Abstract: An assembly is provided for an aircraft propulsion system. This assembly includes an engine core, a bypass duct and a bleed circuit. The engine core includes a compressor section, a combustor section, a turbine section and a core flowpath extending sequentially through the compressor section, the combustor section and the turbine section. The bypass duct includes a bypass flowpath outside of the engine core. The bleed circuit includes a bleed passage and a flow regulator. The bleed circuit is configured to direct bypass gas through the bleed passage from the bypass flowpath into the core flowpath when the flow regulator is in an open position. The bleed circuit is configured to cutoff gas flow through the bleed passage between the bypass flowpath and the core flowpath when the flow regulator is in a closed position.

Abstract: A heat exchanger is provided. The heat exchanger includes one or more exchanger units that each have a core and manifolds. The core of an exchanger unit is formed by multiple unit cells coupled together in flow communication to create a flow distribution grid. Each unit cell has at a first primary channel, a second primary channel, a first secondary channel in flow communication with the first primary channel, and a second secondary channel in flow communication with the second primary channel. The first secondary channel traverses through the second primary channel and the second secondary channel traverses through the first primary channel. Each manifold includes two chambers for separating fluids flowing through the heat exchanger, with one chamber being in flow communication with one of the primary channels and having one or more tubes traversing therethrough to provide flow communication between the other primary channel and the other chamber.

Abstract: An aircraft propulsion system in includes a compressor section, a heat addition system, and an expansion section in serial flow arrangement, wherein the heat addition system is configured to receive a flow of gaseous fuel from a fuel system, and wherein the flow of gaseous fuel is provided to the heat addition system to generate combustion gases. A lubricant system is configured to provide a flow of lubricant through the propulsion system. A hydraulic system is configured to provide a flow of hydraulic fluid to an actuator at the propulsion system. A thermal management system is configured to provide a flow of fluid in thermal communication with the lubricant system and the hydraulic system. The thermal management system comprises a turbine configured to receive the flow of fluid via a fluid circuit extended in fluid communication from the compressor section.

Abstract: An engine system includes an engine core assembly, a fuel system and a cooling circuit. The engine core assembly includes a core flowpath, a compressor section, a combustor section and a turbine section. The combustor section includes a combustor adjacent a plenum. The fuel system includes a fuel flowpath, a fuel reservoir, a fuel-cooling air heat exchanger and a fuel injector. The fuel flowpath is configured to direct hydrogen fuel, received from the fuel reservoir, through the fuel-cooling air heat exchanger to the fuel injector. The fuel injector is configured to introduce the hydrogen fuel into the combustor. The cooling circuit is configured to direct cooling air, bled from the plenum, through the fuel-cooling air heat exchanger to the turbine section. The fuel-cooling air heat exchanger is configured to exchange heat energy between the hydrogen fuel flowing within the fuel flowpath and the cooling air flowing within the cooling circuit.

Abstract: A gas turbine engine comprises a bypass duct and a heat exchanger assembly, the heat exchanger assembly comprising a heat exchanger and a heat exchanger duct having an inlet region, an inflection region and an outlet region. A direction of a centreline of the heat exchanger duct has a tangential component with respect to a principal rotational axis of the gas turbine engine at one or more of the inlet region, the inflection region and the outlet region. The heat exchanger is disposed within the inflection region and configured to transfer heat generated by the gas turbine engine into the flow of air as it passes through the inflection region.

Abstract: An assembly is provided for a turbine engine. This assembly includes an engine core extending along an axis. The engine core includes a first compressor section, a second compressor section, a flowpath, a bleed port, an inner passage, an outer passage and a flow diverter. The flowpath extends longitudinally through the first compressor section and the second compressor section. The bleed port fluidly couples the flowpath to the inner passage and the outer passage in parallel. The bleed port is located longitudinally along the flowpath between the first compressor section and the second compressor section. The flow diverter is located at an inlet into the bleed port from the flowpath. The flow diverter is configured to move between a first position and a second position.

Abstract: A hybrid-electric aircraft system is provided and includes first and second hybrid-electric engines, each of which includes an electric motor to drive operations thereof, and a supplemental power unit (SPU). The SPU is configured as a thermal engine paired with a generator and is configured to generate electrical power. The first and second hybrid-electric engines are operable normally and off, respectively, with electrical power generated by the SPU being diverted to the electric motor of the second hybrid-electric engine.

Abstract: A vehicle drive having a vacuum system for a vehicle having an internal combustion engine that is connected to a supply air line and to an exhaust air line. The vacuum system has an ejector. A propellant gas line opens into a nozzle channel of a propellant nozzle of the ejector. The nozzle channel opens into a mixing chamber of the ejector through a nozzle opening. A suction line opens into the mixing chamber and the mixing chamber on an outlet side opens directly or indirectly into a mixed gas line. The mixed gas line opens into the supply air line. The suction line is connected to a vacuum consumer, and the nozzle opening in the propellant nozzle is formed by a nozzle edge having teeth and/or corrugations.

Abstract: A system is provided for an aircraft. This system includes a propulsor rotor and a powerplant configured to drive rotation of the propulsor rotor. The powerplant includes a turbo-compounded intermittent internal combustion engine and a dual rotor electric machine. The turbo-compounded intermittent internal combustion engine includes a turbine rotor operatively coupled to the propulsor rotor through the dual rotor electric machine. The dual rotor electric machine includes a first rotor, a second rotor, a first stator and a second stator. The first stator and the second stator are arranged radially between and axially aligned with the first rotor and the second rotor.

Abstract: A gas turbine engine includes a compressor section having rotor shaft assembly and a stator shroud assembly including a stator shroud casing surrounding the compressor rotor shaft assembly, a compressor flow passage being defined between the compressor rotor shaft assembly and the stator shroud casing. A compressor bleed air system includes a compressor bleed air duct that includes a sloped inlet portion, such as a NACA submerged inlet portion, in the stator shroud casing. Alternatively, stator vanes at a compressor bleed air passage may have a NACA inlet scoop and an airflow passage through the stator vane providing a bleed airflow to a cavity, which has airflow passages providing bleed air to an upstream side of the bleed air passage.

Abstract: A propulsion system for an aircraft includes an intercooling system with a plurality of injectors for injecting an intercooling waterflow into a portion of a compressor passage. The plurality of injectors are spaced apart about the compressor passage and the injection of the intercooling water flow is varied between the plurality of injectors to influence flow that is communicated through the compressor passage.

Abstract: A cooling system for a turbine engine. The turbine engine includes a compressor, a turbine, and a shaft that drivingly couples the compressor and the turbine. The cooling system includes one or more cavities of the compressor. The cooling system includes a shaft flowpath defined in the shaft. The shaft includes one or more shaft apertures that provide fluid communication between the shaft flowpath and the one or more cavities. Air passes through the one or more shaft apertures and the shaft flowpath during a shutdown of the turbine engine to cool the one or more cavities.

Abstract: The present invention relates to an exhaust-gas treatment device for an aircraft engine, comprising an exhaust-gas channel, through which an exhaust gas of the aircraft engine flows, and a first cooling unit for cooling with ambient air, characterized by a second cooling unit, which is downstream of the first cooling unit with respect to an exhaust-gas flow in the exhaust-gas channel.

Abstract: Variable bleed valves with struts for aerodynamic stability are described. An apparatus described herein includes a variable bleed valve port, and a strut disposed within the variable bleed valve port, the strut defining a plane and a plurality of holes disposed perpendicularly to the plane, the plurality of holes disposed in a pattern to dampen an acoustic response associated with the variable bleed valve port.

Abstract: An auxiliary power unit (10) for an aircraft (1) comprises a gas turbine engine comprising an engine core (12). The engine core comprises a core compressor (14) and core turbine (18) coupled by a core shaft (24). The auxiliary power unit further comprises a load spool comprising a load compressor (30), a load turbine (20) and a permanent magnet electric machine (28), each being coupled by a load shaft (26). The load shaft (26) and core shaft (24) are configured to rotate independently of one another.

Abstract: A medium supply system for delivering one or more mediums to an environmental control system of a vehicle includes a manifold arranged in fluid communication with an inlet of the environmental control system, a first medium source for delivering a first medium to the manifold via a first port; a second medium source for delivering a second medium to the manifold via a second port, and a third medium source for delivering a third medium to the manifold via a third port. At least one of a temperature and pressure of each of the first medium, the second medium, and the third medium varies.

Abstract: The invention concerns a bypass turbomachine (1) with a primary flow path and a secondary flow path, comprising: —a low-pressure body comprising a low-pressure compressor (120) connected to a low-pressure turbine (122) via a low-pressure shaft (124), —a high-pressure body comprising a high-pressure compressor (130) connected to a high pressure turbine (132), via a high-pressure shaft (134), —a low-pressure power take-off system (220) comprising an electrical generator (226), configured to take power (W12) from the low-pressure body, wherein—the turbomachine comprises a debris removal system (500), located between the two compressors (226, 236), —the low-pressure power take-off system (220) is configured to take power (W12) from the low-pressure shaft (124) using the resistive torque of the electrical generator (226), in order to avoid a risk of surging.

Abstract: Diffuser apparatus for turbine frames and outlet guide vanes are disclosed. An example diffuser is integrated into a turbine engine. The example diffuser includes a plurality of struts extending between an inner portion and an outer portion of the turbine engine; and a plurality of flowpath surfaces supported by the plurality of struts. The example plurality of flowpath surfaces extends circumferentially within the turbine engine at a plurality of radial spans to control diffusion of air flow within the turbine engine.

Abstract: An axial turbomachine comprising a compression device with a variable-orientation vane rectifier stage; a combustion chamber downstream of the compression device; a turbine device downstream of the combustion chamber; and a leak passage of air compressed by the compression device, between the compression device and the turbine device for cooling the turbine device; and where the rectifier stage is configured to modulate the section of the leak passage according to the orientation of the vanes of the stage.

Abstract: A valve for an air system in an aircraft engine, comprising: a housing defining a chamber having a valve axis; a body within the chamber about a piston axis collinear with the valve axis, extending from a first surface to a second surface, defining a bore extending from the first to the second surface, having a mating connector defined by the second surface and located radially outward of the bore relative to the piston axis; and a sleeve extending from a first end matingly engaged with the mating connector to a second end along a sleeve axis collinear with the valve axis, the first end axially stacked on the body via the first surface to define a first distance between the first end and the first surface, and via the second surface to define a second distance between the first end and the second surface greater than the first distance.

Abstract: A gas turbine engine includes a propulsor delivering air into a core engine inward of an outer core panel. The core engine has a compressor section, a combustor and a turbine section. A compressor bleed system includes a bleed valve on a conduit communicating with compressed air from the compressor section. The bleed valve is configured to deliver air into a chamber where it can flow outwardly of the core engine at an exit location radially inward of the outer core panel.

Abstract: A secondary air system (SAS) of an aircraft engine that produces secondary airflow from a source of secondary air includes a hollow strut and one or more SAS feed pipes upstream thereof. The hollow strut extends radially through the main gas path of the engine and defines therein a strut conduit extending between a strut inlet and a strut outlet at opposite ends of the hollow strut. The strut outlet is in fluid flow communication with a buffer cavity for feeding the secondary airflow to the engine core. The SAS feed pipe includes an inlet receiving the secondary airflow from the source of secondary air, and an outlet in fluid flow communication with the strut inlet to feed the secondary airflow into the strut conduit. The SAS feed pipe has a sonic orifice therein, between the inlet and the outlet thereof.

Abstract: A turbomachine has a bearing in a bearing compartment. A first seal system isolates the bearing compartment and has: a first runner; a second runner; an inner seal encircling the first runner; and an outer seal encircling the second runner. The inner seal has a static clearance with the first runner. The outer seal has a static clearance with the second runner, greater than the inner seal static clearance. The first runner has an inner portion cantilevered inward. The second runner has an outer portion cantilevered outward.

Abstract: Disclosed is an air intake of an aircraft turbine engine nacelle having a lip, a downstream portion and an internal partition separating the lip and the downstream portion, the lip delimiting an annular recess, the downstream portion having a downstream inner wall and a downstream outer wall, the air intake having an injection channel for injecting a hot air stream into the annular recess, a passage opening formed in the internal partition, an outlet opening formed in the downstream outer wall and a discharge channel for discharging the hot air stream mounted in the downstream portion and having a first end connected to the internal partition, a second end connected to the downstream outer wall and a main body having at least one flexible portion.

Abstract: An airflow arrangement for a gas turbine engine in which air can be selectively directed to and from a bypass channel of the engine and into a duct communicating air into a compressor of a gas turbine engine.

Abstract: Methods, apparatuses and systems for providing dehumidification providing dehumidification for gas detecting components are disclosed herein. An example apparatus may comprise: a flow sensing component configured to detect a flow rate associated with a flowing media in a flow channel of the apparatus; and a controller component in electronic communication with the flow sensing component that is configured to receive a flow rate indication from the flow sensing component, and in response to determining that the flow rate indication satisfies a null condition threshold, apply an exponential smoothing function with a variable alpha value to condition an output signal of the apparatus.

Abstract: A gas turbine engine comprises a turbomachine defining a core flow therethrough during operation. A flow tap is provided in fluid communication with the turbomachine, wherein the flow tap is configured to receive a portion of the core flow therethrough as a bleed flow. A bleed assembly includes a machine load, a bleed flow machine, and a bleed regulator. The bleed flow machine is disposed in fluid communication with the turbomachine through the flow tap, and is configured to drive the machine load. The bleed regulator is configured to regulate a bleed output provided to the bleed flow machine by controlling a capture rate of the bleed flow by the bleed flow machine.

Abstract: A method is provided of controlling a cooled cooling air system for an aeronautical gas turbine engine. The method includes: receiving data indicative of an ambient condition of the aeronautical gas turbine engine, data indicative of a deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or a combination thereof; and modifying a cooling capacity of the cooled cooling air system in response to the received data indicative of the ambient condition of the aeronautical gas turbine engine, data indicative of the deterioration parameter of the aeronautical gas turbine engine, data indicative of an operating condition of the aeronautical gas turbine engine, or the combination thereof.

Abstract: A gas turbine engine includes a propulsor having a plurality of blades, a compressor section including a first compressor and a second compressor aft of the first compressor. The first compressor includes at least one array of first variable guide vanes that control operation of the first compressor. The second compressor includes at least one array of second variable guide vanes that control operation of the second compressor. A turbine section includes a first turbine and a second turbine. A geared architecture is driven by the second turbine for rotating the propulsor.

Abstract: A fuel tank inerting system is provided including an air flow comprising air from a first source having a first temperature and air from a second source having a second temperature. The second temperature is cooler than the first temperature. At least one separating module is configured to separate an inert gas from the air flow.

Abstract: In accordance with at least one aspect of this disclosure, there is provided a system for an aircraft engine. In embodiments, the system includes an accessory box and a fuel accessory located in an interior space within the accessory box, where a vent is defined through a wall of the accessory box. In embodiments, the vent includes a plurality of holes or slots in an outer wall of the accessory box for passage of gaseous fuel from the interior space. In embodiments, the vent is configured for passive ventilation of the interior space.

Abstract: A turbine engine assembly includes a tap is at a location upstream of the combustor section where a bleed airflow is drawn. The bleed air is pressurized in an auxiliary compressor section, heated in an exhaust heat exchanger and expanded through a power turbine that is coupled to drive the auxiliary compressor section.

Abstract: An aircraft power plant for a fuel cell including a turbo assembly, a compressor assembly, a turbo assembly, a compressor assembly controller, a first stage turbo assembly and compressor assembly operation configured to generate a first stage compressed fluid generated from ambient air and excess oxygen exhausted from a fuel cell of an aircraft power plant. A second stage turbo assembly and compressor assembly operation configured to receive the first stage compressed fluid, and a controller bleed valve coupled with the first stage turbo assembly and compressor assembly and the second stage turbo assembly and compressor assembly. An oxygen supply system, the oxygen supply system fluidically coupled with the first stage turbo assembly and compressor assembly wherein a first compressed oxygen is generated by the first stage turbo assembly is combined with a second compressed oxygen generated by the second stage turbo assembly to generate a combined oxygen controlled by the controller bleed valve.

Abstract: A centrifugal compressor for HVAC application includes a rotary component rotatable about an axis, a static component, and a brush seal fixed to one of the static component and the rotary component. The brush seal includes bristles that contact the other of the static component and the rotary component.

Abstract: A gas turbine engine comprising: a duct extending about an axis, the duct including an outer-duct wall having an interior-duct surface circumscribing an interior of the duct and an exterior-duct surface radially outward of the interior-duct surface relative to the axis, the outer-duct wall defining an offtake opening extending from the interior-duct surface to the exterior-duct surface, the offtake opening in fluid communication between an offtake location inside the duct and outside the duct, and a bleed air offtake assembly including: an air line in fluid communication with inside the duct via the offtake opening, the air line having a first-line end defining a line inlet proximate to the outer-duct wall and a second-line end spaced from the first-line end; a valve located outside the duct and fluidly connected to the air line via the second-line end, and a conduit having a conduit inlet in fluid communication with inside the air line at a resonance location between the first-line end and the second-line en

Abstract: A gas turbine engine includes a turbine rotor connected to a main compressor rotor. A tail cone is mounted inward of an exhaust core flow. A generator rotor is adjacent a generator stator. The generator rotor and stator are mounted within the tail cone. A passage connects a bypass flow path to the tail cone. A cooling air compressor is operable within the passage. The turbine rotor drives a shaft to drive the generator rotor and the cooling compressor. A method is also disclosed.

Abstract: A blade for an engine includes an airfoil body having a pressure side and a suction side, a base, and a rib located on the pressure side of the airfoil body. The rib includes a radially outer surface inclined radially outwardly with respect to the pressure side of the airfoil body and a scoop formed by the radially outer surface. The radially outer surface is inclined radially outward with respect to a normal axis to the pressure side of the airfoil body. The rib is angled at a positive angle with respect to the platform. An engine for and a method of directing ingestion material in an engine may employ the rib.

Abstract: A method for thermal management for an aircraft includes extracting a flow of compressed fluid from a compressor section of a propulsion system. The flow of compressed fluid is passed through an anti-ice system. The flow of compressed fluid flows from the anti-ice system to a turbine. The flow of compressed fluid is expanded across the turbine. The expanded flow of compressed fluid then flows to thermal communication with a thermal load.

Abstract: The gas turbine engine can have a pneumatic actuator; an intake device secured to a gas path wall delimiting the gas path, the intake device having a tubular body protruding from the gas path wall into the gas path and an inlet aperture formed in the tubular body, the inlet aperture spaced-apart from the gas path wall and facing downstream relative a flow orientation of the gas path, the intake device having an internal conduit extending from the inlet aperture, along the tubular body, to an outlet across the gas path wall; and a fluid line fluidly connecting the outlet of the intake device to the pneumatic actuator.

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

Abstract: A turboprop gas turbine engine mountable to an aircraft has an engine core and a gearbox driving a propeller, the engine core and the gearbox being enclosed within a nacelle. The propeller is located rearward of the gearbox and the engine core relative to a direction of travel of the aircraft. An air intake is disposed within the nacelle and formed to direct ambient air into the engine core. The air intake includes an air inlet duct, having a forward-facing intake inlet receiving the ambient air, with an upstream section and a downstream section. The upstream section is in fluid communication with the intake inlet and extends downstream from the intake inlet. The downstream section fluidly connects to and directs air from the upstream section into the engine air inlet. A second air outlet duct is located within the nacelle and directs air into an air-cooled-oil-cooler (ACOC).

Abstract: A duct for a bleed system of an aircraft, wherein the duct extends from an inlet section to an outlet section along a longitudinal axis, and wherein the duct comprises a continuous piece arranged on and protruding from the internal wall of the duct. The duct is subject to temperature gradients in order to reduce the temperature of the warmest airflow closer to the inner wall rather than rapidly mix the airflow.

Abstract: The auxiliary duct can branch radially outwardly from the bypass duct, have a proximal end fluidly connecting the bypass duct and a distal end, a valve can be activatable to selectively open and close the auxiliary passage, and a structure can protrude partially from the auxiliary duct into the auxiliary passage, the structure spaced apart from the proximal end, between the proximal end and the valve, the structure generating lesser pressure losses when flow in the auxiliary passage is directed towards the distal end than when the flow is directed towards the proximal end.

Abstract: A system includes a turbo pump to convert compressed gas into power, a storage tank to store the compressed gas, and a fire suppression control valve having a closed position in which the compressed gas is prevented from flowing to the cargo compartment and an open position in which the compressed gas is ported to the cargo compartment to suppress a fire. The system also includes a pump control valve having a closed position in which the compressed gas is prevented from flowing to the turbo pump and an open position in which the compressed gas is ported to the turbo pump to cause the turbo pump to convert the compressed gas into the power. The system also includes an OBIGGS to convert bleed air from a gas turbine engine into an inert gas to provide low rate discharge (LRD) fire suppression to the cargo compartment.

Abstract: Systems and methods are disclosed for controlling surge margin in the compressor section of a gas turbine engine. A first compressor section and a second compressor section are in fluid communication with a bypass conduit. An auxiliary turbine and discharge conduit are positioned in the bypass conduit. Fluid flow from the compressor sections into the bypass conduit is controlled by bypass control valves.

Abstract: The disclosure describes a system that includes an auxiliary power unit (APU), an APU throttle valve, and an environmental control system (ECS) bypass valve. The APU is configured to receive cabin discharge air from an aircraft cabin and receive ECS supply air from an air pressurization system (APS). The APU throttle valve is configured to control flow of cabin discharge air from the cabin to the APU. The ECS bypass valve configured to control flow of ECS supply air from the APS to the APU.

Abstract: A gas turbine engine has a housing exposed to a high temperature environment. The housing has a circumferential wall extending around the engine centerline and circumscribing an oil cavity. The wall has a sealing interface at an inner diameter thereof, the sealing interface having a central axis offset from the engine centerline. A boss is formed on the wall on the offset side relative to the engine centerline and a probe is mounted to the boss. The probe projects into the oil cavity. The oil in the oil cavity thermally shields the probe from the high temperature environment.

Abstract: A combustion engine including at least one combustion chamber, a first bleed air supply fluidly coupled to a portion of the combustion engine upstream the combustion chamber, a second bleed air supply fluidly coupled to a portion of the combustion engine downstream the combustion chamber, a first thermal bus, and a turbomachine including a compressor, a rotary pump, and a first turbine, with the compressor and rotary pump in serial flow arrangement and the rotary pump being fluidly coupled to the first thermal bus.

Abstract: An anti-icing system for a gas turbine system includes multiple nozzles, wherein each nozzle of the multiple nozzles includes one or more outlets that are configured to inject a heated fluid into an airflow within an air intake conduit. The anti-icing system also includes multiple plates disposed upstream of the one or more outlets, wherein each plate of the multiple plates extends laterally across the air intake conduit and is vertically spaced apart from one or more adjacent plates to define one or more vertically-extending gaps. The multiple plates are configured to direct the airflow through the one or more vertically-extending gaps to spread the airflow upstream of the one or more outlets to facilitate mixing of the heated fluid and the airflow.

Abstract: There are described methods, systems, and assemblies for monitoring a bleed valve of a gas turbine engine. The method comprises determining a rate of change of a gas generator speed of the gas turbine engine; determining a rate of change of a parameter indicative of engine power of the gas turbine engine; comparing at least one ratio based on the rate of change of the gas generator speed and the rate of change of the parameter indicative of engine power to at least one range of values; detecting a modulation delay of the bleed valve when the at least one ratio is within the at least one range of values; and transmitting a signal indicative of the bleed valve malfunction in response to detecting the modulation delay.