Turbine engine including a continuous wave detonation chamber and cooling bypass flow and aircraft provided with such a turbine engine

Patent metadata

Field	Value
Publication number	US-9556794-B2
Application number	US-201214116001-A
Country	US
Kind code	B2
Filing date	May 9, 2012
Priority date	May 16, 2011
Publication date	Jan 31, 2017
Grant date	Jan 31, 2017

Abstract

The chamber ( 4 ) of the turbine engine ( 1 ) comprises a continuous detonation wave engine ( 6 ) provided with an annular detonation chamber ( 7 ) and associated means ( 8, 9 ) that can be used to generate a continuous production of hot gases from a detonation mixture of fuel and air. The continuous detonation wave engine ( 6 ) is arranged such as to form, from a flow of incoming air (E), a first flow (F 1 ) which enters the detonation chamber ( 7 ) and which is used by the engine ( 6 ) and a second flow (F 2 ) which bypasses the chamber. The turbine engine ( 1 ) also includes auxiliary means ( 10 ) for mixing the hot gases (F 3 ) leaving the detonation chamber ( 7 ) with the second flow of air (F 2 ) before directing same towards the turbine ( 5 ). A plurality of detonation chambers ( 7 ) are arranged concentrically to one another relative to the axis of the turbine engine.

First claim

The invention claimed is:

1. A turbomachine comprising:
a compressor ( 3 ) configured to generate a compressed air flow from an incoming air flow (E);
a chamber ( 4 ), in which a first portion of said compressed air flow flows through at least one additional compressor ( 11 ) to form a first compressed air flow (F 1 ), and a remainder of said compressed air flow forms a second compressed air flow (F 2 ) that flows around said at least one additional compressor, said chamber ( 4 ) comprising a continuous detonation wave engine ( 6 ), fitted with at least one annular detonation chamber ( 7 ) and associated means ( 8 , 9 ) to generate continuous production of hot gases from an explosive mixture of a fuel and-said first compressed air flow, wherein said fuel is injected into said at least one annular detonation chamber independent of said first compressed air flow;
auxiliary means ( 10 ) for mixing the hot gases (f 3 ) emanating from said at least one annular detonation chamber ( 7 ) with said second compressed air flow (F 2 ) to generate a mixed gas flow;
at least one additional annular detonation chamber ( 7 ) arranged in a concentric manner in relation to at least one annular detonation chamber, and in relation to a central axis ( 2 ) of the turbomachine ( 1 ); and
a turbine ( 5 ) driven in rotation by said mixed gas flow, and wherein said turbine drives said compressor ( 3 ).

2. The turbomachine according to claim 1 , wherein said auxiliary means ( 10 ) comprise an ejector/mixing system which allows the hot gases (F 3 ) to be diluted and part of the momentum thereof to be returned air of said second compressed air flow (F 2 ) in order to obtain a gas mixture which has a temperature which is compatible with the behavior of the turbine ( 5 ).

3. The turbomachine according to claim 1 wherein said at least one additional compressor ( 11 ) is arranged downstream of said compressor ( 3 ) and upstream of said at least one annular detonation chamber ( 7 ) so as to further compress said first compressed air flow (F 1 ) before it enters said at least one annular detonation chamber.

4. The turbomachine according to claim 1 wherein said turbomachine comprises at least one additional compressor which s arranged downstream of said compressor ( 3 ) so as to further compress said second compressed air flow (F 2 ).

5. The turbomachie according to claim 1 wherein said turbomachine comprises at least one circuit for cooling said at least one annular detonation chamber, wherein at least a portion of said fuel is circulated through said circuit before the injection thereof into said chamber.

6. The turbomachine according to claim 5 , wherein said at least one circuit for cooling extends along at least one lateral wall of said at least one annular detonation chambers, over at least part of the length threrof.

7. The turbomaehine according to claim 1 wherein said turbomachine is of a single flow type, comprising a single flow (E), wherein said incoming air flow comprises said single flow (E).

8. The turbomachine according to claim 1 wherein said turbomachine is of the double flow type, comprising a primary flow and a secondary flow, wherein said incoming air flow comprises said primary flow.

9. A flying vehicle. wherein said vehicle is fitted with at least one turbomachine ( 1 ) as specified in claim 1 .

10. A power generation system, wherein said system is fitted with at least one turbomachine ( 1 ) as specified in claim 1 .

11. The ttirbomachine according to claim 1 , wherein said at least one annular detonation chamber comprises a plurality of annular detonation chambers arranged in a concentric manner in relation to each other, relative to an axis ( 2 ) of the turbomachine ( 1 ).

Assignees

• Falempin François
• Le Naour Bruno
• Mbda France

Inventors

• Falempin François
• Le Naour Bruno

Classifications

• F02K7/08

  the jet being continuous · CPC title

• F02C7/22 Primary

  Fuel supply systems · CPC title

• F02C3/14 Primary

  characterised by the arrangement of the combustion chamber in the plant (combustion chambers per se F23R; F02C3/205 takes precedence) · CPC title

• F23R7/00

  Intermittent or explosive combustion chambers · CPC title

• F23R3/50

  Combustion chambers comprising an annular flame tube within an annular casing (toroidal combustion chambers F23R3/52) · CPC title