Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.

Abstract: A de-icing device for an air intake of an aircraft jet engine nacelle extending along an axis in which a flow of air flows from upstream to downstream, the intake comprising an inner wall and an outer wall connected by a leading edge, the inner wall comprising a plurality of blowing lines, each blowing line comprising a plurality of through-openings configured to blow elementary streams from the hot air source in order to de-ice said inner wall, the blowing lines being parallel to one another in a cylindrical projection plane, each blowing line having a depth defined along the axis X and a length defined along the axis Y in the cylindrical projection plane, two adjacent blowing lines being spaced apart by a distance, the ratio of the distances L3/D3 being between 1 and 2.

Abstract: Aircraft propulsion systems including a closed loop-supercritical fluid system having a turbine, a cooler heat exchanger, a compressor, and a recovery heat exchanger arranged along a closed-loop flow path of a supercritical fluid. A shaft is operably coupled to the turbine and configured to be rotationally driven by the turbine. A fan is configured to generate thrust, the fan operably coupled to the shaft to be rotationally driven by the shaft. A burner is configured to combust a fuel and air from the fan to generate a combusted gas and supply said combusted gas to the recovery heat exchanger of the closed loop-supercritical fluid system and out an exhaust nozzle.

Abstract: Variable bleed valves with inner wall controlled-flow circuits are disclosed. An example apparatus disclosed herein includes a casing segment defining a first flow path, a variable bleed valve port defining a second flow path, and a channel formed in the casing segment, the channel including a first opening into the first flow path, and a second opening into the second flow path, the channel defining a third flow path between the first opening and the second opening.

Abstract: A flow control actuator includes a first side wall, a second side wall opposite and substantially parallel to the first side wall, an upstream wall mechanically coupled to upstream ends of the first and second side walls, and a downstream cap mechanically coupled to downstream ends of the first and second side walls. The first side wall, the second side wall, the upstream wall and the downstream cap collectively define an interior of the flow control actuator. An energy source is disposed in at least one of the first sidewall and the second sidewall. At least one fuel injector is disposed in the upstream wall, the first sidewall and/or the second sidewall for dispersing fuel into the flow control actuator. At least one air inlet is disposed in the upstream wall, the first sidewall and/or the second sidewall for introducing air into the flow control actuator. Fuel from fuel injector and air from the air inlet are ignited in the flow control actuator.

Abstract: An example thermal management system may include a first heat exchanger including a bleed air inlet configured to receive input bleed air from a gas turbine engine and a bleed air outlet configured to output cooled bleed air. A turbine including a turbine inlet may be fluidically coupled to the bleed air outlet. The turbine may be configured to drive a shaft mechanically coupled to the turbine in response to expansion of the cooled bleed air through the turbine. A second heat exchanger may include an expanded bleed air inlet fluidically coupled to a turbine outlet of the turbine. The second heat exchanger may be configured to extract heat from at least one heat source using the expanded bleed air.

Abstract: A rocket motor includes at least one inert rod. The inert rod has a groove that extends along the length of the inert rod. The groove may be machined by a lathe, die, and/or CNC machine, such that the groove is configured helically around the rod. A reactive wire is inserted into the groove along the length of the inert rod. The grooved inert rod, and the reactive wire together constitute the burn rate enhancer assembly. The rocket motor is configured such that the burn rate enhancer assembly is inserted into a rocket motor casing. The rocket motor casing is then filled with a burnable propellant grain, the highly loaded grain, which is in a liquid or semi-solid state. The highly loaded grain then cures in the rocket motor casing around the burn rate enhancer assembly.

Abstract: Systems and methods to increase intake air flow to a gas turbine engine of a hydraulic fracturing unit when positioned in an enclosure may include providing an intake expansion assembly to enhance intake air flow to the gas turbine engine. The intake expansion assembly may include an intake expansion wall defining a plurality of intake ports positioned to supply intake air to the gas turbine engine. The intake expansion assembly also may include one or more actuators connected to a main housing of the enclosure and the intake expansion assembly. The one or more actuators may be positioned to cause the intake expansion wall to move relative to the main housing between a first position preventing air flow through the plurality of intake ports and a second position providing air flow through the plurality of intake ports to an interior of the enclosure.

Abstract: A bleed air cooling system for a gas turbine engine includes one or more bleed ports located at one or more axial locations of the gas turbine engine to divert a bleed airflow from a gas turbine engine flowpath, a bleed outlet located at a cooling location of the gas turbine engine and a bleed duct in fluid communication with the bleed port and the configured to convey the bleed airflow from the bleed port to the bleed outlet. One or more safety sensors are configured to sense operational characteristics of the gas turbine engine, and a controller is operably connected to the one or more safety sensors and configured to evaluate the sensed operational characteristics for anomalies in operation of the bleed air cooling system.

Abstract: A method for in-flight monitoring of the health of an aero gas turbine engine including in flow series a compressor system, a core gas generator including a combustor, and a turbine, and further including a shaft system connecting the turbine to the compressor system and forming a torque path there between. The method includes: measuring a first rotational speed of a forward portion of the shaft system; measuring a second rotational speed of a rear portion of the shaft system; measuring other operational parameters of the engine; calculating from the measured rotational speeds the power delivered by the torque path from the turbine to the compressor system; calculating from the other measured operational parameters the power delivered to the turbine by the core gas generator; correlating the power delivered by the torque path at a given time with the power delivered to the turbine at that time.

Abstract: A turbine engine has: a compressor; a combustor; a turbine, a gas flowpath passing consecutively through the compressor, combustor, and turbine; and inlet member along the gas flowpath upstream of the compressor. The inlet member includes the unitarily-formed single piece combination of: a three dimensional (3D) lattice portion; and a nose cap body surrounding the lattice portion.

Abstract: In a convergent-divergent flap pair for a turbojet engine nozzle of the variable-geometry convergent-divergent type, the convergent flap and the divergent flap including respective ducts for circulation of cooling air connected to one another to allow cooling of each of the flaps. The duct of the divergent flap includes an impingement cooling cavity defined by two walls of lateral ends provided with air passage orifices through which the impingement cooling cavity opens towards the outside.

Abstract: A gas turbine engine comprising a variable geometry combustor having pilot fuel injectors and main fuel injectors; a fuel metering system configured to control fuel flow to the pilot fuel injectors and the main fuel injectors; a variable geometry airflow arrangement for the variable geometry combustor, which is configured to vary the airflow through the pilot fuel injectors and/or the main fuel injectors; a control system configured to control the variable geometry airflow arrangement in dependence upon airflow delivered to the combustor, the fuel flow to the pilot fuel injectors and the main fuel injectors, and a target index of soot emissions, thereby controlling airflow through the pilot fuel injectors and/or the main fuel injectors and hence the quantity of soot produced by combustion.

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: An aircraft system includes a turbine engine having a compressor, a combustor having an inlet and an outlet, and a turbine having an inlet portion and an outlet portion. An auxiliary power unit (APU) is operatively connected to the turbine engine. The APU includes a compressor portion, a generator, and a turbine portion. The compressor portion is operatively connected to the turbine portion through the generator. A source of cryogenic fluid is operatively connected to the turbine engine and the APU. A heat exchange member includes an inlet section operatively connected to the source of cryogenic fluid, a first outlet section operatively connected to the turbine engine and a second outlet section operatively connected to the compressor portion.

Abstract: A multi-engine aircraft includes a first engine drivingly engaged to a common rotatable load and a second engine drivingly engaged to the common rotatable load, the second engine having a bleed air system and a control system in communication with a compressed air switching system. The control system controls operation of the second engine and/or the compressed air switching system. The compressed air switching system includes a switching valve that is displaceable between at least a first position and a second position, the first position interconnecting a lower pressure inlet and a switch outlet, and the second position interconnecting a high pressure inlet and the switch outlet. The switch outlet is in communication with the bleed air system of the second engine. The control system actuates the switching valve to switch between the first and second positions.

Abstract: A rocket booster has an annular shape, with a casing defining an annular space therewithin, and a solid rocket fuel in the annular spacing. The rocket booster also includes one or more nozzle pieces, mechanically coupled to the casing, that define one or more nozzles at the aft side of the rocket booster. The rocket booster may be mechanically coupled to an object protruding from the back of a fuselage of a flight vehicle, such as a missile. For example, the rocket booster may be placed around an aft turbojet nozzle of the flight vehicle. This allows the rocket booster to be used in situations where primary propulsion must be running both before and after (and perhaps during) the firing of the rocket booster.

Abstract: A hybrid engine including features to meet aircraft thrust, passenger airflow, and fuel cell requirements. The engine includes a combustor burning the same fuel as the fuel cell. The engine has electric motors to utilize the power output of the fuel cell. The engine shafts have sprags to allow motors to drive the compressors and over run the turbines. The engine has variable flowpath geometry to bypass the combustor.

Abstract: An engine system of an aircraft includes a gas turbine engine comprising at least one spool and at least one electric machine operably coupled with the at least one spool. A controller is configured to detect if the at least one spool of the gas turbine engine is in or is approaching an overspeed condition and apply a load to the at least one spool via the at least one electric machine.

Abstract: Electroaerodynamic devices and their methods of operation are disclosed. In one embodiment, ions are formed by dielectric barrier discharge using a time varying voltage differential applied between a first electrode and a second electrode. The ions are then accelerated in a downstream direction using a second voltage differential applied between a third electrode and the first and/or second electrodes, where the third electrode is located down stream from the first and second electrodes. The ions may then collide with naturally charged molecules and/or atoms within a fluid to accelerate the fluid in the downstream to create an ionic wind and an associated thrust.