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

1. A propulsion system for an aircraft, the aircraft comprising an aircraft fuel supply, the propulsion system comprising: a turbomachine comprising a compressor section, a combustor, and a turbine section arranged in serial flow order, the combustor defining a combustion chamber and an opening at an upstream end of the combustion chamber, the turbomachine defining an axial direction and a radial direction, the combustor configured to receive a flow of aviation fuel from the aircraft fuel supply through the opening; and a reformer stack extended around the combustion chamber and configured to provide output products to the combustion chamber, the reformer stack comprising a plurality of reformers aligned in the radial direction, wherein the reformer stack is positioned at a downstream section of the combustion chamber along the axial direction, wherein the propulsion system defines a downstream distance in the axial direction between the opening of the combustor and an upstream end of the reformer stack, and wherein the downstream distance is at least 30% of a length of the combustion chamber in the axial direction.

2. The propulsion system of claim 1 , wherein the downstream distance is at least half of the length of the combustion chamber in the axial direction.

3. The propulsion system of claim 1 , wherein the downstream distance is greater than two-thirds of the length of the combustion chamber in the axial direction.

4. The propulsion system of claim 1 , wherein the downstream distance in the axial direction is a distance between the opening and a next downstream flow of output products into the combustion chamber.

5. The propulsion system of claim 1 , wherein the reformer stack is a forward-most reformer stack.

6. The propulsion system of claim 1 , wherein the combustor includes an outer liner and an inner liner defining at least in part the combustion chamber, wherein the reformer stack is extended around or integrated into at least one of the outer liner and the inner liner.

7. The propulsion system of claim 1 , the reformer stack comprising a channel extending around one of an outside and an inside of the reformer stack in the radial direction.

8. The propulsion system of claim 1 , wherein the reformer stack is one of a CPO x reformer and an autothermal reformer, wherein the reformer stack is configured to receive air if the reformer stack is a CPO x reformer and the reformer stack is configured to receive steam if the reformer stack is an autothermal reformer.

9. An integrated reformer and combustor assembly for a turbomachine, the turbomachine defining an axial direction and a radial direction, the integrated reformer and combustor assembly comprising: a combustor a 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 when incorporated into the turbomachine; and a reformer stack extended around the combustion chamber and configured to provide output products to the combustion chamber, the reformer stack comprising a plurality of reformers aligned in the radial direction, wherein the reformer stack is positioned at a downstream section of the combustion chamber along the axial direction, wherein the propulsion system defines a downstream distance in the axial direction between the opening of the combustor and an upstream end of the reformer stack, and wherein the downstream distance is at least 30% of a length of the combustion chamber in the axial direction.

10. The propulsion system of claim 9 , wherein the downstream distance is at least half of the length of the combustion chamber in the axial direction.

11. The propulsion system of claim 9 , wherein the downstream distance is greater than two-thirds of the length of the combustion chamber in the axial direction.

12. The propulsion system of claim 9 , wherein the downstream distance in the axial direction is a distance between the opening and a next downstream flow of output products into the combustion chamber.

13. The propulsion system of claim 9 , wherein the reformer stack is a forward-most reformer stack.

14. The integrated reformer and combustor assembly of claim 9 , wherein the combustor includes an outer liner and an inner liner defining at least in part the combustion chamber, wherein the reformer stack is extended around or integrated into at least one of the outer liner and the inner liner.

15. The integrated reformer and combustor assembly of claim 9 , the reformer stack comprising a channel extending around one of an outside and an inside of the reformer stack in the radial direction.

16. A method of operating a propulsion system comprising a turbomachine and a reformer stack, the turbomachine defining an axial direction, the method comprising: providing a flow of aviation fuel to a combustion chamber of a combustor of the turbomachine through an opening defined at an upstream end of the combustion chamber to initiate an initial combustion within the combustion chamber; and providing a flow of output products from the reformer stack to the combustion chamber at a downstream section of the combustion chamber to initiate a secondary combustion within the combustion chamber at a location downstream of the initial combustion within the combustion chamber, wherein the reformer stack is positioned at a downstream section of the combustion chamber along the axial direction, wherein the propulsion system defines a downstream distance in the axial direction between the opening of the combustor and an upstream end of the reformer stack, and wherein the downstream distance is at least 30% of a length of the combustion chamber in the axial direction.