Engine component having support with intermediate layer

Patent metadata

Field	Value
Publication number	US-10590780-B2
Application number	US-201414782014-A
Country	US
Kind code	B2
Filing date	Apr 1, 2014
Priority date	Apr 2, 2013
Publication date	Mar 17, 2020
Grant date	Mar 17, 2020

Abstract

Disclosed is a gas turbine engine component, and a method for forming the component. The component includes a first portion, a second portion formed separately from the first portion, and an intermediate layer provided between the first portion and the second portion.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
a compressor section, a combustor section, and a turbine section, the turbine section including a stationary stage and a rotating stage;
a component provided in one of the stationary stage and the rotating stage, the component including a first portion, wherein the first portion includes an airfoil section and a root of a turbine blade, a second portion formed separately from the first portion and at least partially disposed radially outward of a rotor disc of the turbine section, and an intermediate layer provided between the first portion and the second portion, wherein, in an initial condition before any wear or de-bonding of the intermediate layer, the first portion abuts the second portion indirectly via the intermediate layer at a location, and wherein, after the initial condition, the intermediate layer wears and becomes de-bonded from the first portion at least at the location such that the first portion directly abuts the second portion at the location.

2. The gas turbine engine as recited in claim 1 , wherein the component is the turbine blade, and is provided in the rotating stage of the turbine section.

3. The gas turbine engine as recited in claim 2 , wherein the turbine section includes a high pressure section and a low pressure section, the turbine blade being provided in the high pressure section.

4. The gas turbine engine as recited in claim 2 , wherein the turbine blade is made of a ceramic matrix composite (CMC) material.

5. The gas turbine engine as recited in claim 2 , wherein the second portion of the turbine blade is a platform.

6. The gas turbine engine as recited in claim 5 , wherein an outer face of the platform directly abuts an inner face of a slot of the rotor disc.

7. The gas turbine engine as recited in claim 1 , wherein the second portion is at least partially moveable relative to the first portion of the component during operation of the gas turbine engine.

8. The gas turbine engine as recited in claim 1 , wherein, in the initial condition, the intermediate layer prevents the first portion from directly contacting the second portion.

9. The gas turbine engine as recited in claim 8 , wherein, in the initial condition, the intermediate layer has a thickness between 0.005 to 0.0005 inches (0.00127 to 0.0127 cm).

10. A gas turbine engine component, comprising:
a first portion, wherein the first portion includes an airfoil section and a root of a turbine blade;
a second portion formed separately from the first portion, wherein the second portion includes a platform of the turbine blade, the second portion at least partially disposed radially outward of a rotor disc of a turbine section; and
an intermediate layer provided between the first portion and the second portion, wherein, in an initial condition before any wear or de-bonding of the intermediate layer, the first portion abuts the second portion indirectly via the intermediate layer at a location, and wherein, after the initial condition, the intermediate layer wears and becomes de-bonded from the first portion at least at the location such that the first portion directly abuts the second portion at the location.

11. The gas turbine engine as recited in claim 10 , wherein the first portion is at least partially moveable relative to the second portion of the component.

12. A method of forming a component for a gas turbine engine comprising:
forming a first portion of the component;
providing an intermediate layer over the component; and
forming a second portion of the component over the intermediate layer, wherein the first portion of the component includes an airfoil section and a root of a turbine blade, and the second portion of the component is a supporting structure configured to support the first portion and includes a platform, the second portion at least partially disposed radially outward of a rotor disc of a turbine section, and wherein, in an initial condition before any wear or de-bonding of the intermediate layer, the first portion abuts the second portion indirectly via the intermediate layer at a location, and wherein, after the initial condition, the intermediate layer wears and becomes de-bonded from the first portion at least at the location such that the first portion directly abuts the second portion at the location.

13. The method as recited in claim 12 , wherein the component is a turbine blade.

14. The method as recited in claim 12 , wherein the intermediate layer is provided over at least a portion of the root.

15. The method as recited in claim 12 , wherein the airfoil section and root are formed by molding ceramic matrix composite (CMC) material.

16. The method as recited in claim 12 , wherein the platform is formed by molding ceramic matrix composite (CMC) material.

17. The method as recited in claim 16 , wherein the intermediate layer includes one of carbon, boron nitride, and silicon.

18. A gas turbine engine, comprising:
a compressor section, a combustor section, and a turbine section, the turbine section including a stationary stage and a rotating stage;
a component provided in one of the stationary stage and the rotating stage, the component including a first portion, a second portion formed separately from the first portion, and an intermediate layer provided between the first portion and the second portion, wherein the intermediate layer wears and becomes at least partially de-bonded from the first portion during operation of the gas turbine engine, and
wherein the intermediate layer becomes completely worn away during operation of the gas turbine engine.

Assignees

• United Technologies Corp

Inventors

• Mccaffrey Michael G

Classifications

• F05D2240/80

  Platforms for stationary or moving blades · CPC title

• F02C3/04

  having a turbine driving a compressor (power transmission arrangements F02C7/36; control of working fluid flow F02C9/16) · CPC title

• F01D9/041

  using blades (F01D5/148 takes precedence) · CPC title

• C04B2237/38

  Fiber or whisker reinforced · CPC title

• C04B2237/08

  Non-oxidic interlayers · CPC title