall of the annulus. The aerofoils lean at an acute angle to the surface with the first flank facing toward the inner wall and adjoining a bleed duct opening. The bleed duct having a bleed duct passage and a submerged scoop.

Claims

1. A bleed duct assembly for a gas turbine bypass duct comprising

a circumferentially extending surface defining an inner wall of an annulus onto which a bleed duct opens at a bleed duct opening and an aerofoil projecting from the surface and extending across the annulus between the inner wall and an outer wall of the annulus, the aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge,

wherein the aerofoil is leant at an acute angle to the surface with the first flank facing toward the inner wall and adjoining the bleed duct opening.

2. A bleed duct assembly according to claim 1, wherein the aerofoil provides an axially extending edge to the bleed duct.

3. A bleed duct assembly according to claim 1, wherein the bleed duct opening has a leading edge at a first axial location and a trailing edge at a second axial location axially rearward of the first axial location

4. A bleed duct assembly according to claim 3, wherein the first axial location is at or before the axial location of the leading edge of the aerofoil.

5. A bleed duct assembly according to claim 3, wherein the second axial location is rearward of the axial location of the leading edge of the aerofoil and ahead of the axial location of the trailing edge of the aerofoil.

6. A bleed duct assembly according to claim 3, wherein the second axial location is at or aft of the axial location of the trailing edge of the aerofoil.

7. A bleed duct assembly according to claim 3 wherein the leading edge of the bleed duct opening and the trailing edge of the bleed duct opening extend substantially circumferentially, wherein the circumferential length of the leading edge is less than the circumferential length of the trailing edge.

8. A bleed duct assembly according to claim 1, wherein the bleed duct has a bleed duct passage and a scoop recessed radially inside the inner wall.

9. A bleed duct assembly according to claim 7, wherein the recessed scoop has a side wall, the first flank projecting radially inside the surface and providing at least part of the scoop side wall.

10. A support unit for a gas turbine engine, the support unit having a pair of aerofoils with a first aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge, the first aerofoil leaning at an acute angle to a circumferentially extending surface defining an inner wall of an annulus onto which a bleed duct opens at a bleed duct opening and wherein the first flank faces toward the surface;

and the second aerofoil having a leading edge, a trailing edge and first and second flanks connecting the leading edge and the trailing edge, the aerofoil leaning at an acute angle to the circumferentially extending surface with the first flank facing toward the surface, wherein the first flank of the first aerofoil and the first flank of the second aerofoil face towards each other;

wherein the first flank of the first aerofoil adjoins the bleed duct opening.

11. A support unit for a gas turbine engine according to claim 10, wherein the aerofoils extend across an annular bypass duct surrounding an engine core.

12. A gas turbine engine having a support unit according to claim 11.