Compressor secondary flow aft cone cooling scheme

Patent metadata

Field	Value
Publication number	US-2017292532-A1
Application number	US-201615094583-A
Country	US
Kind code	A1
Filing date	Apr 8, 2016
Priority date	Apr 8, 2016
Publication date	Oct 12, 2017
Grant date	—

Abstract

The present disclosure provides an axial flow compressor comprising a high pressure compressor section having a core flow path, an aft stage and a forward stage; a diffuser in fluid communication with the core flow path and coupled to the aft stage; a plenum coupled to the diffuser; a pre-swirl nozzle coupled to the plenum, an exit of the pre swirl nozzle being directed at an aft stage rotor disk and configured to impart a swirl to a cooling fluid. The axial flow compressor further may further comprise an aft stage rotor cavity defined by a portion of the aft stage rotor disk and having an aft stage axial overlap seal, wherein a portion of the cooling fluid returns to the core flow path though the aft stage labyrinth seal. The present disclosure provides a method of high pressure compressor aft stage cooling.

First claim

What is claimed is:

1 . An axial flow compressor comprising:
a high pressure compressor section having a core flow path, an aft stage, and a forward stage; a diffuser in fluid communication with the core flow path coupled to the aft stage; a plenum coupled to the diffuser; a pre-swirl nozzle coupled to the plenum, an exit of the pre-swirl nozzle being directed at a rotor disk of the aft stage and configured to impart a swirl to a cooling fluid.

2 . The axial flow compressor of claim 1 , further comprising an aft stage rotor cavity defined by a portion of the rotor disk and having an aft stage axial overlap seal,
wherein a portion of the cooling fluid returns to the core flow path though the aft stage axial overlap seal.

3 . The axial flow compressor of claim 2 , wherein the aft stage rotor cavity further comprises an aft cone wherein a portion of the cooling fluid travels along the aft cone and exits the high pressure compressor section through a labyrinth seal.

4 . The axial flow compressor of claim 1 , further comprising a forward stage axial overlap seal, wherein the cooling fluid returns to the core flow path through the forward stage axial overlap seal.

5 . The axial flow compressor of claim 1 , wherein the rotor disk is at least one of a segmented bladed disk or an integrally bladed disk having cooling slots.

6 . The axial flow compressor of claim 5 , wherein the rotor disk is in fluid communication with the pre-swirl nozzle and configured to pass the cooling fluid from the pre-swirl nozzle through the rotor disk to the forward stage.

7 . The axial flow compressor of claim 1 , wherein the pre-swirl nozzle, the aft stage, and the forward stage are in fluid communication.

8 . The axial flow compressor of claim 1 , wherein the plenum further comprises a heat exchanger in fluid communication with the pre-swirl nozzle.

9 . The axial flow compressor of claim 3 , wherein the labyrinth seal is at least one of integrated with the aft cone or coupled to the aft cone.

10 . The axial flow compressor of claim 9 , wherein a portion of the cooling fluid exits through the labyrinth seal.

11 . The axial flow compressor of claim 1 , wherein the swirl coincides with a rotation of the rotor disk.

12 . The axial flow compressor of claim 1 , wherein the pre-swirl nozzle comprises a least one of steel, stainless steel, nickel, nickel alloy, titanium, or titanium alloy.

13 . A gas turbine engine comprising:
an axial flow compressor having a core flow path; a combustor; a diffuser coupled between the axial flow compressor and the combustor; a plenum coupled to the diffuser; and a pre-swirl nozzle coupled to the plenum, an exit of the pre-swirl nozzle being directed at an aft stage rotor disk and configured to impart a swirl to a cooling fluid.

14 . The gas turbine engine of claim 13 , wherein the diffuser comprises an airfoil disposed within the core flow path.

15 . The gas turbine engine of claim 14 , wherein the airfoil comprises an aperture proximate a trailing edge of the airfoil.

16 . The gas turbine engine of claim 13 , wherein the aft stage rotor disk comprises at least one of a segmented bladed disk or integrally bladed disk having cooling slots.

17 . A method of high pressure compressor aft stage cooling comprising:
drawing a coolant from a core flow path of a gas turbine engine, wherein the coolant is drawn from the core flow path between an exit of a high pressure compressor and an entrance of a combustor; feeding the coolant through a pre-swirl nozzle, wherein the pre-swirl nozzle exit is directed at an aft stage rotor disk of the high pressure compressor; and returning the coolant to the core flow path through an axial overlap seal.

18 . The method of claim 17 , further comprising directing a portion of the coolant along an aft stage cone,
wherein the aft stage cone is coupled to a labyrinth seal, wherein the portion of coolant exits through the labyrinth seal.

19 . The method of claim 17 , further comprising directing a portion of the coolant forward through the aft stage rotor disk to a forward stage and returning the portion of coolant to the core flow path through a forward stage axial overlap seal.

20 . The method of claim 17 , further comprising reducing the temperature of the coolant prior to feeding the coolant through the pre-swirl nozzle.

Assignees

• United Technologies Corp

Inventors

• Wall Jordan T

Classifications

• F01D11/001

  for sealing space between stator blade and rotor · CPC title

• F04D29/083

  especially adapted for elastic fluid pumps · CPC title

• F05D2240/126

  Baffles or ribs · CPC title

• F04D29/321

  for axial flow compressors · CPC title

• F04D29/541

  Specially adapted for elastic fluid pumps (F04D29/56 takes precedence) · CPC title