Low pressure recovery inlet for gas turbine engine turbomachinery on aircraft
A gas turbine engine system adapted for use with an aircraft includes a gas turbine engine. The gas turbine engine system includes a gas turbine engine and an auxiliary compressor operably coupled with the gas turbine engine. The auxiliary compressor arranged to receive air from an inlet that faces away from a direction of travel of the aircraft such that the air flowing to the auxiliary compressor has a lower mass flow rate of air at an inflight airspeed than a mass flow rate of air of a forward facing inlet so that the auxiliary compressor consumes less work from the turbine inflight.
1 . A gas turbine engine system adapted for use with an aircraft, the gas turbine engine system comprising:
a gas turbine engine including a gearbox and an engine core having a first compressor, a combustor, and a turbine coupled with the first compressor and the gearbox to drive the first compressor and the gearbox,
a second compressor operably coupled with the gas turbine engine and driven by the turbine via the gearbox during operation of the gas turbine engine to produce a flow of auxiliary compressed air usable by the aircraft, and
a nacelle extending from a fore end and an aft end thereof, a direction of a flow of air over the aircraft flowing from the fore end to the aft end, the nacelle arranged around the gas turbine engine and formed to define (i) a first inlet at the fore end of the nacelle and arranged to direct a first portion of the flow of air aft into the first compressor, (ii) an outlet formed in the aft end of the nacelle and configured to direct exhaust gases from the turbine away from the nacelle, and (iii) a second inlet in direct fluid communication with ambient air and arranged to direct a second portion of the flow of air into the second compressor, the second portion of the flow of air being separate from the first portion of the flow of air,
wherein the second inlet is oriented such that the second portion of the flow of air flows through the second inlet in a direction defined by the aft end of the nacelle toward the fore end of the nacelle.
2 . The gas turbine engine system of claim 1 , wherein the second inlet has a lower mass flow rate of the second portion of the flow of air at an inflight airspeed than a mass flow rate of the first portion of the flow of air at the first inlet due to the lower pressure so that the second compressor consumes less work from the turbine when the aircraft is inflight.
3 . The gas turbine engine system of claim 1 , wherein the nacelle further includes a plenum disposed between the second inlet and the second compressor configured to receive the second portion of the flow of air from the second inlet before the second portion of the flow of air enters the second compressor.
4 . The gas turbine engine system of claim 3 , wherein the plenum includes at least one anti-swirl structure configured to reduce distortion of the second portion of the flow of air.
5 . The gas turbine engine system of claim 4 , wherein the plenum includes a plurality of anti-swirl vanes.
6 . The gas turbine engine system of claim 1 , wherein the nacelle further includes a valve coupled to the second inlet and configured to move with respect to the nacelle to at least partially block the second portion of the flow of air through the second inlet.
7 . The gas turbine engine system of claim 6 , wherein the valve is configured to pivot or rotate with respect to the nacelle to at least partially block the second portion of the flow of air through the second inlet.
8 . The gas turbine engine system of claim 1 , where the second inlet includes at least one inlet guide vane that is configured to control an air flow rate into the second compressor.
9 . The gas turbine engine system of claim 1 , wherein the second inlet and the second compressor are oriented in the same direction.
10 . The gas turbine engine system of claim 1 , wherein the second inlet is oriented parallel to the direction of the flow of air over the nacelle.
11 . The gas turbine engine system of claim 1 , wherein the second inlet is disposed at an angle relative a central axis of the second compressor.
12 . The gas turbine engine of claim 1 , wherein the second portion of the flow of air does not flow through the engine core prior to flowing through the second compressor.
13 . A gas turbine engine system adapted for use with an aircraft, the gas turbine engine system comprising:
a gas turbine engine including a gearbox and an engine core having a first compressor, a combustor, and a turbine,
a second compressor operably coupled with the gas turbine engine and driven by the turbine via the gearbox, and
an airframe including fuselage, a wing, and a nacelle, a nose and a tail disposed aft of the nose such that a direction of a flow of air over the airframe moves from the nose of the airframe aft towards the tail, the nacelle extending from a fore end and an aft end thereof, the nacelle arranged around the gas turbine engine and formed to define a first inlet at the fore end of the nacelle and arranged to direct a first portion of the flow of air aft into the first compressor, and an outlet formed in the aft end of the nacelle and configured to direct exhaust gases from the turbine away from the nacelle,
wherein the airframe is formed to define a second inlet in direct fluid communication with ambient air and arranged to direct a second portion of the flow of air into the second compressor, the second portion of the flow of air being separate from the first portion of the flow of air, the second inlet is oriented such that the second portion of the flow of air flows through the second inlet in a direction defined by the aft end of the nacelle toward the fore end of the nacelle.
14 . The gas turbine engine system of claim 13 , wherein the second inlet is oriented parallel to the direction of the flow of air over the nacelle.
15 . The gas turbine engine system of claim 13 , wherein the second inlet is disposed at an angle relative a central axis of the second compressor.
16 . The gas turbine engine system of claim 13 , wherein the airframe further includes a valve coupled to the second inlet and configured to move with respect to the airframe to at least partially block the second portion of the flow of air through the second inlet.
17 . The gas turbine engine of claim 16 , wherein the valve is configured to move between an open position in which the second portion of the flow of air is free to enter the second inlet and a closed position in which the second portion of the flow of air is blocked from entering the second inlet.
18 . The gas turbine engine system of claim 17 , wherein the airframe further includes a plenum disposed between the valve and the second compressor and configured to receive the second portion of the flow of air from the second inlet before the second portion of the flow of air enters the second compressor.
19 . The gas turbine engine of claim 18 , wherein the plenum includes a plurality of anti-swirl vanes located therein.