Constant volume turbomachine combustion chamber

Patent metadata

Field	Value
Publication number	US-12173646-B2
Application number	US-202017626637-A
Country	US
Kind code	B2
Filing date	Jul 8, 2020
Priority date	Jul 15, 2019
Publication date	Dec 24, 2024
Grant date	Dec 24, 2024

Abstract

A combustion chamber of an aircraft turbomachine having a main axis includes: a body of revolution coaxial with the main axis having a plurality of combustion tubes extending mainly in the direction of the main axis and being distributed in a ring about the main axis; a first perforated rotary disc mounted at a first axial end of the body and rotatable about the main axis to selectively open or close a first end of each of the combustion tubes; and a second perforated rotary disc mounted at a second axial end of the body and rotatable about the main axis to selectively open or close a second end of each of the combustion tubes. The body includes a plurality of cooling segments extending mainly in the direction of the main axis, which are distributed in a ring about the main axis and around the combustion tubes.

First claim

The invention claimed is:

1. A combustion chamber of an aircraft turbomachine having a main axis and including:
a body of revolution coaxial with the main axis in which a plurality of combustion tubes are formed, said combustion tubes extending mainly in the direction of the main axis of the body and being distributed in a ring about said main axis,
a first perforated rotary disc which is mounted at a first axial end of the body and which is movable in rotation about the main axis to selectively open or close a first end of each of said combustion tubes, and
a second perforated rotary disc which is mounted at a second axial end of the body and which is movable in rotation about the main axis to selectively open or close a second end of each of said combustion tubes,
wherein the body includes a plurality of cooling segments which extend mainly in the direction of the main axis, which are distributed in a ring about said main axis and around the combustion tubes and are configured to cool a radially outer portion of the body, and
wherein each cooling segment is open at each axial end of the body and is passed through by a compressed air flow,
wherein each rotary disc is coaxial with the main axis, includes a series of lumens intended to face one end of each combustion tube and in a selective manner, and includes a series of orifices intended to face one end of each cooling segment in a selective manner,
wherein each rotary disc is configured to selectively open or close a corresponding end of the cooling segments when the orifices are facing an associated axial end of the cooling segments.

2. The combustion chamber according to claim 1 , wherein each rotary disc includes the same number of lumens as orifices, and wherein the lumens are circumferentially offset relative to the orifices.

3. The combustion chamber according to claim 1 , wherein the body is cylindrical and includes a radially central cylindrical housing and each rotary disc includes a central orifice which communicates with the associated end of the radially central cylindrical housing of the body.

4. An aircraft turbomachine including a primary air flow path wherein at least one compressor, a combustion chamber and at least one turbine are disposed, in the direction of air flow in the primary flow path,
wherein the combustion chamber is a combustion chamber according to claim 1 .

5. The turbomachine according to claim 4 , which includes a shaft connecting one of said at least one compressor to one of said at least one turbine, said shaft being coaxial with the main axis of the body of the combustion chamber, wherein one and/or the other of the rotary discs is driven in rotation by said shaft.

6. The turbomachine according to claim 5 , which includes:
a speed reducer interposed between each disc and the shaft to drive the discs.

7. The turbomachine according to claim 6 , wherein the speed reducer is disengageable to hold the discs in a defined position.

8. The combustion chamber according to claim 1 , wherein the lumens and/or the orifices of each rotary disc are elliptical or ovoid in shape.

Assignees

• Safran Aircraft Engines

Inventors

• Escure Didier René André
• Tantot Nicolas Jérôme Jean

Classifications

• F23R7/00 Primary

  Intermittent or explosive combustion chambers · CPC title

• F02C5/12 Primary

  the combustion chambers having inlet or outlet valves, e.g. Holzwarth gas-turbine plants · CPC title