Abstract: The gas turbine engine for an aircraft includes at least a low pressure spool with a low pressure turbine shaft operatively connected to at least one turbine, the low pressure turbine shaft rotatable about an engine axis, and a low pressure compressor operatively connected to a low pressure compressor shaft that is independently rotatable relative to the low pressure turbine shaft. A differential gearbox has an input operatively connected to the low pressure turbine shaft, a first output and a second output, the first output of the differential gearbox operatively connected to the low pressure compressor shaft and the second output of the differential gearbox operatively connected to an output shaft of the gas turbine engine. The differential gearbox permits the output shaft, the low pressure compressor shaft and the low pressure turbine shaft to rotate at different speeds.

Abstract: An inertial particle separator (IPS) for a gas turbine engine, has: inner and outer walls extending about a central axis, an inlet defined between the inner and outer walls and oriented axially; swirling vanes extending at least radially between the inner and outer walls and circumferentially distributed around the central axis, the swirling vanes configured for inducing a circumferential component in an airflow flowing between the swirling vanes; a plenum between the inner and outer walls downstream of the swirling vanes, the plenum circumferentially extending about the central axis, the outer wall converging toward the central axis in a direction of the airflow; and a splitter radially between the inner and outer walls downstream of the plenum and circumferentially extending around the central axis, a particle outlet radially between the splitter and the outer wall, an air outlet radially between the inner wall and the splitter.

Abstract: The aircraft can have a first engine secured to a first wing on a first side of a fuselage, and a second engine secured to a second wing on a second side of the fuselage, the second wing having a proximal end secured to the fuselage, and a distal end extending away from the fuselage. While operating the first engine, compressed gas can be conveyed from the first engine to a thrust generating device located at the distal end of the second wing.

Abstract: The gas turbine engine includes a combustion section including an annular swirl combustor having a combustor inlet, and a compressor section including a centrifugal compressor with an impeller, the impeller compressing and swirling an airflow and discharging the compressed and swirled airflow from the impeller outlet into the combustor inlet. The turbine section includes a radial turbine having a turbine fuel inlet and a turbine fuel outlet, the radial turbine receiving a flow of fuel at the turbine fuel inlet and discharging the flow of fuel from the turbine fuel outlet of the radial turbine into the combustor inlet.

Abstract: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. In one embodiment, such system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid-driven fluid propeller configured to drive the bleed air through the heat exchanger.

Abstract: The aircraft can have a first engine secured to a first wing on a first side of a fuselage, and a second engine secured to a second wing on a second side of the fuselage, the second wing having a proximal end secured to the fuselage, and a distal end extending away from the fuselage. While operating the first engine, compressed gas can be conveyed from the first engine to a thrust generating device located at the distal end of the second wing.

Abstract: An inertial particle separator (IPS) for a gas turbine engine, has: a plenum circumferentially extending about a central axis and defined between an outer wall and an inner wall, the plenum having an inlet facing a circumferential direction relative to the central axis, a radius of the outer wall decreasing in an axial direction relative to the central axis between the inlet and an annular splitter extending circumferentially around the central axis and located downstream of the inlet radially between the outer wall and the inner wall, a particle outlet including an annulus radially between the outer wall and the splitter, an air outlet fluidly connectable to a compressor of the gas turbine engine and defined radially between the splitter and the inner wall.

Abstract: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. The system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid propeller configured to drive the bleed air through the heat exchanger. The fluid propeller is disposed downstream of the heat exchanger.

Abstract: An inertial particle separator (IPS) for a gas turbine engine, has: inner and outer walls extending about a central axis, an inlet defined between the inner and outer walls and oriented axially; swirling vanes extending at least radially between the inner and outer walls and circumferentially distributed around the central axis, the swirling vanes configured for inducing a circumferential component in an airflow flowing between the swirling vanes; a plenum between the inner and outer walls downstream of the swirling vanes, the plenum circumferentially extending about the central axis, the outer wall converging toward the central axis in a direction of the airflow; and a splitter radially between the inner and outer walls downstream of the plenum and circumferentially extending around the central axis, a particle outlet radially between the splitter and the outer wall, an air outlet radially between the inner wall and the splitter.

Abstract: An inertial particle separator (IPS) for a gas turbine engine, has: a plenum circumferentially extending about a central axis and defined between an outer wall and an inner wall, the plenum having an inlet facing a circumferential direction relative to the central axis, a radius of the outer wall decreasing in an axial direction relative to the central axis between the inlet and an annular splitter extending circumferentially around the central axis and located downstream of the inlet radially between the outer wall and the inner wall, a particle outlet including an annulus radially between the outer wall and the splitter, an air outlet fluidly connectable to a compressor of the gas turbine engine and defined radially between the splitter and the inner wall.

Abstract: The gas turbine engine for an aircraft includes at least a low pressure spool with a low pressure turbine shaft operatively connected to at least one turbine, the low pressure turbine shaft rotatable about an engine axis, and a low pressure compressor operatively connected to a low pressure compressor shaft that is independently rotatable relative to the low pressure turbine shaft. A differential gearbox has an input operatively connected to the low pressure turbine shaft, a first output and a second output, the first output of the differential gearbox operatively connected to the low pressure compressor shaft and the second output of the differential gearbox operatively connected to an output shaft of the gas turbine engine. The differential gearbox permits the output shaft, the low pressure compressor shaft and the low pressure turbine shaft to rotate at different speeds.

Abstract: The gas turbine engine includes a combustion section including an annular swirl combustor having a combustor inlet, and a compressor section including a centrifugal compressor with an impeller, the impeller compressing and swirling an airflow and discharging the compressed and swirled airflow from the impeller outlet into the combustor inlet. The turbine section includes a radial turbine having a turbine fuel inlet and a turbine fuel outlet, the radial turbine receiving a flow of fuel at the turbine fuel inlet and discharging the flow of fuel from the turbine fuel outlet of the radial turbine into the combustor inlet.

Abstract: A gas turbine engine system includes a boost compressor configured to compress air; a combustor in which the compressed air is mixed with fuel and ignited to generate a stream of combustion gases; and a turbine configured to extract energy from the combustion gases, the turbine being drivingly coupled to the boost compressor and to an output shaft via a differential gearbox configured to apportion an input torque from the turbine between a first output torque applied to the output shaft and a second output torque applied to the boost compressor.

Abstract: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. The system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid propeller configured to drive the bleed air through the heat exchanger. The fluid propeller is disposed downstream of the heat exchanger.

Abstract: Systems and methods for conditioning a fluid using bleed air from a bypass duct of a turbofan engine are disclosed. In one embodiment, such system comprises a heat exchanger configured to facilitate heat transfer between a flow of bleed air from the bypass duct of the turbofan engine and the fluid, and a fluid-driven fluid propeller configured to drive the bleed air through the heat exchanger.

Abstract: An auxiliary power unit for an aircraft includes a rotary intermittent internal combustion engine drivingly engaged to an engine shaft, a turbine section having an inlet in fluid communication with an outlet of the engine(s), the turbine section including at least one turbine compounded with the engine shaft, and a compressor having an inlet in fluid communication with an environment of the aircraft and an outlet in fluid communication with a bleed duct for providing bleed air to the aircraft, the compressor having a compressor rotor connected to a compressor shaft, the compressor shaft drivingly engaged to the engine shaft. The driving engagement between the compressor shaft and the engine shaft is configurable to provide at least two alternate speed ratios between the compressor shaft and the engine shaft.

Abstract: The aircraft engine can have a core gas path extending sequentially across a core compressor, a core combustor, and a core turbine; a boost gas path extending from an intake to the core compressor, across a boost compressor, a bypass gas path extending from the intake to the core compressor, and a bypass valve operable to selectively open and close the bypass gas path. The intake flow can be directed either across the boost gas path for increased power output, or be directed to bypass the boost gas path via the bypass gas path.

Abstract: The gas turbine engine can have a core gas path extending sequentially across a core compressor, a core combustor, and a core turbine, an auxiliary air intake path and a bypass intake path leading in parallel to the core compressor, an auxiliary compressor in the auxiliary air intake path, an auxiliary gas path downstream of the core compressor, the auxiliary gas path extending in sequence across an auxiliary combustor and an auxiliary turbine, in parallel with the core combustor and core turbine, and valves operable to control the flow through the bypass gas path and the auxiliary gas path. Accordingly, the auxiliary components can be operated to increase power output, or deactivated while allowing the core components to run efficiently while meeting a lower power output.

Abstract: The aircraft engine can have a core gas path having a first combustor, a second gas path parallel to the core gas path, the second gas path having a second combustor, a turbine driven by the second gas path, a gearbox driven by the turbine, and a valve configured for selectively opening and closing the second gas path.

Abstract: The aircraft engine can have a core gas path extending from an intake across a core compressor, and then manifolding to a plurality turbine intake paths, each turbine intake path leading to a respective turbine unit via a respective combustor unit, and gearing drivingly connecting the collective rotary power of the turbine units to at least one power output shaft. During operation, the different core-turbine units can be operated simultaneously and efficiently, or one or more of the core-turbine units can be selectively shut down while the other core-turbine units continue to operate efficiently to lower the power output.