Gas turbine engine with integral combustion liner and turbine nozzle

Patent metadata

Field	Value
Publication number	US-11248789-B2
Application number	US-201816212879-A
Country	US
Kind code	B2
Filing date	Dec 7, 2018
Priority date	Dec 7, 2018
Publication date	Feb 15, 2022
Grant date	Feb 15, 2022

Abstract

A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

First claim

What is claimed is:

1. A core engine article, comprising:
a combustor liner defining a combustion chamber therein, the combustor liner including a plurality of injector ports, the plurality of injector ports having a shape that tapers to a corner on a side of the injector ports; and
a turbine nozzle, the turbine nozzle including a plurality of airfoils, an inner annulus, and an outer annulus, the inner annulus extends into the combustion chamber, and the combustor liner is arranged between the inner annulus and the outer annulus;
wherein the combustor liner and turbine nozzle are integral with one another.

2. The core engine article of claim 1 , wherein the fuel injector ports are diamond-shaped.

3. The core engine article of claim 1 , further comprising a webbing extending outward from the combustor liner along an extent of a periphery of the fuel injector ports.

4. The core engine article of claim 3 , wherein the webbing is configured to support a fuel injector.

5. The core engine article of claim 1 , wherein the fuel injector ports extend into the combustion chamber.

6. The core engine article of claim 1 , wherein a rib extending from the inner annulus contacts the combustor liner.

7. The core engine article of claim 1 , wherein the airfoils extend from an exterior of the outer annulus.

8. The core engine article of claim 1 , wherein an end of the combustor liner includes one or more scallops.

9. The core engine article of claim 8 , wherein the scallops are configured to accelerate air flowing through the core engine article.

10. The core engine article of claim 8 , wherein the scallops are configured to accelerate airflow through the core engine article in a direction that is a reverse of the direction of the airflow.

11. The core engine article of claim 1 , wherein the airfoils are each arranged along an airfoil axis, and an angle α between the airfoil axis and a central axis of the core engine article is 45 degrees.

12. The core engine article of claim 1 , wherein each of the plurality of fuel injector ports are configured to receive atomized fuel into the combustion chamber.

13. The core engine article of claim 1 , wherein the combustion chamber is configured to receive airflow in a direction parallel to a central axis of the core engine article.

Assignees

• United Technologies Corp
• Raytheon Tech Corp

Inventors

• Binek Lawrence
• Boyer Jesse R
• Voytovych Dmytro Mykolayovych

Classifications

• F23R3/002 Primary

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

• B22F10/25

  Direct deposition of metal particles, e.g. direct metal deposition [DMD] or laser engineered net shaping [LENS] · CPC title

• F23R3/30

  comprising fuel prevapourising devices · CPC title

• F05D2250/82

  Micromachines · CPC title

• F05D2230/31

  Layer deposition · CPC title