Reverse flow gas turbine engine having electric machine
View Patent ↗An aircraft engine assembly includes a gas turbine engine having an intake channel configured to receive an incoming flow of air and thereby form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction, and an electric machine coupled with the low pressure shaft and located at the aft end of the gas turbine engine proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated.
1. An aircraft engine assembly comprising:
a gas turbine engine having a high pressure compressor, a high pressure turbine, a high pressure shaft coupling the high pressure compressor with the high pressure turbine, a low pressure turbine, and a low pressure shaft coupled to the low pressure turbine, the high pressure turbine located forward of the high pressure compressor, and the low pressure turbine located on a forward end of the gas turbine engine;
a propeller located on a forward end of the gas turbine engine and coupled via the low pressure shaft with the low pressure turbine;
an intake channel of the gas turbine engine configured to receive an incoming flow of air and form an intake flow of air, the intake channel configured to turn the received incoming flow of air from an incoming flow direction to a first axial direction of the gas turbine engine, the incoming flow direction reverse of the first axial direction; and
an electric machine coupled with the low pressure shaft and located on a side of the high pressure compressor opposite of the high pressure turbine and proximate the intake channel, the electric machine in heat exchange communication with the intake flow of air such that the electric machine transfers heat to the incoming flow of air within the intake channel when the electric machine is operated,
wherein the high pressure compressor includes:
a shaft comprising a forward end portion, wherein the forward end portion defines an outer surface and rotates with the shaft;
a first row of compressor rotor blades coupled to the shaft downstream from the forward end portion;
an outer casing at least partially surrounding the first row of compressor rotor blades and the outer surface of the forward end portion of the shaft, the outer casing at least partially defining an inlet to the high pressure compressor; and
an inlet guide vane comprising a mounting portion, a tip portion, a leading-edge portion, and a trailing-edge portion, wherein the mounting portion is coupled to the outer casing upstream from the first row of compressor rotor blades, wherein the tip portion extends towards the outer surface of the forward end portion of the shaft, and wherein a radial gap is defined between the tip portion and the outer surface.
2. The aircraft engine assembly of claim 1 , wherein the forward end portion of the shaft and an inner wall of an inlet frame upstream from the inlet guide vane define an aperture therebetween, wherein the aperture is in fluid communication with a purge-air source, wherein the aperture is oriented to direct a flow of purge-air towards the radial gap.
3. The aircraft engine assembly of claim 1 , wherein the forward end portion of the shaft terminates axially forward of the leading-edge portion of the inlet guide vane.
4. The aircraft engine assembly of claim 1 , further comprising a seal body disposed within the radial gap.
5. The aircraft engine assembly of claim 1 , wherein the forward end portion of the shaft and an inner wall of an inlet frame define an aperture therebetween, wherein the aperture is located between the leading-edge portion and the trailing-edge portion of the inlet guide vane, wherein the aperture is in fluid communication with a purge-air source, and wherein the aperture is oriented to direct a flow of purge air into the radial gap.
6. The aircraft engine assembly of claim 1 , wherein the forward end portion of the rotor shaft terminates axially aft from the leading-edge portion of the inlet guide vane.
7. The aircraft engine assembly of claim 1 , wherein the forward end portion of the shaft includes a plurality of protrusions circumferentially spaced about and extending radially outwardly from the outer surface, wherein the plurality of protrusions is disposed upstream of the radial gap.
8. The aircraft engine assembly of claim 1 , wherein the inlet guide vane includes an internal air passage having an inlet in fluid communication with a purge air source and an outlet defined along the leading-edge portion of the inlet guide vane upstream from the radial gap.
9. The aircraft engine assembly of claim 1 , wherein the inlet guide vane includes an internal air passage having an inlet in fluid communication with a purge air source and an outlet defined within the radial gap.
10. The aircraft engine assembly of claim 1 , wherein the outer casing includes an inlet frame, wherein the mounting portion of the inlet guide vane is coupled to the inlet frame.
11. A turboprop aircraft powerplant comprising:
a gas turbine engine having a high pressure compressor and a high pressure turbine, the gas turbine engine further having a high pressure shaft coupling the high pressure compressor with the high pressure turbine, the gas turbine engine also having a first axial flow direction from the high pressure compressor to the high pressure turbine;
a propeller coupled to a low pressure turbine of the gas turbine engine using a low pressure shaft, the low pressure shaft located coaxial with the high pressure shaft, the propeller configured to receive a free stream flow of air oriented in a freestream direction and impart work upon the free stream flow of air, the propeller located on an upstream side of the freestream direction from the high pressure turbine;
an intake channel defining an intake flow of air in fluid communication with the gas turbine engine, the intake channel configured to reverse the intake flow of air initially flowing in the freestream direction to the first axial flow direction of the gas turbine engine; and
an electric machine coupled to the low pressure shaft and located on an opposite side of the gas turbine engine from the propeller, the electric machine positioned to be cooled by a portion of the intake flow of air defined by the intake channel,
wherein the high pressure compressor includes:
a shaft comprising a forward end portion, wherein the forward end portion defines an outer surface and rotates with the shaft;
a first row of compressor rotor blades coupled to the shaft downstream from the forward end portion;
an outer casing at least partially surrounding the first row of compressor rotor blades and the outer surface of the forward end portion of the shaft, the outer casing at least partially defining an inlet to the high pressure compressor; and
an inlet guide vane comprising a mounting portion, a tip portion, a leading-edge portion, and a trailing-edge portion, wherein the mounting portion is coupled to the outer casing upstream from the first row of compressor rotor blades, wherein the tip portion extends towards the outer surface of the forward end portion of the shaft, and wherein a radial gap is defined between the tip portion and the outer surface.
12. The turboprop aircraft powerplant of claim 11 , wherein the forward end portion of the shaft and an inner wall of an inlet frame upstream from the inlet guide vane define an aperture therebetween, wherein the aperture is in fluid communication with a purge-air source, wherein the aperture is oriented to direct a flow of purge-air towards the radial gap.
13. The turboprop aircraft powerplant of claim 11 , wherein the forward end portion of the shaft terminates axially forward of the leading-edge portion of the inlet guide vane.
14. The turboprop aircraft powerplant of claim 11 , further comprising a seal body disposed within the radial gap.
15. The turboprop aircraft powerplant of claim 11 , wherein the forward end portion of the shaft and an inner wall of an inlet frame define an aperture therebetween, wherein the aperture is located between the leading-edge portion and the trailing-edge portion of the inlet guide vane, wherein the aperture is in fluid communication with a purge-air source, and wherein the aperture is oriented to direct a flow of purge air into the radial gap.
16. The turboprop aircraft powerplant of claim 11 , wherein the forward end portion of the rotor shaft terminates axially aft from the leading-edge portion of the inlet guide vane.
17. The turboprop aircraft powerplant of claim 11 , wherein the forward end portion of the shaft includes a plurality of protrusions circumferentially spaced about and extending radially outwardly from the outer surface, wherein the plurality of protrusions is disposed upstream of the radial gap.
18. The turboprop aircraft powerplant of claim 11 , wherein the inlet guide vane includes an internal air passage having an inlet in fluid communication with a purge air source and an outlet defined along the leading-edge portion of the inlet guide vane upstream from the radial gap.
19. The turboprop aircraft powerplant of claim 11 , wherein the inlet guide vane includes an internal air passage having an inlet in fluid communication with a purge air source and an outlet defined within the radial gap.
20. The turboprop aircraft powerplant of claim 11 , wherein the outer casing includes an inlet frame, wherein the mounting portion of the inlet guide vane is coupled to the inlet frame.