Dual cooling airflow to blades

Patent metadata

Field	Value
Publication number	US-2018291744-A1
Application number	US-201715483752-A
Country	US
Kind code	A1
Filing date	Apr 10, 2017
Priority date	Apr 10, 2017
Publication date	Oct 11, 2018
Grant date	—

Abstract

An airfoil may comprise a root and an airfoil body radially outward of the root. The airfoil body may define a first cooling chamber and a second cooling chamber. A first passage may be defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber. A second passage may be defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber.

First claim

What is claimed is:

1 . An airfoil, comprising:
a root; an airfoil body radially outward of the root, the airfoil body defining a first cooling chamber and a second cooling chamber; a first passage defined within the root and configured to direct a first airflow radially outward through the root into the first cooling chamber; and a second passage defined within the root and configured to direct a second airflow radially outward through the root and into the second cooling chamber.

2 . The airfoil of claim 1 , further comprising a first inlet defined in an axially forward surface of the root.

3 . The airfoil of claim 2 , further comprising a second inlet defined in an axially aft surface of the root.

4 . The airfoil of claim 3 , further comprising a leading edge and a trailing edge, wherein the first cooling chamber is disposed at the leading edge and the second cooling chamber is disposed at the trailing edge.

5 . The airfoil of claim 4 , wherein the airfoil body defines a first plurality of holes at the leading edge, and wherein the first airflow is directed out the airfoil through the first plurality of holes.

6 . The airfoil of claim 5 , wherein the airfoil body defines a second plurality of holes at the trailing edge, and wherein the second airflow is directed out the airfoil through the second plurality of holes.

7 . The airfoil of claim 1 , wherein the airfoil comprises a blade.

8 . An engine section of a gas turbine engine, comprising:
a disk configured to rotate about an axis; a blade coupled to the disk, the blade defining a first cooling chamber and a second cooling chamber; a first flow guide disposed upstream of the disk, the first flow guide defining a first airflow path and a second airflow path; a first passage configured to direct the first airflow path to the first cooling chamber; and a second passage configured to direct the second airflow path to the second cooling chamber.

9 . The engine section of claim 8 , further comprising a first tangential onboard injector (TOBI) disposed in the first airflow path.

10 . The engine section of claim 9 , wherein the first passage includes a first inlet defined in an axially forward surface of the disk, the first TOBI configured to direct the first airflow path into the first inlet.

11 . The engine section of claim 10 , further comprising a second TOBI disposed in the second airflow path, wherein the second passage includes a second inlet defined in the axially forward surface of the disk, the second TOBI configured to direct the second airflow path into the second inlet.

12 . The engine section of claim 8 , wherein the first passage includes a first inlet defined in an axially forward surface of a root, and wherein the second passage includes a second inlet defined in the axially forward surface of the disk.

13 . The engine section of claim 8 , further comprising a radial onboard injector (ROBI) disposed in the second airflow path forward of the disk and configured to direct the second airflow path radially inward.

14 . The engine section of claim 13 , further comprising a minidisk coupled to an axially aft surface of the disk, wherein the second airflow path is directed radially outward between the minidisk and the axially aft surface of the disk.

15 . The engine section of claim 8 , wherein the blade defines a first plurality of holes at a leading edge, and wherein the first airflow path is directed out the blade through the first plurality of holes.

16 . A gas turbine engine, comprising:
a compressor section configured to supply airflow to at least one of a first airflow path or a second airflow path; a turbine section configured to receive airflow from the first airflow path and the second airflow path, the turbine section comprising:
a disk configured to rotate about an axis,
a blade coupled to the disk, the blade defining a first cooling chamber and a second cooling chamber,
a first passage configured to direct a first airflow path to the first cooling chamber; and
a second passage configured direct the second airflow path to the second cooling chamber.

17 . The gas turbine engine of claim 16 , further comprising:
a first flow guide disposed upstream of the disk, the first flow guide defining the first airflow path and the second airflow path; and a first tangential onboard injector (TOBI) disposed in the first airflow path, wherein the first passage includes a first inlet defined in an axially forward surface of the disk, the first TOBI configured to direct the first airflow path into the first inlet.

18 . The gas turbine engine of claim 17 , further comprising a second TOBI disposed in the second airflow path, wherein the second passage includes a second inlet defined in the axially forward surface of the disk, the second TOBI configured to direct the second airflow path into the second inlet.

19 . The gas turbine engine of claim 17 , further comprising a radial onboard injector (ROBI) disposed in the second airflow path forward of the disk and configured to direct the second airflow path radially inward.

20 . The gas turbine engine of claim 19 , further comprising a minidisk coupled to an axially aft surface of the disk, wherein the second airflow path is directed radially outward between the minidisk and the axially aft surface of the disk.

Assignees

• United Technologies Corp

Inventors

• Devore Matthew A
• Ortiz Jonathan

Classifications

• F05D2240/81

  Cooled platforms · CPC title

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title

• F05D2260/201

  by impingement of a fluid · CPC title

• F01D5/081 Primary

  Cooling fluid being directed on the side of the rotor disc or at the roots of the blades (F01D5/087 takes precedence) · CPC title

• F01D5/187

  Convection cooling · CPC title