Recovered-cycle aircraft turbomachine

Patent metadata

Field	Value
Publication number	US-12286926-B2
Application number	US-202118001441-A
Country	US
Kind code	B2
Filing date	Jun 15, 2021
Priority date	Jun 17, 2020
Publication date	Apr 29, 2025
Grant date	Apr 29, 2025

Abstract

Aircraft turbomachine including a centrifugal compressor, a combustion chamber, the combustion chamber being supplied by the compressor via a diffuser and via a straightener, and a heat exchanger, the exchanger including a first circuit, supplied with exhaust gas from the turbomachine, and a second circuit, which are connected by volutes on the one hand to an outlet of the diffuser and on the other hand to an inlet of the straightener, the volutes having reversed winding directions such that their connection ports to the exchanger are independent of one another and are substantially diametrically opposed, and such that the minimum cross section of each duct is situated at a larger cross section of the other duct.

First claim

The invention claimed is:

1. An aircraft turbomachine, comprising:
a centrifugal compressor extending around an axis A,
an annular combustion chamber extending around the axis A,
a system for diffusing and rectifying an air flow leaving the centrifugal compressor to supply the combustion chamber, this system comprising:
an annular diffuser which is oriented substantially radially and which comprises an inlet supplied by the centrifugal compressor, and
an annular rectifier which comprises an outlet for supplying the combustion chamber, and

a heat exchanger, this exchanger comprising:
a first circuit supplied with exhaust gases from the turbomachine, and
a second circuit comprising an inlet connected by a first volute to an outlet of the diffuser, and an outlet connected by a second volute to an inlet of the rectifier, the first and second volutes being joined together and each comprising an annular conduit wound around the axis A, the first and second volutes each comprising a first port located at the external periphery of the conduit and oriented in a tangential direction, and a second port located at the internal periphery of the conduit and oriented in a radial direction, the conduit of each of the volutes having an evolving passage cross-section which is maximum at the level of the first port and minimum at a circumferential end of the conduit opposite the first port,

wherein the volutes have reversed winding directions so that their first ports are formed by conduit portions spaced apart from each other and the minimum cross-section of each conduit is located at the level of a larger cross-section of the other conduit.

2. The turbomachine according to claim 1 , wherein each of the volutes has circular or oval shaped passage cross-sections extending over an angle (β) of at least 220°.

3. The turbomachine according to claim 1 , wherein the second port of each volute comprises two annular walls extending about the axis A and defining between them an air passage duct.

4. The turbomachine according to claim 3 , wherein the two walls of the second port are substantially parallel and project from an annular skin of the volute, this skin extending about the axis A and having in axial cross-section a circular or oval shape to form said conduit.

5. The turbomachine according to claim 4 , wherein one of the walls of the second port of the first volute is coincident with one of the walls of the second port of the second volute.

6. The turbomachine according to claim 4 , wherein the walls of the second port of the first volute have free ends opposite the skin, which define a substantially radially oriented connector for connection to the outlet of the diffuser, and the walls of the second port of the second volute have free ends opposite the skin, which define a substantially axially oriented connector for connection to the inlet of the rectifier.

7. The turbomachine according to claim 4 , wherein the walls of the second ports have flat shapes and perpendicular to the axis A, or have frustoconical shapes converging from upstream to downstream towards the interior.

8. The turbomachine according to claim 4 , wherein the skin and the walls have substantially a same thickness.

9. The turbomachine according to claim 1 , wherein it further comprises an external casing which extends around the axis A and surrounds the combustion chamber, the volutes being spaced from the casing and attached to the latter by flanges.

10. The turbomachine according to claim 1 , wherein the winding direction of a volute is the direction in which the conduit of the volute winds from the first port of the volute.

11. The turbomachine according to claim 1 , wherein one of the first and second volutes has a winding direction which is clockwise, the other of the first and second volutes having a winding direction which is counterclockwise.

12. The turbomachine according to claim 1 , wherein the first port of the first volute and the first port of the second volute are distinct and spaced from each other.

13. The turbomachine according to claim 1 , wherein the first port of the first volute and the first port of the second volute are opposite with respect to the axis A.

14. A turbomachine comprising:
a combustion chamber; and
a volute assembly, said volute assembly comprising first and second volutes which are axially joined together, each including an annular conduit wound around an axis A, the first and second volutes each comprising a first port located at the external periphery of the conduit and oriented in the tangential direction, and a second port located at the internal periphery of the conduit and oriented in the radial direction, the second ports of the volutes opening onto an outside of the volute assembly, the conduit of each of the volutes having an evolving passage cross-section which is maximum at the level of the first port and minimum at a circumferential end of the conduit opposite the first port, the volutes having reversed winding directions so that their first ports are formed by conduit portions spaced apart from each other and the minimum cross-section of each conduit is located at the level of a larger cross-section of the other conduit,
wherein at least one of the second ports of the volutes extends from an annular rectifier of the combustion chamber.

15. The assembly according to claim 14 , wherein the first ports of the volutes are connected to each other by a bypass conduit.

16. The assembly according to claim 14 , wherein the winding direction of a volute is the direction in which the conduit of the volute winds from the first port of the volute.

17. The assembly according to claim 14 , wherein one of the first and second volutes has a winding direction which is clockwise, the other of the first and second volutes having a winding direction which is counterclockwise.

18. The assembly according to claim 14 , wherein the first port of the first volute and the first port of the second volute are distinct and spaced from each other.

19. The assembly according to claim 14 , wherein the first port of the first volute and the first port of the second volute are opposite with respect to the axis A.

20. The assembly according to claim 14 , wherein the second port of the first volute projects radially from the conduit of the first volute towards the axis A, the second port of the second volute projecting radially from the conduit of the second volute towards the axis A.

21. A method for manufacturing an aircraft turbomachine, said turbomachine comprising:
a centrifugal compressor extending around an axis A,
an annular combustion chamber extending around the axis A,
a system for diffusing and rectifying an air flow leaving the centrifugal compressor to supply the combustion chamber, this system comprising:
an annular diffuser which is oriented substantially radially and which comprises an inlet supplied by the centrifugal compressor, and
an annular rectifier which comprises an outlet for supplying the combustion chamber,
wherein the method comprises a step of: equipping the turbomachine with a volute assembly, said assembly comprising first and second volutes which are axially joined together, each including an annular conduit wound around an axis A, the first and second volutes each comprising a first port located at the external periphery of the conduit and oriented in the tangential direction, and a second port located at the internal periphery of the conduit and oriented in the radial direction, the seco

Assignees

• Safran Helicopter Engines

Inventors

• Pons Bernard Claude
• Puerto Alphonse
• Tarnowski Laurent Pierre
• Vive Lois Pierre Denis

Classifications

• F04D29/4233

  with volutes extending mainly in axial or radially inward direction · CPC title

• F02C3/08

  the compressor comprising at least one radial stage (F02C3/10 takes precedence) · CPC title

• F05D2220/329

  in helicopters · CPC title

• F02C7/08

  Heating air supply before combustion, e.g. by exhaust gases · CPC title

• B64C27/12

  Rotor drives · CPC title