Abstract: A solid fuel propulsion system includes a housing, an air inlet configured to enable air to flow into the housing, an air duct, a rotation detonation chamber, and a nozzle assembly. The air duct is disposed in the housing and is in fluid communication with the air inlet. The air duct is configured to supply air for combustion of a solid fuel configured to sublimate into a gaseous fuel mixture. The rotation detonation chamber is disposed in the housing aft of the air duct and is configured to combust the gaseous fuel mixture and air to generate a rotating detonation wave. The nozzle assembly is configured to expand and exhaust the flow prompted by the rotating detonation wave to generate thrust.

Abstract: Pressure regulators together with associated methods and systems are provided. An embodiment of a pressure regulator includes a chamber having an inlet port for receiving a fluid and an outlet port for delivering the fluid at a regulated pressure. The pressure regulator includes a diaphragm defining at least part of the chamber, and the diaphragm defining a first orifice therethrough. The pressure regulator includes a spring configured to interact with the diaphragm. The pressure regulator includes an interface operatively disposed between the spring and the diaphragm. The interface includes a contact surface for contacting the diaphragm, and the contact surface is convex toward the diaphragm. The interface defines a second orifice therethrough that is in fluid communication with the first orifice in the diaphragm. The first and second orifices define a relief port for venting the chamber to an ambient environment external to the chamber.

Abstract: A power plant is provided including a heat recovery steam generator positioned to receive a flow of an exhaust gas and having a heating surface, an exhaust gas recirculation line branching off at an extraction point within the heat recovery steam generator and opening into the heat recovery steam generator at an injection point upstream of the extraction point within the heat recovery steam generator, a thermal storage system arranged between the extraction point and the injection point in the exhaust gas recirculation line wherein the thermal energy storage system stores thermal energy, and a blower arranged in the exhaust gas recirculation line to push air or exhaust gas through the thermal energy storage system.

Abstract: An engine driven environmental control system (ECS) air circuit includes a gas turbine engine having a compressor section. The compressor section includes a plurality of compressor bleeds. A selection valve selectively connects each of said bleeds to an input of an intercooler. A second valve is configured to selectively connect an output of said intercooler to at least one auxiliary compressor. The output of each of the at least one auxiliary compressors is connected to an ECS air input.

Abstract: A solid-propellant gas generator assembly may comprise a bulkhead having an orifice disposed in a housing. The bulkhead may be disposed between a first end and a second end of the housing. The bulkhead and the first end may define a propellant cavity. The bulkhead and the second end may define a pressure chamber. A fast burning solid-propellant may be disposed in the propellant cavity. The solid-propellant gas generator assembly may be configured to replace a slow burning solid-propellant gas generator system in a solid-propellent gas generator system.

Abstract: An exhaust diffuser hub disposed at a longitudinal center of an exhaust diffuser is provided. The exhaust diffuser hub includes a hub extension extending from a downstream end thereof in a longitudinal direction of the exhaust diffuser. A transverse cross-sectional area of the hub extension is smaller than a transverse cross-sectional area of the hub.

Abstract: A system and method of increasing efficiency and power output of a gas turbine system using a compressed air storage system including delivering a compressed air charge from the compressed air storage system, the compressed air charge having a pressure greater than ambient pressure and a temperature less than ambient temperature, the compressed air charge being delivered to the gas turbine and the compressed air charge operable to cool at least a portion of the gas turbine.

Abstract: A gas turbine engine exhaust system comprises an instrumented exhaust duct including a set of exhaust temperature probes installed at the exhaust end of the duct in an air space between the engine outer case and a surrounding engine nacelle. A shield is removably installed at the end of the exhaust duct to protect the probes and associated wiring in the air space.

Abstract: The invention relates to a propelling nozzle for a turbofan engine on a supersonic aircraft, the propelling nozzle including: a propelling nozzle wall, a duct, which is radially outwardly bounded by the propelling nozzle wall, and a central body arranged in the duct. According to the invention, the central body is connected to the propelling nozzle wall via at least one brace.

Abstract: Liquid fuel supply system (12) for a combustion system (14), in particular a gas turbine, including at least one storage tank (16) for liquid fuel supplying at least one injector (34) connected to a combustion chamber (32) of the combustion system (14), said liquid fuel supply system (12) including a first piping section (18) disposed downstream of the tank (16) and a second piping section (20) disposed downstream of the first piping section (18) and upstream of fuel nozzle (34) in each combustion chamber (32), said first piping section (18) including at least one pressurizing means (22), and at least one injecting point or entering (24) for a water-soluble product, and the second piping section (20) including a mixing and distribution flow device (26) configured to create an emulsion and distributing the emulsion flow rate to at least one piping (28) connected to said nozzle (34).

Abstract: A propulsion unit includes a gas turbine engine arranged along an axis and an exhaust system coupled to the gas turbine engine. The gas turbine engine includes an engine core configured to discharge a core flow and a fan configured to be driven by the engine core to discharge a bypass flow. The exhaust system receives the mixed bypass and core flows from the gas turbine engine.

Abstract: A twin-engine system includes a gas turbine engine comprising a core and a first output shaft drivable by the core. An electric engine has an electric motor configured to drive a second output shaft. A reduction gear box (RGB) has an RGB input drivingly engaged to both the first output shaft and the second output shaft. The RGB has an RGB output to provide rotational output to a rotatable load.

Abstract: A ring assembly for holding a double-skin combustor liner has an inner skin of ceramic matrix composite and an outer skin. The assembly comprises an annular end body, a first projecting portion projecting relative to the annular end body at least partially in an axial direction, the first projecting portion configured for supporting the inner skin, a second projecting portion projecting relative to the annular end body at least partially in the axial direction, the second projecting portion configured for supporting the outer skin. The first and second projecting portions are radially spaced apart to provide a radial spacing between the inner skin and the outer skin. An expansion joint is between the ring assembly and the inner skin configured to permit motion resulting from a variation in thermal expansion between the inner skin and the ring assembly.

Abstract: An active device for attenuating the acoustic emissions of an aircraft turbojet engine includes circulation conduits for a pressurized air flow rate supplying rotary elements each having a pulsation system for the delivered air. The rotary elements are controlled in amplitude and phase and deliver, to outlet diffusers, a pulsed air flow rate with a pulsation at the frequency of the noise to be attenuated having an amplitude and a phase adjusted according to a local feedback law with microphones to attenuate the radiated acoustic power.

Abstract: An air-breathing turbojet engine (101) for a hypersonic vehicle is shown. The engine comprises a pump for pumping a cryogenic fuel, an inlet (102) configured to compress inlet air by one or more shocks, a cooler (103) to cool the compressed inlet air using the cryogenic fuel, and a turbo-compressor (104) to compress the air further. A combustor (105) receives compressed cooled air from the turbo-compressor and a first portion of the cryogenic fuel for combustion. A first turbine (106) expands and is driven by combustion products, and a second turbine (107) expands and is driven by a second portion of the cryogenic fuel. The first turbine and the second turbine drive the turbo-compressor via a shaft. An afterburner (109) receives combustion products from the first turbine and the second portion of the cryogenic fuel from the second turbine for combustion therein.

Abstract: Herein provided are methods and systems for a method for controlling autothrottle of an engine. A digital power request is obtained from an autothrottle controller, the digital power request based on an autothrottle input to the autothrottle controller. A manual input mode for the engine is terminated, the manual input mode based on a second power request obtained from a manual input associated with the engine. An autothrottle mode for the engine is engaged to control the engine based on the digital power request.

Abstract: A mounting apparatus for mounting a gas turbine engine to an aircraft. The mounting apparatus includes a thrust strut operably connected to the gas turbine engine and the aircraft. The thrust strut defines a thrust strut axis. The mounting apparatus further includes a restraining structure. The restraining structure includes a bracket circumferentially disposed on the thrust strut and at least one elongate member connected to the bracket and the gas turbine engine. The restraining structure radially and circumferentially constrains the thrust strut with respect to the thrust strut axis of the thrust strut while allowing the thrust strut to move in the direction of thrust strut axis.

Abstract: Turbine engine (10) with a contra-rotating turbine for an aircraft, the turbine engine comprising a contra-rotating turbine (22) whose first rotor (22a) is configured to rotate in a first direction of rotation and is connected to a first turbine shaft (36), and a second rotor (22b) is configured to rotate in an opposite direction of rotation and is connected to a second turbine shaft (38), the first rotor comprising turbine wheels (36a) inserted between turbine wheels (38a) of the second rotor, the turbine engine further comprising a mechanical reduction gear (42) with an epicyclic gear train which comprises a sun gear (44) driven in rotation by said second shaft, a ring gear (40) driven in rotation by said first shaft, and a planet carrier (46) fixed to a stator casing (28) of the turbine engine situated downstream from the contra-rotating turbine with respect to direction of flow of the gases in the turbine engine, the turbine engine further comprising a bearing (56) for guiding the second shaft with res

Abstract: To prevent or reduce an oversensitive change in the degree of opening of a bleed valve regulating the flow volume of a cooling air used to heat a fuel. A gas turbine includes: a compressor compressing an air; a combustor burning a fuel mixed with the air compressed by the compressor; a turbine configured to be driven by a combustion gas generated by the combustor; a heat exchanger causing the fuel to be supplied to the combustor to be heated through a cooling air bled from the compressor to be supplied to the turbine; a bleed valve regulating the flow volume of the cooling air; a first sensor measuring a fuel state value; a second sensor measuring a cooling air state value; and a control computer controlling the degree of opening of the bleed valve.

Abstract: A method of monitoring bleed air provided from a first engine to a second engine of a multi-engine aircraft includes operating the first engine in a powered mode to provide motive power to the multi-engine aircraft; and when the first engine is in the powered mode: actuating an air valve to open an air flow path from a compressor section of the first engine to a plurality of parts of the second engine, after the step of actuating the air valve, determining a change in a temperature in a main gas path of the first engine at a location in the main gas path at or downstream of a combustor of the first engine, and in response to determining that the change is below a threshold, executing an action with respect to the first engine, the air valve, and/or the second engine.