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 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 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 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.

Abstract: The invention relates to a bleed air extraction device for a turbomachine, which has: an axial compressor, formed in a flow path and having at least one compressor stage, which comprises a rotor and a stator, and a bleed air duct, which is provided and designed to guide a bleed air flow branched off from the flow path of the axial compressor. In this case, the bleed air duct comprises an inlet opening, which is formed downstream of a stator of the axial compressor in the radially outer flow path boundary, an axially forward wall adjoining the inlet opening, and an axially rearward wall adjoining the inlet opening. Guide means are provided, which are provided and designed for the purpose of guiding at least a portion of the bleed air flow branched off from the flow path in the direction of the axially forward wall of the bleed air duct.

Abstract: A duct for evacuating hot air from an aircraft engine is extended by a movable stack that can project from the wall such that the hot air is ejected a distance from the wall without risk of damaging the wall. The stack can, however, be retracted by a control device under circumstances of moderate engine speed, with the advantage that the drag of the wall, generally an outer nacelle, is then reduced.

Abstract: In addition to a first fuel gas heater utilizing the heated water from the outlet of an economizer of a heat recovery steam boiler, there is provided a second fuel gas heater utilizing as the heat source the bleed air of a compressor of a gas turbine. A control device opens a bleed air control valve of the piping for supplying bleed air to the second fuel gas heater at the time of starting the gas turbine combined cycle system to heat a fuel gas by the bleed air.

Abstract: Disclosed is a hybrid power generation facility. The hybrid power generation facility includes a gas turbine including a compressor configured to compress air introduced from an outside, a combustor configured to mix the compressed air with fuel and to combust the air and fuel mixture, and a turbine configured to produce power with first combustion gas discharged from the combustor, a boiler including a combustion chamber and configured to burn a mixture of the first combustion gas and air, a first water heat exchanger configured to pass second combustion gas discharged from the boiler and to heat water through heat exchange with the second combustion gas, a water supply device configured to supply water to the first water heat exchanger, a steam turbine through which steam generated in the combustion chamber passes, and a first air preheater configured to pass second combustion gas discharged from the first water heat exchanger and to pass air supplied to the boiler.

Abstract: A propulsion system is provided including a first vane extended along the radial direction. The first vane is configured to rotate relative to a vane axis extended along the radial direction. A second vane is extended along the radial direction. The second vane is positioned aft along the axial direction of the first vane. The second vane forms an inlet opening proximate to a second vane leading edge, and the second vane forms an outlet opening proximate to a second vane trailing edge. The inlet opening and the outlet opening together allow a flow of fluid through the second vane. A heat exchanger is positioned within the second vane. The inlet opening and the outlet opening allow the flow of fluid in fluid communication with the heat exchanger.

Abstract: A gas turbine engine includes a compressor section having a high pressure compressor, the high pressure compressor including an aft-most compressor stage. The gas turbine engine also includes a combustion section having a stage of discharge nozzles, the stage of discharge nozzles located downstream of the aft-most compressor stage and upstream of a diffuser cavity. The gas turbine engine also includes a high pressure spool drivingly coupled to the high pressure compressor, the high pressure spool forming in part a compressor discharge pressure seal and including a forward spool section. The forward spool section extends between the compressor discharge pressure seal and the aft-most compressor stage, the forward spool section defining an airflow cavity for providing a cooling airflow from the diffuser cavity to the aft-most compressor stage.

Abstract: A conduit assembly may comprise: a pipe; a plurality of hollow passages disposed through the pipe; and a plurality of flow guides disposed in the pipe, each flow guide in the plurality of flow guides at least partially defining a respective hollow passage in the plurality of hollow passages. The conduit assembly may act as a heat exchanger.

Abstract: A method of controlling a gas turbine engine capable of operating in at least a high output power range, a medium-high output power range, a medium-low output power range and a low output power range. The method includes during the medium-high output power range bleeding a gas from a downstream part of the compressor to an upstream part of the compressor so that a first predetermined temperature of the combustor is maintained, during the medium-low output power range bleeding a gas from a downstream part of the compressor to an upstream part of the compressor and bleeding a gas from the downstream part of the compressor to the exhaust so that a second predetermined temperature of the combustor is maintained.

Abstract: A compound engine assembly for use as an auxiliary power unit for an aircraft and including an engine core with internal combustion engine(s), a compressor having an outlet in fluid communication with an engine core inlet, a bleed conduit in fluid communication with the compressor outlet through a bleed air valve, and a turbine section having an inlet in fluid communication with the engine core outlet and configured to compound power with the engine core. The turbine section may include a first stage turbine having an inlet in fluid communication with the engine core outlet and a second stage turbine having an inlet in fluid communication the first stage turbine outlet. A method of providing compressed air and electrical power to an aircraft is also discussed.

Abstract: A supersonic gas turbine engine for an aircraft that comprises a nacelle, a fan, an engine core including a primary duct configured to guide a core airflow through the engine core, a bypass duct extending between the engine core and an engine casing and configured to guide a bypass airflow through the bypass duct, an intake located upstream of the fan, and a tertiary airflow duct extending between the engine casing and the nacelle and configured to guide a tertiary airflow. The intake is configured to extract air from the intake and guide it to the tertiary airflow duct in which the extracted air flows as tertiary airflow. It is provided that at least one heat exchanger is mounted in the tertiary airflow duct.

Abstract: A gas turbine engine section includes a plurality of spaced rotor stages, with a static guide vane intermediate the spaced rotor stages. The static guide vane provides swirl into air passing toward a downstream one of the spaced rotor stages, and an outer housing surrounding the spaced rotor stages. A diverter diverts a portion of air radially outwardly through the outer housing, and across at least one heat exchanger. The diverted air passes back into a duct radially inwardly through the outer housing, and is exhausted toward the downstream one of the spaced rotor stages.

Abstract: A thermal management system for one or more aircraft components includes a bleed airflow source at a gas turbine engine of the aircraft, one or more turbines configured to expand the bleed airflow, thus lowering a temperature of the bleed airflow, the bleed airflow driving rotation of the one or more turbines. One or more heat exchangers are in fluid communication with the one or more turbines. The one or more heat exchangers are configured to exchange thermal energy between the bleed airflow and a thermal load. One or more electrical generators are operably connected to the one or more turbines. The one or more electrical generators are configured to convert rotational energy of the one or more turbines to electrical energy.

Abstract: A gas turbine engine comprising a central power shaft; a rotatable turbine impeller on the shaft, a compressor configured to receive power from the central power shaft, the compressed air exiting the compressor entering a cavity of the impeller. Ejector(s) mounted on a periphery of the impeller having a combustion chamber including an outer wall; a mixing chamber downstream of the combustion chamber, a passageway to the mixing chamber from outside the combustion chamber for ejected (outside) air to enter the ejector and travel to the mixing chamber where the ejected air mixes with a flow of hot gases that has exited the combustion chamber to create a mixed flow of gases. A convergent-divergent nozzle at an exit of the mixing chamber accelerates the mixed flow of gases moving through the nozzle, thereby creating a reaction thrust and a moment of force on the shaft.

Abstract: A blade track for a gas turbine engine includes a plurality of blade track segments. The blade track segments are arranged circumferentially around a central axis to form the blade track.

Abstract: A method of controlling a gas turbine engine capable of operating in at least a high, medium-high, medium, medium-low, and low output power ranges. The method includes during the medium-high output power range varying the angle of the variable guide vanes so that a third predetermined temperature of the combustor is maintained, during the medium output power range the variable guide vanes are closed and bleeding a gas from a downstream part of the compressor to an upstream part of the compressor so that a first predetermined temperature of the combustor is maintained, during the medium-low output power range the variable guide vanes are closed and bleeding a gas from a downstream part of the compressor to an upstream part of the compressor and bleeding a gas from the downstream part of the compressor to the exhaust so that a second predetermined temperature of the combustor is maintained.

Abstract: A cooling system for a gas turbine engine comprises a passage capable of receiving cooling air, a compartment radially adjacent thereto and axially aligned therewith, an opening therebetween, a valve within the opening, and a heat exchanger received in the compartment. The valve is moveable between a maximum open position and a minimum open position for increasing or decreasing airflow from the passage into the compartment. At the valve minimum open position, a leakage path is provided between the passage and the compartment, whereby cooling air is capable of passing from the passage to the compartment and toward the heat exchanger at all valve positions. A gas turbine engine is also disclosed.

Abstract: A vane for a gas turbine engine, the vane having: an airfoil; and a root portion disposed on a side of the airfoil and including a platform, the platform having a vane forward rail and an extension extending from the platform, the extension defining portions of an outer diameter platform cavity and an airfoil leading edge cavity. The extension extends from the platform such that an upper portion of each of the outer diameter platform cavity and the airfoil leading edge cavity is spaced equidistant from the platform.

Abstract: A gas turbine engine with a bearing that is supplied with fuel and compressed air for both cooling and lubrication of the bearing. The bearing housing contains a serpentine flow cooling circuit with each leg formed from a plurality of parallel flow channels, and where fuel is supplied to the cooling circuit and then discharged into a combustor. Some of the fuel is also bled off and passed into the bearing compartment along with compressed air bled off from the compressor to both cool and lubricate the bearing.

Abstract: An assembly is provided for a turbine engine with an axial centerline. This assembly includes a turbine engine structure and a valve assembly. The turbine engine structure includes an outer duct wall, an inner duct wall, a first flow path and a second flow path. The inner duct wall is radially inward of the outer duct wall. The first flow path is radially inward of the inner duct wall. The second flow path is radially outward of the inner duct wall and is radially inward of the outer duct wall. The valve assembly includes a valve element and a valve actuator. The valve element is configured to regulate flow of fluid between the first flow path and the second flow path. The valve actuator is configured to move the valve element. The valve actuator is positioned entirely radially outward of the outer duct wall.

Abstract: The invention relates to a method for supplying air at a controlled temperature to a cabin of a surface vehicle in which at least one air cycle device is used, comprising at least one motorized turbocompressor. The air inlet of the turbine is arranged to receive a compressed airflow from the compressor. At least one exchanger is interposed between the air outlet of the compressor and the air inlet of the turbine. The air inlet of the compressor is arranged to receive air at a pressure greater than or equal to atmospheric pressure. The invention likewise relates to a surface vehicle comprising at least one such air cycle device.

Abstract: A turbine ring assembly including ring sectors forming a turbine ring and a ring support structure, each ring sector having, along a section plane defined by an axial direction and a radial direction of the ring, a portion forming an annular base with, in the radial direction an inner face defining the inner face of the ring and an outer face from which a first and a second attachment tabs protrude, the structure including a central shroud from which a first and a second radial clamps protrude between which the first and second attachment tabs of each ring sector are maintained. It includes a first and a second annular flanges removably fastened to the first radial clamp of the central shroud and separated from each other by a contact abutment.

Abstract: Various embodiments may include a drive system for an aircraft comprising: an electrical generator; and an internal combustion engine with a gas expansion turbine. The generator includes a rotor rotating on a rotor shaft. The gas expansion turbine rotates on a turbine shaft. The rotor shaft and the turbine shaft are couplable to transmit torque such that the internal combustion engine can deliver mechanical energy to the generator. A supply air stream impinges at least one component of the generator.

Abstract: A bleed valve system comprises a duct, featuring a central longitudinal axis and allowing a main flow of fluid to pass from a first environment at a first static pressure to a second environment at a second static pressure along a bleed direction, a valve comprising a valve member arranged within the duct between the first and second environments and movable to partially obstruct the duct and deviate the main flow of fluid to direct at least a part of it towards a portion of an internal wall of the duct; and an ejector, arranged within the duct, downstream of the valve member and offset from the central longitudinal axis in correspondence of said portion of the internal wall, adapted to supply an additional flow of fluid within the duct to accelerate the main flow of fluid and reduce the second static pressure.

Abstract: A CAES power generation device includes a plurality of compression/expansion/combined machines, a pressure accumulation unit for storing compressed air, a plurality of heat exchangers, a heat storage unit for storing a heating medium, and a plurality of first containers 10 having a rectangular parallelepiped shape for accommodating the plurality of compression/expansion/combined machines and the plurality of heat exchangers. The plurality of first containers are arranged side by side so that long side surfaces face each other. The long side surface of the first container is provided with at least one vent being an outlet of a pipe for communicating the compression/expansion/combined machine with the outside of the container.

Abstract: A thermal management system for an aircraft engine. The thermal management system extracts interstage engine bleed air from the engine's compressor and performs three stage cooling of the interstage engine bleed air using two heat exchangers and a vapor cycle system evaporator. The cooled air is then further cooled by expansion through a set of turbines and subsequently used as a heat sink for heat loads on the aircraft.

Abstract: A gas turbine engine includes: a compressor, a combustor, and a turbine arranged in sequential flow relationship along a primary flowpath, the turbine being connected in mechanical driving relationship to the compressor, so as to define at least one engine rotor that is rotatable about a centerline axis of the engine; a secondary flowpath connected in flow communication with the primary flowpath; and an air cycle machine including an air cycle rotor carrying an at least one air cycle compressor and at least one air cycle expander, wherein: the air cycle rotor is coupled in mechanical driving relationship with the at least one engine rotor; the air cycle rotor is coupled in fluid flow communication with the secondary flowpath; and the air cycle rotor is coupled in fluid flow communication with at least one heat exchanger.

Abstract: The invention relates to an intermediate housing hub (2) comprising a discharge stream conduit (18) having a given length (L1), and discharge fins (22) extending in the discharge stream conduit (18) and having a given chord length (L2), the chord length (L2) of at least one of the discharge fins (22) being at least equal to 50% of the length of conduit (L1).

Abstract: An aircraft gas turbine engine system comprises first and second gas turbine engines connected by an inter-engine gas path. The first gas turbine engine has a first spool with a first compressor section, and a second spool with a second compressor section downstream of and rotationally independent from the first compressor section. The second gas turbine engine is configured to provide power to at least one of the first and second spools of the first gas turbine engine. The inter-engine gas path is disposed to receive gas flow bled from a bleed location in the first gas turbine engine downstream of the first compressor section, and to supply this gas flow to an inlet of the second gas turbine engine.

Abstract: A thrust reverser system for a gas turbine engine includes at least one hinge coupled to the thrust reverser system so as to be adjacent to at least one opening defined in the thrust reverser system. The thrust reverser system includes at least one body coupled to the at least one hinge. The at least one body has a first body end and an opposing second body end. The body pivotally coupled to the hinge such that a portion of the body is positionable within the at least one opening and the body includes at least one counterweight at the first body end or the second body end. The body is positioned within the at least one opening based on an operating condition of the gas turbine engine.

Abstract: Systems and methods for cooling electrical components disposed in a jet engine. An example system includes an evaporation chamber configured to contain the electrical components in contact with a coolant liquid. The coolant vapor formed during the heat transfer from the electrical components to the coolant liquid flows to a condenser assembly having a fuel-cooled condenser and an air-cooled condenser. The air-cooled condenser cools the coolant vapor to condensation using either fan stream air or engine bleed air from the intermediate pressure compressor or the high pressure compressor. An air cycle machine cools the engine bleed air. A controller may be used to select a coolant source for condensing the coolant vapor based on operating conditions of the aircraft. Spent air from the air-cooled condenser may be recycled back to the engine for engine cooling, added thrust, oil sump buffering, oil or fuel cooling, or blade tip clearance control.

Abstract: An aircraft incorporating a bleed system for extracting compressed air from the aircraft main engines to be used as a source of pressurized air for the aircraft. The bleed air system includes a first pre-cooler installed at one of the main engines nacelle and coupled with the bleed duct, and adapted for cooling down the bleed air extracted from the main engine, and a second pre-cooler installed at the pylon and coupled with bleed duct and downstream the first pre-cooler. The working temperature of the aircraft bleed system is reduced, down to max 200° C., so that the dimensions of an Over Heat Detection System (OHDS) is reduced.

Abstract: The present disclosure is directed to a gas turbine including a primary compressor, a combustion section disposed downstream from the primary compressor and a turbine disposed downstream from the combustion section. A pipe assembly defines an extraction-air circuit between an extraction port of the primary compressor and an inlet port of the turbine. The gas turbine further includes an extraction-air conditioning system. The extraction-air conditioning system includes a secondary compressor having an inlet and an outlet. Both the inlet and the outlet are fluidly coupled to the extraction-air circuit at respective locations defined along the pipe assembly between the extraction port of the primary compressor and the inlet port of the turbine.

Abstract: A heat exchanger system includes a plurality of additively manufactured heat transfer surfaces in a plurality of sectors, a first of the plurality of sectors including a first subset of the plurality of the additively manufactured heat transfer surfaces having a characteristic different than a characteristic in a second of the plurality of sectors.

Abstract: A cabin air control system for an aircraft comprising a plurality of compressed air sources, each operable to provide compressed air, at least one air outlet, a conduit system connected to the air sources and the air outlet and arranged to conduct air from the air sources downstream to the air outlet, and a plurality of particulate matter detectors configured to detect the presence of particulate matter in air from the air sources. Each detector is associated with a different predefined subset of the air sources by being arranged in the conduit system at a location from which it is only possible to reach the associated subset of the air sources when traversing the conduit system from the respective location towards the air sources in an upstream direction. Any two subsets of the different subsets associated with the different particulate matter detectors include different ones of the air sources.

Abstract: A pressurization system includes a first compressor that receives a ram air, a fan air, or engine air; a first turbine that is on a common shaft with the first compressor and wherein the first turbine receives an engine air; a main heat exchanger downstream of the first compressor and the first turbine; an internal environment suitable for human occupants and downstream of the main heat exchanger; a second turbine downstream of the internal environment; the second turbine may be on the common shaft with the first compressor and first turbine; or a generator downstream of the second turbine; a motor downstream of the generator; and wherein the motor drives the first compressor.

Abstract: A system includes an air turbine starter having an inlet, a turbine air exhaust outlet, an output shaft, and a turbine in fluid communication with the inlet and the turbine air exhaust outlet. The turbine is operably coupled to the output shaft. The system also includes an outlet valve assembly configured to adjust an exhaust area of the turbine air exhaust outlet.

Abstract: A system and method for harnessing latent heat to generate energy. The system and method provide a fully closed latent heat recovery system that utilizes a vapor source to generate vapor. A plurality of conduits carries the vapor and resultant gas, expanded energy, and condensate to: a vapor expander, a compressor, a heat exchanger, an accumulator, and a vapor condenser for expansion, compression, and conversion between states of the vapor. The latent heat generated from the expansion and energy release from the vapors and gases produces work for driving a load.