Aircraft power plant with detonation combustion tube

Patent metadata

Field	Value
Publication number	US-12203429-B2
Application number	US-202218066302-A
Country	US
Kind code	B2
Filing date	Dec 15, 2022
Priority date	Dec 15, 2022
Publication date	Jan 21, 2025
Grant date	Jan 21, 2025

Abstract

An aircraft power plant, has: a combustion engine having an outlet outputting combustion gases; a turbine downstream of the combustion engine; a detonation combustion tube fluidly connecting the combustion engine to the turbine; a member having an open position in which the outlet of the combustion engine is fluidly connected to the turbine and a closed position in which the combustion engine is fluidly disconnected from the turbine; a fuel injector fluidly connected to the detonation combustion tube; an igniter operatively connected to the detonation combustion tube; and a controller operatively connected to the fuel injector and to the igniter, the controller configured to, in response to receiving of a command: inject fuel into the detonation combustion tube via the fuel injector, and once the member is in the closed position, power the igniter to ignite a mixture of the combustion gases and the fuel into the detonation combustion tube.

First claim

The invention claimed is:

1. An aircraft power plant, comprising:
a combustion engine having an outlet outputting combustion gases;
a turbine located downstream of the combustion engine relative to the combustion gases;
a detonation combustion tube fluidly connecting the combustion engine to the turbine, the detonation combustion tube having a tube inlet and a tube outlet;
a member disposed within the detonation combustion tube downstream of the tube inlet and upstream of the tube outlet, the member having an open position in which the tube inlet is fluidly connected to the tube outlet through the member and a closed position in which the tube inlet is fluidly disconnected from the tube outlet by the member;
a fuel injector fluidly connected to the detonation combustion tube downstream of the outlet of the combustion engine;
an igniter located downstream of the member and operatively connected to the detonation combustion tube; and
a controller operatively connected to the fuel injector and to the igniter, the controller configured to, in response to receiving of a command:
a) inject a fuel into the detonation combustion tube via the fuel injector,
b) once the member is in the closed position, power the igniter to ignite a mixture of the combustion gases and the fuel injected into the detonation combustion tube, and
c) repeat steps a) and b) for each of a plurality of successive cycles of combustion sequentially occurring one after the other.

2. The aircraft power plant of claim 1 , wherein the controller is configured to:
inject the fuel into the detonation combustion tube at fuel injection events in which fuel is injected into the detonation combustion tube via the fuel injector, the fuel injection events occurring at an injection frequency corresponding to a pulse frequency of pulses of the combustion gases generated by the combustion engine and entering the detonation combustion tube, the fuel injection events occurring between the pulses of the combustion gases.

3. The aircraft power plant of claim 2 , wherein the controller is operatively connected to the member and configured to:
move the member between the open position and the closed position at a frequency corresponding to the injection frequency.

4. The aircraft power plant of claim 3 , wherein the member is a thrust wall having a first disc defining at least one first aperture and a second disc defining at least one second aperture, the first disc rotatable relative to the second disc, the at least one first aperture being in register with the at least one second aperture in the open position, the at least one first aperture being circumferentially offset from the at least one second aperture in the closed position.

5. The aircraft power plant of claim 4 , wherein the first disc is drivingly engaged to the combustion engine.

6. The aircraft power plant of claim 5 , wherein the first disc is drivingly engaged to the combustion engine via a gearbox.

7. The aircraft power plant of claim 6 , wherein the controller is operatively connected to the member via the gearbox, the controller being configured to:
determine a required speed ratio between the first disc and a shaft of the combustion engine such that a frequency at which the member moves from the open position to the closed position corresponds to the pulse frequency of the pulses of the combustion gases generated by the combustion engine; and
operate the gearbox at the required speed ratio.

8. The aircraft power plant of claim 1 , wherein the combustion engine includes a plurality of rotary engine units, a manifold having a plurality of manifold inlets fluidly connected to a manifold outlet, each of the plurality of rotary engine units fluidly connected to a respective one of the plurality of manifold inlets, the manifold outlet fluidly connected to the tube inlet of the detonation combustion tube.

9. The aircraft power plant of claim 1 , wherein the fuel injector is an electrically-controllable injector.

10. The aircraft power plant of claim 1 , wherein the fuel injector is located upstream of the member.

11. The aircraft power plant of claim 1 , wherein the controller is configured to operate the combustion engine at an air-fuel equivalence ratio being greater than one.

12. The aircraft power plant of claim 11 , wherein the detonation combustion tube is devoid of a dedicated air inlet.

13. A method of operating an aircraft power plant having a detonation combustion tube fluidly connecting a combustion engine to a turbine, the method comprising:
generating combustion gases with the combustion engine;
receiving the combustion gases generated by the combustion engine into the detonation combustion tube via a tube inlet of the detonation combustion tube;
injecting fuel into the detonation combustion tube; and
a) fluidly disconnecting the tube inlet from a tube outlet of the detonation combustion tube;
b) igniting a mixture of the combustion gases and the fuel received into the detonation combustion tube to detonate the mixture thereby generating a pulse of exhaust gases;
c) extracting energy from the pulse of the exhaust gases with the turbine; and
d) repeating steps a) to c) for each of a plurality of successive cycles of combustion sequentially occurring one after the other.

14. The method of claim 13 , wherein the receiving of the combustion gases includes receiving pulses of the combustion gases at a pulse frequency, the injecting of the fuel into the detonation combustion tube includes injecting the fuel into the detonation combustion tube at fuel injection events having an injection frequency corresponding to the pulse frequency, the fuel injection events occurring between the pulses of the combustion gases.

15. The method of claim 14 , wherein the fluidly disconnecting of the detonation combustion tube from the combustion engine includes moving a member from an open position in which in which an outlet of the combustion engine is fluidly connected to the turbine to a closed position in which the combustion engine is fluidly disconnected from the turbine by the member.

16. The method of claim 15 , wherein the member is a thrust wall having a first disc defining at least one first aperture and a second disc defining at least one second aperture, the moving of the member from the open position to the closed position includes rotating the first disc relative to the second disc until the at least one first aperture is offset from the at least one second aperture.

17. The method of claim 16 , wherein the first disc is drivingly engaged to the combustion engine via a gearbox, the method comprising determining a speed ratio between the first disc and a shaft of the combustion engine such that a frequency at which the member moves from the open position to the closed position corresponds to the pulse frequency of the pulses of the combustion gases generated by the combustion engine, and operating the gearbox at the speed ratio.

18. The method of claim 13 , wherein the receiving the combustion gases generated by the combustion engine into the detonation combustion tube includes converging the combustion gases outputted by a plurality of rotary engine units into the detonation combustion tube.

19. The method of claim 13 , wherein the generating of the combustion gases includes generating the combustion gases by operating the combustion engine at an air-fuel equivalence ratio being greater than one.

20. The method of claim 19 , comprising injecting air into the detonation combustion tube s

Assignees

• Pratt & Whitney Canada

Inventors

• Plamondon Etienne
• Bousquet Michel
• Nguyen Kevin

Classifications

• F05D2220/80

  in supersonic vehicles excluding hypersonic vehicles or ram, scram or rocket propulsion · CPC title

• F02C6/12

  Turbochargers, i.e. plants for augmenting mechanical power output of internal-combustion piston engines by increase of charge pressure · CPC title

• F02C3/14

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

• F01C11/008

  and of complementary function, e.g. internal combustion engine with supercharger · CPC title

• F01C1/22

  of internal-axis type with equidirectional movement of co-operating members at the points of engagement, or with one of the co-operating members being stationary, the inner member having more teeth or tooth- equivalents than the outer member · CPC title