Combustor particulate deflector

Patent metadata

Field	Value
Publication number	US-11262071-B2
Application number	US-201916520853-A
Country	US
Kind code	B2
Filing date	Jul 24, 2019
Priority date	Jul 24, 2019
Publication date	Mar 1, 2022
Grant date	Mar 1, 2022

Abstract

A gas turbine engine, including: a diffuser case defining an inner shroud and an outer shroud; and a combustor housed within the diffuser case between the inner shroud and the outer shroud, the combustor including: a shell; a forward dome attached to the shell at a forward end of the combustor; and a deflector attached to the forward dome and extending away from the forward dome.

First claim

What is claimed is:

1. A gas turbine engine, comprising:
a diffuser case defining an inner plenum and an outer plenum, the inner plenum being radially inward of the outer plenum relative to a longitudinal axis of the gas turbine engine; and
a combustor housed within the diffuser case between the inner plenum and the outer plenum, the combustor comprising:
a radially outward shell;
a radially inward shell;
a forward dome attached to the radially outward shell and the radially inward shell at a forward end of the combustor; and
a deflector attached to the forward dome and extending axially aft and radially inward away from the forward dome and the combustor,
wherein the forward dome further comprises:
a curved dome portion; and
a radially inward linear portion,
the forward dome being attached to the radially inward shell at the radially inward linear portion,

wherein the forward dome transitions from the curved dome portion to the radially inward linear portion at a radially inward transition point,
wherein the deflector is attached to the forward dome proximate the radially inward transition point,
wherein the deflector is oriented at an angle relative to the longitudinal axis of the gas turbine engine about equal to an angle of the curved dome portion proximate the radially inward transition point relative to the longitudinal axis of the gas turbine engine, and
wherein the deflector is linear, and the angle relative to the longitudinal axis of the gas turbine engine is maintained across the entire length of the deflector.

2. The gas turbine engine of claim 1 , wherein the deflector extends into the inner plenum.

3. The gas turbine engine of claim 1 , wherein the deflector is attached to a radially inward side of the forward dome.

4. The gas turbine engine of claim 1 , wherein the deflector is a solid body.

5. The gas turbine engine of claim 1 , wherein the deflector further comprises:
a radially inward forward side;
a radially outward aft side opposite the radially inward forward side; and
one or more orifices extending from the radially inward forward side to the radially outward aft side.

6. The gas turbine engine of claim 5 , wherein the one or more orifices are oriented at an angle less than or equal to 90 degrees relative to the radially inward forward side.

7. The gas turbine engine of claim 5 , wherein the one or more orifices each include corners that are rounded in shape, each of the corners of each of the one or more orifices being adjacent to one of the radially inward forward side and the radially outward aft side.

8. The gas turbine engine of claim 1 , further comprising:
a catcher located opposite of the deflector on a radially inward wall of the diffuser case,
the catcher forming an internal chamber with the radially inward wall,
wherein the catcher comprises a forward opening to allow particulate to enter into the internal chamber.

9. The gas turbine engine of claim 8 , wherein the catcher further comprises a rear backstop located aft of the forward opening.

10. The gas turbine engine of claim 8 , further comprising:
a radially inward exit orifice in the radially inward wall,
the radially inward exit orifice being fluidly connected to the internal chamber of the catcher.

11. A combustor for use in a gas turbine engine, the combustor comprising:
a radially outward shell;
a radially inward shell, radially inward of the radially outward shell relative to a longitudinal axis of the gas turbine engine;
a forward dome attached to the radially outward shell and the radially inward shell at a forward end of the combustor; and
a deflector attached to the forward dome and extending axially aft and radially inward away from the forward dome and the combustor,
wherein the forward dome further comprises:
a curved dome portion; and
a radially inward linear portion, the forward dome being attached to the radially inward shell at the radially inward linear portion,
wherein the forward dome transitions from the curved dome portion to the radially inward linear portion at a radially inward transition point,
wherein the deflector is attached to the forward dome proximate the radially inward transition point, and

wherein the deflector is oriented at an angle relative to the longitudinal axis of the gas turbine engine about equal to an angle of the curved dome portion proximate the radially inward transition point relative to the longitudinal axis of the gas turbine engine, and
wherein the deflector is linear, and the angle relative to the longitudinal axis of the gas turbine engine is maintained across the entire length of the deflector.

12. The combustor of claim 11 , wherein the deflector is attached to a radially inward side of the forward dome.

13. The combustor of claim 11 , wherein the deflector is a solid body.

14. The combustor of claim 11 , wherein the deflector further comprises: a radially inward forward side; a radially outward aft side opposite the radially inward forward side; and one or more orifices extending from the radially inward forward side to the radially outward aft side.

15. The combustor of claim 14 , wherein the one or more orifices are oriented at an angle less than or equal to 90 degrees relative to the radially inward forward side.

16. The combustor of claim 14 , wherein the one or more orifices each include corners that are rounded in shape, each of the corners of each of the one or more orifices being adjacent to one of the radially inward forward side and the radially outward aft side.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Prenter Robin
• Lundgreen Ryan

Classifications

• F23R3/002 Primary

  Wall structures (F23R3/02 and F23R3/007 take precedence) · CPC title

• F23R3/04

  Air inlet arrangements · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F23R2900/00004

  Preventing formation of deposits on surfaces of gas turbine components, e.g. coke deposits · CPC title

• F23R3/42

  characterised by the arrangement or form of the flame tubes or combustion chambers · CPC title