ENGINE ASSEMBLY WITH DIRECT DRIVE OF GENERATOR

• US 20170267370A1
• Filed: 06/07/2017
• Published: 09/21/2017
• Est. Priority Date: 06/25/2015
• Status: Active Grant

• Current Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Original Assignee: Pratt & Whitney Canada Corporation (Raytheon Technologies Corporation d/b/a RTX)
• Inventors: ULLYOTT, Richard, JONES, Anthony, JULIEN, Andre, THOMASSIN, Jean
• Granted Patent: US 9,994,332 B2
• Time in Patent Office: Days
• CPC Class Codes: B64D 2013/0611 combined with auxiliary power units (APU's)
• CPC Class Codes: B64D 41/00 Power installations for auxiliary purposes
• CPC Class Codes: F01C 1/22 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 Class Codes: F01C 11/008 and of complementary function, e.g. internal combustion engine with supercharger
• CPC Class Codes: F02C 3/10 with another turbine driving an output shaft but not driving the compressor
• CPC Class Codes: H02K 7/1815 structurally associated with reciprocating piston engines general aspects of generating sets, e.g. housing, F02B63/04
• 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 engine assembly for use as an aircraft auxiliary power unit, having internal combustion engine(s) in driving engagement with an engine shaft, a generator having a generator shaft directly engaged to the engine shaft such as to be rotatable at a same speed, a compressor having an outlet in communication with the internal combustion engine inlet, and a turbine having an inlet in communication with the internal combustion engine outlet. The turbine may be a first stage turbine, and the assembly may include a second stage turbine having an inlet in communication with the first stage turbine outlet. A method of providing electrical power to an aircraft is also discussed.

Claims

1. An engine assembly for use as an auxiliary power unit for an aircraft, the engine assembly comprising:

at least one internal combustion engine in driving engagement with an engine shaft;

a generator having a generator shaft in driving engagement with the engine shaft to provide electrical power for the aircraft, the generator and engine shafts being directly engaged to one another such as to be rotatable at a same speed;

a compressor having an outlet in fluid communication with an inlet of the at least one internal combustion engine; and

a turbine having an inlet in fluid communication with an outlet of the at least one internal combustion engine.

2. The engine assembly as defined in claim 1, wherein the turbine is mechanically coupled to the compressor.

3. The engine assembly as defined in claim 1, further comprising 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.

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 4, wherein the generator has a nominal frequency of 400 Hz.

6. The engine assembly as defined in claim 5, wherein the generator has a nominal rotational speed of at least 5700 rotations per minute.

7. The engine assembly as defined in claim 5, wherein the generator is selected from the group consisting of:

a 6 pole, 3 phases, alternative current generator having a design speed of from 7600 to 8400 rotations per minute;

a 8 pole, 3 phases, alternative current generator having a design speed of from 5700 to 6300 rotations per minute; and

a 4 pole, 3 phases, alternative current generator having a design speed of from 11400 to 12600 rotations per minute.

8. The engine assembly as defined in claim 1, wherein the generator is a first generator, the turbine includes at least one rotor engaged on a turbine shaft rotatable independently of the engine shaft, the assembly further comprising a second generator having a second generator shaft directly engaged to the turbine shaft such that the turbine and second generator shafts are rotatable at a same speed.

9. The engine assembly as defined in claim 1, wherein the turbine is a first stage turbine, the assembly further comprising a second stage turbine having an inlet in fluid communication with an outlet of the first stage turbine.

10. The engine assembly 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.

11. 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.

12. An engine assembly for use as an auxiliary power unit for an aircraft, the engine assembly comprising:

at least one internal combustion engine in driving engagement with an engine shaft;

a generator having a generator shaft in driving engagement with the engine shaft to provide electrical power for the aircraft, the generator and engine shafts being directly engaged to one another such as to be rotatable at a same speed;

a compressor having an outlet in fluid communication with an inlet of the at least one internal combustion engine;

a first stage turbine having an inlet in fluid communication with an outlet of the at least one internal combustion engine; and

a second stage turbine having an inlet in fluid communication with an outlet of the first stage turbine;

wherein at least one of the turbines is in driving engagement with the compressor.

13. The engine assembly as defined in claim 12, further comprising 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, and a bleed air valve selectively opening and closing the fluid communication between the end of the bleed conduit and the outlet of the compressor.

14. The engine assembly as defined in claim 12, 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 12, wherein the generator has a nominal frequency of 400 Hz.

16. The engine assembly as defined in claim 15, wherein the generator has a nominal rotational speed of at least 5700 rotations per minute.

17. The engine assembly as defined in claim 15, wherein the generator is selected from the group consisting of:

a 6 pole, 3 phases, alternative current generator having a design speed of from 7600 to 8400 rotations per minute;

a 8 pole, 3 phases, alternative current generator having a design speed of from 5700 to 6300 rotations per minute; and

a 4 pole, 3 phases, alternative current generator having a design speed of from 11400 to 12600 rotations per minute.

18. The engine assembly as defined in claim 12, wherein rotors of the first and second stage turbines and of the compressor are engaged on a turbine shaft rotatable independently of the engine shaft, the generator is a first generator, the assembly further comprising a second generator having a second generator shaft directly engaged to the turbine shaft such that the turbine and second generator shafts are rotatable at a same speed.

19. The engine assembly as defined in claim 12, 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 reaction ratio higher than that of the first stage turbine.

20. A method of providing electrical power to an aircraft, the method comprising:

flowing compressed air from an outlet of a compressor to an inlet of at least one internal combustion engine of an engine assembly;

rotating an engine shaft with the at least one internal combustion engine;

driving a generator providing electrical power to the aircraft with the engine shaft through a direct drive engagement between a shaft of the generator and the engine shaft causing the shaft of the generator and the engine shaft to rotate at a same speed; and

driving a turbine of the engine assembly with exhaust from the at least one internal combustion engine.