Auxiliary power unit with excess air recovery

• US 10,696,417 B2
• Filed: 06/25/2015
• Issued: 06/30/2020
• Est. Priority Date: 06/25/2015
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Original Assignee: Pratt & Whitney Canada Corporation (Rtx Corporation)
• Inventors: Ullyott, Richard, Jones, Anthony, Julien, Andre, Thomassin, Jean
• Primary Examiner(s): Laurenzi, Mark A
• Assistant Examiner(s): Kebea, Jessica L
• Application Number:
  US14/750,187
• Publication Number:
  US 20160376022A1
• Time in Patent Office: 1,832 Days
• CPC Class Codes: B64D 41/00 Power installations for auxiliary purposes
• CPC Class Codes: F02B 2053/005 Wankel engines
• CPC Class Codes: F02B 41/10 in exhaust turbines use of exhaust turbines for charging F02B37/00
• CPC Class Codes: F02B 53/02 Methods of operating
• CPC Class Codes: F02C 3/04 having a turbine driving a compressor power transmission arrangements F02C7/36; control of working fluid flow F02C9/16
• CPC Class Codes: H02K 7/1807 Rotary generators H02K7/006 takes precedence
• CPC Class Codes: H02K 7/1823 structurally associated with turbines or similar engines
• CPC Class Codes: Y02T 50/40 Weight reduction
• CPC Class Codes: Y02T 50/50 On board measures aiming to increase energy efficiency

Abstract

An auxiliary power unit for an aircraft, having an engine core, a compressor having an outlet in fluid communication with the engine core inlet, a turbine section in fluid communication with the engine core outlet, and an excess air duct having a first end in fluid communication with the compressor outlet and a second end in fluid communication with a turbine inlet of the turbine section. The excess air duct defines a flow path between the compressor outlet and the turbine section separate from the engine core. The auxiliary power unit may include a generator in driving engagement with the engine core to provide electrical power for the aircraft. A method of providing compressed air and electrical power to an aircraft is also discussed.

Claims

1. An auxiliary power unit for an aircraft, comprising:

an internal combustion engine core having a shaft;

a compressor having an outlet in fluid communication with an inlet of the engine core;

a turbine section in fluid communication with an outlet of the engine core, 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, the first stage turbine having at least one rotor on a turbine shaft of the turbine section, the second stage turbine having at least one rotor on the turbine shaft, the turbine shaft in driving engagement with the shaft of the engine core to compound power with the engine core;

a bleed conduit having a first end in fluid communication with the outlet of the compressor and a second end configured for connection with a pneumatic system of the aircraft;

an excess air duct having a first end in fluid communication with the outlet of the compressor and a second end in fluid communication with the turbine section, the excess air duct defining a flow path between the outlet of the compressor and the turbine section separate from the engine core, wherein the second end of the excess air duct is in fluid communication with the second stage turbine of the turbine section downstream of the first stage turbine such that the flow path bypasses the first stage turbine; and

a diverter valve selectively opening and closing a fluid communication through the excess air duct between the first and second ends thereof.

2. The auxiliary power unit as defined in claim 1, wherein the bleed conduit is 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 second end of the bleed conduit configured for connection to the pneumatic system.

3. The auxiliary power unit as defined in claim 1, wherein the engine core includes at least one rotary internal combustion engine each including 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.

4. The auxiliary power unit as defined in claim 1, wherein the turbine section is in driving engagement with the compressor.

5. The auxiliary power unit as defined in claim 1, wherein the second stage turbine has a higher reaction ratio than that of the first stage turbine, the second end of the excess air duct being in fluid communication with the turbine section by being in fluid communication with the inlet of the second stage turbine.

6. The auxiliary power unit as defined in claim 5, further comprising a generator in driving engagement with the turbine shaft of the turbine section.

7. The auxiliary power unit as defined in claim 1, further comprising a heat exchanger including at least one first conduit in heat exchange relationship with at least one second conduit, the excess air duct being in fluid communication with the turbine section through the at least one first conduit, the at least one second conduit being in fluid communication with an exhaust of the turbine section.

8. The auxiliary power unit as defined in claim 1, further comprising variable inlet guide vanes, a variable diffuser or a combination thereof at an inlet of the compressor.

9. An auxiliary power unit for an aircraft, comprising:

a compressor;

an internal combustion engine core having an inlet in fluid communication with an outlet of the compressor, the engine core having a shaft;

a generator in driving engagement with the shaft of the engine core to provide electrical power for the aircraft;

a turbine section in fluid communication with an outlet of the engine core, 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, the first stage turbine having at least one rotor on a turbine shaft of the turbine section, the second stage turbine having at least one rotor on the turbine shaft, the turbine shaft in driving engagement with the shaft of the engine core to compound power with the engine core;

a bleed conduit having a first end in fluid communication with the outlet of the compressor and a second end configured for connection with a pneumatic system of the aircraft;

an excess air duct having a first end in fluid communication with the outlet of the compressor and a second end in fluid communication with an inlet of a turbine of the turbine section, the excess air duct defining a flow path between the outlet of the compressor and the turbine section separate from the engine core, wherein the second end of the excess air duct is in fluid communication with the second stage turbine of the turbine section downstream of the first stage turbine such that the flow path bypasses the first stage turbine; and

a diverter valve selectively opening and closing a fluid communication through the excess air duct between the first and second ends thereof.

10. The auxiliary power unit as defined in claim 9, further comprising bleed air valve selectively opening and closing the fluid communication between the second end of the bleed conduit and the outlet of the compressor.

11. The auxiliary power unit as defined in claim 9, wherein the engine core includes at least one rotary internal combustion engine each including 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.

12. The auxiliary power unit as defined in claim 9, 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.

13. The auxiliary power unit as defined in claim 9, further comprising a heat exchanger including at least one first conduit in heat exchange relationship with at least one second conduit, the excess air duct being in fluid communication with the turbine section through the at least one first conduit, the at least one second conduit being in fluid communication with an exhaust of the turbine section.

14. The auxiliary power unit as defined in claim 9, further comprising a heat exchanger including at least one first conduit in heat exchange relationship with at least one second conduit, the excess air duct being in fluid communication with the inlet of the second stage turbine through the at least one first conduit, the at least one second conduit being in fluid communication with an exhaust of the second stage turbine.

15. A method of providing compressed air and electrical power to an aircraft, the method comprising:

flowing compressed air from an outlet of a compressor to an inlet of an internal combustion engine core of an auxiliary power unit;

driving at least one generator providing electrical power to the aircraft with the engine core;

flowing exhaust gas from the engine core to a turbine section of the auxiliary power unit, the turbine section including a first stage turbine and a second stage turbine, wherein the first stage turbine has an exhaust inlet in fluid communication with an exhaust outlet of the engine core and the second stage turbine has an exhaust inlet in fluid communication with an exhaust outlet of the first stage turbine;

flowing excess compressed air from the outlet of the compressor to a compressed air inlet of the turbine section of the auxiliary power unit through a flow path separate from the engine core, the compressed air inlet of the turbine section being downstream of an exhaust outlet of the first stage turbine and upstream of the exhaust inlet of the second stage turbine such that the flow path bypasses the first stage turbine, the first stage turbine having at least one rotor engaged on a turbine shaft in driving engagement with a shaft of the engine core, the second stage turbine having at least one rotor engaged on the turbine shaft; and

flowing compressed air from the outlet of the compressor to a system of the aircraft.

16. The method as defined in claim 15, further comprising exchanging heat between the compressed air flowed through the flow path separate from the engine core and an exhaust of the turbine section.