Thermal structure for outer diameter mounted turbine blades

Patent metadata

Field	Value
Publication number	US-10876407-B2
Application number	US-201715434658-A
Country	US
Kind code	B2
Filing date	Feb 16, 2017
Priority date	Feb 16, 2017
Publication date	Dec 29, 2020
Grant date	Dec 29, 2020

Abstract

The present disclosure is directed to a gas turbine engine defining a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction. The gas turbine engine defines a core flowpath extended generally along the longitudinal direction. The gas turbine engine includes a first turbine rotor. The first turbine rotor includes an annular outer band disposed outward of the core flowpath along the radial direction. The first turbine rotor further includes a plurality of airfoils coupled to an inner diameter of the outer band in which the plurality of airfoils are extended generally inward along the radial direction. The outer band defines a plurality of airfoil cooling passages in which the plurality of airfoil cooling passages are extended at least partially in the radial direction in fluid communication with the plurality of airfoils.

First claim

What is claimed is:

1. A gas turbine engine, wherein the gas turbine engine defines a radial direction, a circumferential direction, an axial centerline along a longitudinal direction, and an upstream end and a downstream end along the longitudinal direction, and further wherein the gas turbine engine defines a core flowpath extended generally along the longitudinal direction, the gas turbine engine comprising:
a first turbine rotor comprising an annular outer band disposed outward of the core flowpath along the radial direction, and wherein the first turbine rotor further comprises a plurality of airfoils coupled to an inner diameter of the outer band, the plurality of airfoils extended generally inward along the radial direction, and wherein the outer band defines a plurality of airfoil cooling passages, the plurality of airfoil cooling passages extended at least partially in the radial direction in fluid communication with the plurality of airfoils.

2. The gas turbine engine of claim 1 , wherein the outer band defines a first aperture and a second aperture, and wherein the first aperture defines a first aperture area and the second aperture defines a second aperture area, and further wherein the first aperture is defined at an outer diameter of the outer band and the second aperture is defined at an inner diameter of the outer band.

3. The gas turbine engine of claim 2 , wherein the first aperture area defines a greater area than the second aperture area.

4. The gas turbine engine of claim 2 , wherein each airfoil cooling passage of the plurality of airfoil cooling passages is defined between the first aperture and the second aperture.

5. The gas turbine engine of claim 4 , wherein the outer band comprises a first cooling passage wall and a second cooling passage wall at each of the plurality of airfoil cooling passages between the first aperture and the second aperture.

6. The gas turbine engine of claim 4 , wherein each of the first cooling passage walls and second cooling passage walls together define a serpentine structure between the outer diameter and inner diameter of the outer band.

7. The gas turbine engine of claim 4 , wherein the airfoil cooling passages defined between the first aperture and the second aperture define a decreasing volume of the cooling passage.

8. The gas turbine engine of claim 1 , wherein the outer band further defines a plurality of axial passages extended generally along the longitudinal direction.

9. The gas turbine engine of claim 1 , wherein the plurality of airfoils defines one or more cooling fluid orifices at an outer diameter of the airfoil.

10. The gas turbine engine of claim 1 , wherein each airfoil cooling passage of the plurality of airfoil cooling passages is extended at least partially along the circumferential direction in the same direction of rotation as the first turbine rotor along the circumferential direction.

11. The gas turbine engine of claim 1 , the engine further comprising:
a second turbine rotor interdigitated among the first turbine rotor along the longitudinal direction, wherein the second turbine rotor includes a plurality of second turbine airfoils extended outward in the radial direction.

12. The gas turbine engine of claim 11 , wherein the engine defines, in serial flow arrangement from the upstream end to the downstream end, a plurality of upstream airfoils of the first turbine rotor, the plurality of second turbine airfoils of the second turbine rotor, and the plurality of downstream airfoils of the first turbine rotor.

13. The gas turbine engine of claim 11 , the engine further comprising:
a combustion section disposed upstream of the first turbine rotor and the second turbine rotor.

14. The gas turbine engine of claim 13 , wherein the engine defines, in serial flow arrangement, the combustion section, a plurality of upstream airfoils of the first turbine rotor, and the plurality of second turbine airfoils of the second turbine rotor, and the plurality of downstream airfoils of the first turbine rotor.

15. The gas turbine engine of claim 1 , the engine further comprising:
a turbine casing surrounding the first turbine rotor along the longitudinal direction and the circumferential direction.

16. The gas turbine engine of claim 15 , wherein the turbine casing and the first turbine rotor together define a first seal interface disposed upstream of the plurality of airfoils of the first turbine rotor and a second seal interface disposed downstream of the plurality of airfoils of the first turbine rotor.

17. The gas turbine engine of claim 16 , wherein the turbine casing and the first turbine rotor define a cooling cavity between the first seal interface, the second seal interface, the turbine casing, and the outer band of the first turbine rotor.

18. The gas turbine engine of claim 16 , wherein the turbine casing comprises a plurality of shrouds disposed inwardly along the radial direction, and wherein the first turbine rotor comprises a plurality of knife edge seals disposed outwardly along the radial direction toward the plurality of shrouds.

19. The gas turbine engine of claim 18 , wherein an upstream portion of the plurality of shrouds and an upstream portion of the plurality of knife edge seals define the first seal interface, and wherein a downstream portion of the plurality of shrouds and a downstream portion of the plurality of knife edge seals define the second seal interface.

20. The gas turbine engine of claim 1 , further comprising a connecting airfoil disposed downstream of the plurality of airfoils, wherein the connecting airfoil is coupled to the outer band on a radially outward end, and wherein the connecting airfoil is coupled to a rotor on a radially inward end.

Assignees

• Gen Electric

Inventors

• Miller Brandon Wayne
• Clements Jeffrey Donald
• Waslo Daniel
• Kirk Joel Francis

Classifications

• F01D11/08

  for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20) · CPC title

• F05D2260/202

  by film cooling · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F01D5/187

  Convection cooling · CPC title

• F01D5/03

  Annular blade-carrying members having blades on the inner periphery of the annulus and extending inwardly radially, i.e. inverted rotors · CPC title