Compound engine assembly with bleed air

Patent metadata

Field	Value
Publication number	US-11313273-B2
Application number	US-202016794806-A
Country	US
Kind code	B2
Filing date	Feb 19, 2020
Priority date	Jun 25, 2015
Publication date	Apr 26, 2022
Grant date	Apr 26, 2022

Abstract

A compound engine assembly for use as an auxiliary power unit for an aircraft and including an engine core with internal combustion engine(s), a compressor having an outlet in fluid communication with an engine core inlet, a bleed conduit in fluid communication with the compressor outlet through a bleed air valve, and a turbine section having an inlet in fluid communication with the engine core outlet and configured to compound power with the engine core. The turbine section may include a first stage turbine having an inlet in fluid communication with the engine core outlet and a second stage turbine having an inlet in fluid communication the first stage turbine outlet. A method of providing compressed air and electrical power to an aircraft is also discussed.

First claim

The invention claimed is:

1. An engine assembly for use as an auxiliary power unit for an aircraft, comprising:
an engine core including at least one internal combustion engine, the engine core in driving engagement with an engine shaft;
a generator drivingly engaged by the engine core via the engine shaft;
a compressor having an outlet in fluid communication with an inlet of the engine core;
a turbine section having an inlet in fluid communication with an outlet of the engine core, the turbine section in driving engagement with the compressor via a turbine shaft;
a bleed conduit having an end configured for connection with a pneumatic system of the aircraft, the bleed conduit in fluid communication with the outlet of the compressor through a bleed air valve selectively opening and closing the fluid communication between the outlet of the compressor and the end of the bleed conduit configured for connection to the pneumatic system;
an excess air duct in fluid communication with the outlet of the compressor, the excess air duct fluidly connected to the turbine section; and
a heat exchanger having at least one first conduit in heat exchange relationship with at least one second conduit, the at least one first conduit fluidly connected to the excess air duct, the at least one second conduit fluidly connected to an outlet of the turbine section such that excess compressed air from the compressor is heated by combustion gases outputted by the turbine section.

2. The engine assembly of claim 1 , wherein the turbine shaft rotates independently from the engine shaft.

3. The engine assembly as defined in claim 1 , wherein the excess air duct is fluidly connected to the inlet of the turbine section, a valve connected to the excess air duct and having an open configuration in which the outlet of the compressor is fluidly connected to the inlet of the turbine section through the valve and a closed configuration in which the valve blocks fluid communication between the outlet of the compressor and the inlet of the turbine section via the excess air duct.

4. The engine assembly as defined in claim 1 , wherein each of the at least one internal combustion engine includes a rotor sealingly and rotationally received within a respective internal cavity to provide rotating chambers of variable volume in the respective internal cavity, the rotor having three apex portions separating the rotating chambers and mounted for eccentric revolutions within the respective internal cavity, the respective internal cavity having an epitrochoid shape with two lobes.

5. The engine assembly as defined in claim 1 , wherein the turbine section includes a first stage turbine having an inlet in fluid communication with the outlet of the engine core, and a second stage turbine having an inlet in fluid communication with an outlet of the first stage turbine.

6. The engine assembly as defined in claim 5 , wherein the first stage turbine is configured as an impulse turbine with a pressure-based reaction ratio having a value of at most 0.25, the second stage turbine having a higher reaction ratio than that of the first stage turbine.

7. The engine assembly as defined in claim 1 , further comprising an inlet conduit in fluid communication with the inlet of the engine core and an outlet conduit in fluid communication with the outlet of the compressor, the inlet conduit and bleed conduit being both in fluid communication with the outlet conduit.

8. The engine assembly as defined in claim 1 , wherein the bleed conduit is in fluid communication with the outlet of the compressor at least in part through an intercooler configured to reduce a temperature of compressed air circulating from the compressor to the bleed conduit.

9. The engine assembly as defined in claim 1 , further comprising variable inlet guide vanes, a variable diffuser or a combination thereof at an inlet of the compressor.

10. An engine assembly for use as an auxiliary power unit for an aircraft, the engine assembly comprising:
an engine core including at least one internal combustion engine in driving engagement with an engine shaft;
a generator in driving engagement with the engine shaft to provide electrical power for the aircraft;
a compressor having an outlet in fluid communication with an inlet of the engine core;
a bleed conduit having an end configured for connection to a system of the aircraft, the bleed conduit being in fluid communication with the outlet of the compressor;
a bleed air valve selectively opening and closing the fluid communication between the end of the bleed conduit and the outlet of the compressor; and
a turbine section having a first stage turbine having an inlet in fluid communication with an outlet of the engine core, and a second stage turbine having an inlet in fluid communication with an outlet of the first stage turbine, the first stage turbine being configured as an impulse turbine, the second stage turbine having a reaction ratio higher than that of the first stage turbine,
one or both of the first stage turbine and the second stage turbine drivingly engaged to the compressor via a turbine shaft.

11. The engine assembly of claim 10 , wherein the turbine shaft rotates independently from the engine shaft.

12. The engine assembly as defined in claim 10 , comprising an excess air duct in fluid communication with the outlet of the compressor, the excess air duct fluidly connected to the inlet of the turbine section, a valve connected to the excess air duct and having an open configuration in which the outlet of the compressor is fluidly connected to the inlet of the turbine section through the valve and a closed configuration in which the valve blocks fluid communication between the outlet of the compressor and the inlet of the turbine section via the excess air duct.

13. The engine assembly as defined in claim 12 , comprising a heat exchanger having at least one first conduit in heat exchange relationship with at least one second conduit, the at least one first conduit fluidly connected to the excess air duct, the at least one second conduit fluidly connected to an outlet of the turbine section such that excess compressed air from the compressor is heated by combustion gases outputted by the turbine section.

14. The engine assembly as defined in claim 10 , wherein each of the at least one internal combustion engine includes a rotor sealingly and rotationally received within a respective internal cavity to provide rotating chambers of variable volume in the respective internal cavity, the rotor having three apex portions separating the rotating chambers and mounted for eccentric revolutions within the respective internal cavity, the respective internal cavity having an epitrochoid shape with two lobes.

15. The engine assembly as defined in claim 10 , wherein the turbine shaft is drivingly engaged to rotors of the compressor and of both of the first and second stage turbines.

16. The engine assembly as defined in claim 10 , wherein the turbine shaft is radially offset from the engine shaft relative to a rotation axis of the turbine shaft.

17. The engine assembly as defined in claim 10 , wherein the first stage turbine has a pressure-based reaction ratio having a value of at most 0.25.

18. The engine assembly as defined in claim 10 , further comprising an inlet conduit in fluid communication with the inlet of the engine core and an outlet conduit communicating with the outlet of the compressor, the inlet conduit and bleed conduit being both in fluid communication with the outlet conduit.

19

Assignees

• Pratt & Whitney Canada

Inventors

• Ullyott Richard
• Jones Anthony
• Julien Andre
• Thomassin Jean

Classifications

• Y02T10/12

  Improving ICE efficiencies · CPC title

• B64D41/00 Primary

  Power installations for auxiliary purposes · CPC title

• Y02T50/40

  Weight reduction · 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

• F05B2260/40

  Transmission of power · CPC title