Gas turbine engine cooling arrangement

Patent metadata

Field	Value
Publication number	US-10087775-B2
Application number	US-201414199099-A
Country	US
Kind code	B2
Filing date	Mar 6, 2014
Priority date	Mar 26, 2013
Publication date	Oct 2, 2018
Grant date	Oct 2, 2018

Abstract

A gas turbine engine comprises a compressor, a combustion chamber, an outer casing, an inner casing and a cooling arrangement. The outer casing surrounds the compressor and the combustion chamber and the combustion chamber has turbine nozzle guide vanes. The compressor has load carrying outlet guide vanes connected to the outer casing and the inner casing. The turbine nozzle guide vanes connect the outer casing and the inner casing. The cooling arrangement comprises a cooling air duct located between the compressor and the combustion chamber. The compressor outlet guide vanes carry at least one aerodynamic fairing. A support structure supports the cooling air duct from the inner casing at two spaced positions and the support structure forms a chamber with the inner casing. The support structure comprises at least one hollow duct and each hollow duct locates behind a respective one of the aerodynamic fairings.

First claim

The invention claimed is:

1. A gas turbine engine comprising a compressor, a diffuser, a combustion chamber, an outer casing, an inner casing and a cooling arrangement,
the outer casing surrounding the compressor, the diffuser and the combustion chamber,
the combustion chamber having turbine nozzle guide vanes,
the compressor having structural load carrying compressor outlet guide vanes connected to the outer casing and the inner casing, the compressor outlet guide vanes extending radially between and in contact with an inner wall and an outer wall, the inner wall and the outer wall diverging away from each other to form the diffuser,
the turbine nozzle guide vanes being connected to the inner casing to transmit structural loads via the compressor outlet guide vanes to the outer casing,
the cooling arrangement comprising a cooling air duct located between the compressor and the combustion chamber, the cooling arrangement being located radially between the outer wall and the outer casing,
the diffuser being arranged axially between the compressor and the combustion chamber,
the compressor outlet guide vanes supporting at least one aerodynamic fairing, the at least one aerodynamic fairing extending radially across the diffuser, the at least one aerodynamic fairing extending radially between the inner wall and the outer wall,
a first support structure to support the cooling air duct from the inner casing, the first support structure forming a chamber with the inner casing, the first support structure comprising at least one hollow duct, each hollow duct at least one hollow duct being located within and behind a respective one of the at least one aerodynamic fairings, the at least one hollow duct extending radially across the diffuser from the cooling air duct to the first support structure, the at least one hollow duct extending through the respective one of the at least one aerodynamic fairings, and
the cooling air duct, the at least one hollow duct, and the support structure being free to expand and contract independently of the inner casing.

2. The gas turbine engine as claimed in claim 1 wherein the compressor outlet guide vanes carries a plurality of aerodynamic fairings and the support structure comprises a plurality of hollow ducts.

3. The gas turbine engine as claimed in claim 1 wherein the cooling air duct comprises an annular manifold and the support structure is arranged around and surrounds the inner casing to form an annular chamber with the inner casing.

4. The gas turbine engine as claimed in claim 1 wherein the cooling air duct comprises a part annular manifold and the support structure is arranged around and surrounds the inner casing to form an annular chamber with the inner casing.

5. The gas turbine engine as claimed in claim 1 wherein
the inner casing comprises at least one aperture to supply cooling air to cool a component, and
the component is selected from the group consisting of a combustion chamber component and a turbine component.

6. The gas turbine engine as claimed in claim 5 wherein the turbine component comprises a turbine disc and the at least one aperture in the inner casing is arranged to supply cooling air from the chamber to cool the turbine disc.

7. The gas turbine engine as claimed in claim 1 wherein the support structure comprises a first end and a second end.

8. The gas turbine engine as claimed in claim 7 wherein
the inner casing comprises a flange, and
the second end of the support structure comprises a flange that is secured to the flange of the inner casing.

9. The gas turbine engine as claimed in claim 7 wherein
the inner casing comprises a slot, and
the second end of the support structure is located in the slot in the inner casing to form a sliding joint.

10. The gas turbine engine as claimed in claim 7 further comprising a ring seal provided between the second end of the support structure and the inner casing.

11. The gas turbine engine as claimed in claim 10 wherein the ring seal is selected from the group consisting of a piston ring seal, a brush seal and a labyrinth seal.

12. The gas turbine engine as claimed in claim 7 further comprising a ring seal provided between the first end of the support structure and the inner casing.

13. The gas turbine engine as claimed in claim 12 wherein the ring seal is selected from the group consisting of a piston ring seal, a brush seal and a labyrinth seal.

14. The gas turbine engine as claimed in claim 1 wherein the cooling air duct is supported from the outer casing.

15. The gas turbine engine as claimed in claim 1 wherein
the cooling arrangement comprises a supply duct, a heat exchanger and a return duct, where the supply duct is arranged to supply air from the compressor to the heat exchanger, and the return duct is arranged to return cool air from the heat exchanger to the cooling air duct.

16. The gas turbine engine as claimed in claim 15 , wherein
the gas turbine engine is a turbofan gas turbine engine having a bypass duct and the heat exchanger is arranged to exchange heat with air in the bypass duct of the turbofan gas turbine engine.

17. The gas turbine engine as claimed in claim 15 , further comprising a fuel system, wherein the heat exchanger is arranged to exchange heat with fuel in the fuel system of the gas turbine engine.

18. The gas turbine engine as claimed in claim 1 wherein the at least one aerodynamic fairing has an open downstream end.

19. A gas turbine engine comprising a compressor, a diffuser, a combustion chamber, an outer casing, an inner casing and a cooling arrangement,
the outer casing surrounding the compressor, the diffuser and the combustion chamber,
the combustion chamber having turbine nozzle guide vanes,
the compressor having structural load carrying compressor outlet guide vanes connected to the outer casing and the inner casing, the compressor outlet guide vanes extending radially between and in contact with an inner wall and an outer wall, the inner wall and the outer wall diverging away from each other to form the diffuser,
the turbine nozzle guide vanes being connected to the inner casing to transmit structural loads via the compressor outlet guide vanes to the outer casing,
the cooling arrangement comprising at least a part annular manifold located axially between the compressor and the combustion chamber, the at least a part annular manifold being located radially between the outer wall and the outer casing,
the diffuser being arranged axially between the compressor and the combustion chamber,
the compressor outlet guide vanes supporting at least one aerodynamic fairing, the at least one aerodynamic fairing extending radially across the diffuser, the at least one aerodynamic fairing extending radially between the inner wall and the outer wall,
a first support structure to support the at least a part annular manifold from the inner casing, the first support structure forming a chamber with the inner casing, the first support structure comprising at least one hollow duct, each hollow duct of the at least one hollow duct being located within and behind a respective one of the at least one aerodynamic fairings, the at least one hollow duct extending radially across the diffuser from the at least a part annular manifold to the first support structure, the at least one hollow duct extending radially through the respective one of the at least one aerodynamic fairings, and
the cooling air duct, the at least a part annular manifold, the at least one hollow duct, and the first sup

Assignees

• Rolls Royce Plc

Inventors

• Mulcaire Thomas Gerard

Classifications

• Y02T50/671

  Cross-Sectional Technologies · mapped topic

• Y02T50/676

  Cross-Sectional Technologies · mapped topic

• F02C7/185

  Cooling means for reducing the temperature of the cooling air or gas · CPC title

• F05D2260/213

  by the provision of a heat exchanger within the cooling circuit · CPC title

• F05D2240/90

  Mounting on supporting structures or systems · CPC title