Combustor size rating for a gas turbine engine using hydrogen fuel

Patent metadata

Field	Value
Publication number	US-12460573-B2
Application number	US-202418673577-A
Country	US
Kind code	B2
Filing date	May 24, 2024
Priority date	Dec 3, 2021
Publication date	Nov 4, 2025
Grant date	Nov 4, 2025

Abstract

A gas turbine engine includes a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow, a compressor section configured to compress air flowing therethrough to provide a compressed air flow, and a combustor including a combustion chamber having a burner length and a burner dome height. The combustion chamber is configured to combust a mixture of the hydrogen fuel flow and the compressed air flow. The combustion chamber can be characterized by a combustor size rating between one inch and seven inches. In more detail, the combustion chamber can be characterized by the combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, in which the combustor size rating is a function of the core air flow parameter.

First claim

What is claimed is:

1 . A gas turbine engine comprising:
a hydrogen fuel delivery assembly configured to deliver a hydrogen fuel flow; a compressor section configured to compress air flowing therethrough to provide a compressed air flow; and a combustor configured to operate without diluent, the combustor including a fuel premixer, an inner liner, an outer liner, and a combustion chamber, the combustion chamber characterized by a combustor size rating between one inch and seven inches at a core air flow parameter between two and one half kN and sixty kN, wherein the combustor size rating is a function of the core air flow parameter, and wherein the combustor size rating is defined by:

L
2

H

wherein His a maximum height of the combustion chamber measured by a forward line extending from an inner surface of the outer liner to an inner surface of the inner liner and L is a length of the combustion chamber measured from a midpoint of the forward line to a midpoint of an aft line, the aft line extending from the inner surface of the inner liner to the inner surface of the outer liner at a leading edge of a turbine nozzle,
and, wherein the core air flow parameter is defined by:

Thrust

Bypass
⁢

Ratio

;

and
the fuel premixer comprising an annular shroud defining an interior with a mixing chamber, a center body located within the interior and having a first fuel nozzle emitting fuel into the mixing chamber, and an annular swirler located within the interior, circumscribing the center body, with the annular swirler spaced from the annular shroud to define an outer annular passage, and also spaced from the center body to define an inner annular passage.

2 . The gas turbine engine of claim 1 , wherein the outer annular passage is configured to emit a first non-swirling airflow into the mixing chamber as a first air film located along the annular shroud, the inner annular passage is configured to emit a second non-swirling airflow into the mixing chamber as a second air film located along the center body, the annular swirler is configured to emit a swirling airflow into the mixing chamber between the first non-swirling airflow and the second non-swirling airflow and the first fuel nozzle is configured to emit a swirling fuel flow into the mixing chamber.

3 . The gas turbine engine of claim 2 , further comprising a set of second fuel nozzles configured to emit a second fuel flow into the swirling airflow.

4 . The gas turbine engine of claim 3 , further comprising a second fuel passage located within the interior and fluidly coupled to the set of second fuel nozzles.

5 . The gas turbine engine of claim 4 , wherein the annular swirler comprises a set of vanes, with the set of second fuel nozzles positioned on the set of vanes.

6 . The gas turbine engine of claim 5 , wherein the set of vanes comprises a trailing edge, with the second fuel passage located one of downstream or upstream of the trailing edge.

7 . The gas turbine engine of claim 4 , wherein the second fuel passage extends to at least one outlet on at least one wall of the annular swirler, with the at least one outlet fluidly coupling the second fuel passage to at least one of the inner annular passage or the outer annular passage.

8 . The gas turbine engine of claim 4 , wherein the second fuel passage extends into a wall of the annular swirler, with the set of second fuel nozzles configured to emit the second fuel flow outside of the annular swirler.

9 . The gas turbine engine of claim 2 , wherein the center body comprises a nozzle cap and a set of fuel orifices in the nozzle cap configured to emit the swirling fuel flow.

10 . The gas turbine engine of claim 1 , further comprising a third fuel nozzle in the center body fluidly coupling a first fuel passage in the first fuel nozzle to the inner annular passage.

11 . The gas turbine engine of claim 1 , further comprising a cooling aperture in the annular shroud downstream of the annular swirler and fluidly coupled to the mixing chamber.

12 . The gas turbine engine of claim 1 , wherein the combustor size rating is between two inches and three and one quarter inches at the core air flow parameter between two and one half kN and fifty kN.

13 . The gas turbine engine of claim 1 , wherein the combustor size rating is based on a thrust of the gas turbine engine.

14 . The gas turbine engine of claim 13 , wherein the thrust is between sixty kN and five hundred kN.

15 . The gas turbine engine of claim 1 , wherein the annular swirler comprises a set of vanes.

16 . The gas turbine engine of claim 15 , wherein the annular swirler has a first annular wall and a second annular wall, wherein the outer annular passage is a first annular passage defined by the first annular wall and the annular shroud, wherein a second annular passage is defined between the first annular wall and a second annular wall of the annular swirler, and the inner annular passage is a third annular passage defined by the second annular wall and the center body.

17 . A method of mixing fuel in the combustor of the gas turbine engine of claim 1 , the method comprising:
emitting a first non-swirling airflow into the mixing chamber within the combustor; emitting a second non-swirling airflow into the mixing chamber and spaced from the first non-swirling airflow; emitting a first swirling airflow into the mixing chamber between the first non-swirling airflow and the second non-swirling airflow; and emitting a first flow of the fuel into the combustor adjacent the second non-swirling airflow.

18 . The method of claim 17 , further comprising emitting a second flow of fuel into the first swirling airflow.

19 . The method of claim 18 , wherein the second flow of fuel is emitted from at least one vane of the annular swirler.

20 . The method of claim 19 , wherein the second flow of fuel is emitted from fuel orifices upstream of a trailing edge of the at least one vane.

Assignees

• Gen Electric

Inventors

• Benjamin Michael A
• Giridharan Manampathy G

Classifications

• F23R3/002

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

• F05D2220/323

  for aircraft propulsion, e.g. jet engines · CPC title

• F23R3/286

  having fuel-air premixing devices (F23R3/30 takes precedence) · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F02C7/22

  Fuel supply systems · CPC title