Systems and methods for providing output products to a combustion chamber of a gas turbine engine

Patent metadata

Field	Value
Publication number	US-12037952-B2
Application number	US-202217568359-A
Country	US
Kind code	B2
Filing date	Jan 4, 2022
Priority date	Jan 4, 2022
Publication date	Jul 16, 2024
Grant date	Jul 16, 2024

Abstract

Systems and methods including a plurality of reformer stacks extended around the combustion chamber. The reformer stacks are distributed along a length of the combustion chamber in the axial direction and configured to provide output products to the combustion chamber.

First claim

We claim:

1. A propulsion system for an aircraft, the aircraft comprising an aircraft fuel supply, the propulsion system comprising:
a turbomachine defining an axial direction and comprising a compressor section, a combustor, and a turbine section arranged in serial flow order, the combustor including a liner partially defining a combustion chamber and an opening at an upstream end of the combustion chamber, the combustor configured to receive a flow of aviation fuel from the aircraft fuel supply through the opening; and
a plurality of reformer stacks located on the liner and extended around the combustion chamber, each reformer stack of the plurality of reformer stacks is a fuel processing unit being configured to generate and provide output products, including hydrogen, to the combustion chamber, the plurality of reformer stacks distributed along a length of the combustion chamber in the axial direction;
wherein the plurality of reformer stacks have respective fuel inputs that are independently controlled;
wherein the plurality of reformer stacks have respective oxidant inputs that are independently controlled;
wherein the respective fuel inputs and the respective oxidant inputs are configured to be controlled to control respective hydrogen conversion rates of the plurality of reformer stacks;
wherein the plurality of reformer stacks comprises a first reformer stack including a first housing circumferentially surrounded by a first channel, and a second reformer stack including a second housing circumferentially surrounded by a second channel; and wherein the first reformer stack defines a first height extending in a radial direction outward from the liner, wherein the second reformer stack defines a second height extending in the radial direction outward from the liner, and wherein the first height is greater than the second height; and
wherein the liner is an outer liner defining at least in part the combustion chamber, wherein the plurality of reformer stacks are extended around or integrated into the outer liner.

2. The propulsion system of claim 1 , wherein the plurality of reformer stacks are connected one to the next in a series flow arrangement.

3. The propulsion system of claim 1 , wherein the second reformer stack is positioned upstream of the first reformer stack.

4. The propulsion system of claim 3 , wherein the first height includes a greater amount of catalyst than the second height.

5. The propulsion system of claim 3 , wherein the first housing of the first reformer stack extends a first length in the axial direction, wherein the second housing of the second reformer stack extends a second length in the axial direction, and wherein the first length is different than the second length.

6. The propulsion system of claim 3 , wherein the first reformer stack includes a greater amount of catalyst than the second reformer stack such that the first reformer stack is configured to have a higher conversion rate than the second reformer stack.

7. The propulsion system of claim 3 , further comprising
a first fuel flowline in fluid communication with the first channel of the first reformer stack; and
a second fuel flowline in fluid communication with the second channel of the second reformer stack, wherein the first fuel flowline and the second fuel flowline are configured to be independently controlled.

8. The propulsion system of claim 7 , wherein the first fuel flowline comprises a first valve and wherein the second fuel flowline comprises a second valve;
the propulsion system further comprising a controller, wherein the controller is configured to control the first valve and the second valve independently.

9. An integrated reformer and combustor assembly for a turbomachine, the turbomachine defining an axial direction, the integrated reformer and combustor assembly comprising:
a combustor including a liner, at least partially defining a combustion chamber and an opening at an upstream end of the combustion chamber, the combustor configured to receive a flow of aviation fuel through the opening; and
a first reformer stack and a second reformer stack located on the liner and extended around the combustion chamber, the first reformer stack including a first channel defining a first fuel inlet and a first oxidant inlet, the second reformer stack including a second channel defining a second fuel inlet and a second oxidant inlet, wherein the first reformer stack and the second reformer stack are distributed along a length of the combustion chamber in the axial direction;
wherein the first fuel inlet and the second fuel inlet have independent fuel inputs that are independently controlled;
wherein the first oxidant inlet and the second oxidant inlet have independent oxidant inputs that are independently controlled;
wherein the respective fuel inputs and the respective oxidant inputs are configured to be controlled to control respective hydrogen conversion rates of the first reformer stack and the second reformer stack;
wherein the first reformer stack defines a first length in the axial direction between an upstream end of the first reformer stack and a downstream end of the first reformer stack, wherein the second reformer stack defines a second length in the axial direction between an upstream end of the second reformer stack and a downstream end of the second reformer stack, and wherein the first length is different than the second length;
wherein the first reformer stack defines a first height extending in a radial direction outward from the liner, wherein the second reformer stack defines a second height extending in the radial direction outward from the liner, and wherein the first height is greater than the second height; and
wherein the liner is an outer liner defining at least in part the combustion chamber, wherein the first reformer stack and the second reformer stack are extended around or integrated into the outer liner.

10. The integrated reformer and combustor assembly of claim 9 , wherein the first reformer stack and the second reformer stack are connected one to the next in a series flow arrangement.

11. The integrated reformer and combustor assembly of claim 9 , wherein the second reformer stack is positioned upstream of the first reformer stack.

12. The integrated reformer and combustor assembly of claim 11 , wherein the first reformer stack includes a greater amount of catalyst than the second reformer stack such that the first reformer stack is configured to have a higher conversion rate than the second reformer stack.

13. An integrated reformer and combustor assembly for a turbomachine, the turbomachine defining an axial direction, the integrated reformer and combustor assembly comprising:
a combustor, the combustor including a liner at least partially defining a combustion chamber and an opening at an upstream end of the combustion chamber, the combustor configured to receive a flow of aviation fuel through the opening; and
a first reformer stack and a second reformer stack located on the liner and extended around the combustion chamber, the first reformer stack including a first channel defining a first fuel inlet and a first oxidant inlet, the second reformer stack including a second channel defining a second fuel inlet and a second oxidant inlet, each of the first reformer stack and the second reformer stack is a fuel processing unit being configured to generate and provide output products, including hydrogen, to the combustion chamber, wherein the first reformer stack and the second reformer stack are distributed along a length of the liner in the axial direction;
wherein the first fuel inlet and the second fuel inlet

Assignees

• Gen Electric

Inventors

• Wang Honggang
• Benjamin Michael Anthony

Classifications

• F23R3/36 Primary

  Supply of different fuels · CPC title

• F05D2220/32

  in gas turbines · CPC title

• F05D2240/35

  Combustors or associated equipment · CPC title

• F02C7/22

  Fuel supply systems · CPC title

• F02C3/30

  Adding water, steam or other fluids {for influencing combustion, e.g. to obtain cleaner exhaust gases (F02C7/141, F02C7/30, F01D21/00, F01K21/04, F23D11/10 take precedence)} · CPC title