Abstract: An assembly including a support panel of an internal fixed structure, a thermal protection element and a plurality of fastening systems. Each fastening system includes a female part secured to the support panel and which has, in its interior, a circular shoulder, and a male part including a sleeve pierced with a main bore, a slide movable in translation in the main bore and a plurality of balls. For each ball, the sleeve has a secondary bore, and each ball is movable along the secondary bore between a locking position and an unlocking position. The slide is movable in translation between a blocking position and an unlocking position. Such an assembly makes it possible to fasten the thermal protection element to the support panel in a simple and rapid manner, also permitting easy removal if necessary.

Abstract: A fan for a gas turbine engine is provided. The fan includes a plurality of fan blades, a disk, and a trunnion mechanism for attaching the fan blades to the disk. The disk can be formed of a plurality of individual disk segments, with the trunnion mechanism attaching one of the plurality of fan blades to a respective disk segment. A retention member is also provided. The retention member includes a means for catching a portion of the trunnion mechanism should a primary attachment system of the trunnion mechanism fail.

Abstract: This control device for a gas turbine connected to a power generator capable of linking to a power system is configured to be able to switch operation modes of the gas turbine between a normal operation mode in which the gas turbine rotationally drives the power generator and a reverse power operation mode in which the gas turbine is rotationally driven by the power generator actuated as a motor by receiving power from the power system, the control device comprising: an index acquisition unit configured to acquire an index indicating the pressure ratio of a compressor of the gas turbine; and a signal output unit configured to output a signal for protecting the compressor when the index becomes less than a threshold value while the gas turbine is operating in the reverse power operation.

Abstract: A thrust unit for a propulsion device includes a thrust engine, arranged to provide a thrust force oriented in a direction so as to provide substantially vertical take-off and landing capability, and a deflector assembly comprising a pair of deflecting elements arranged to selectively divert the ejected fluid and movably mounted in the fluid outlet path. In order to reduce the bulk of the thrust unit and improve reliability and responsiveness, the invention relates more particularly to the positioning of the deflecting elements opposite the ejected fluid.

Abstract: The subject matter of this specification can be embodied in, among other things, an engine assembly having a nacelle surrounding an engine, and a thrust reverser coupled to the nacelle and having a thrust-reversing element movable relative to the nacelle between a stowed and deployed position, a first hydraulic actuator configured to move the thrust-reversing element between the stowed position and deployed position, the first hydraulic actuator being connected to a fluid source and a return reservoir, and a second hydraulic actuator configured to move the thrust-reversing element between the stowed and deployed position, the second hydraulic actuator being connected to the fluid source and the return reservoir, and a control system having an electrohydraulic servo valve operable to selectively route fluid between the fluid source, the second hydraulic actuator, and the return reservoir, and a controller configured to operate the electrohydraulic servo valve.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, a plurality of air scoops, and a plurality of exhaust gas conduit banks. The housing has an interior cavity, a bottom side, and first and second lateral sides. The exhaust gas conduit banks are disposed in the interior cavity of the housing and are configured to form an interior bypass air chamber. Each exhaust gas conduit bank includes alternating exhaust gas conduits and bypass air passages. The exhaust gas conduits extend axially between the forward and aft ends, and are configured to conduct exhaust gases to the nozzle. The bypass air passages are configured to receive bypass air from the air scoops and direct the bypass air between the exhaust gas conduits and into the interior bypass air chamber.

Abstract: A combustion system comprising: a combustion chamber extending in an axial direction between an inlet and an outlet, the combustion chamber configured to receive an airflow through the inlet and to discharge the airflow through the outlet; a fuel injection port configured to inject fuel into the airflow to form an air-fuel mixture; an ignition system for igniting the air-fuel mixture in the combustion chamber, the ignition system comprising an array of electrical plasma initiation points disposed downstream of the fuel injection port, and distributed radially and circumferentially around the combustion chamber, wherein each electrical plasma initiation point comprises a pair of electrodes configured to apply a voltage across an electrode gap between the pair of electrodes to produce plasma within the air-fuel mixture passing between the electrodes, thereby igniting the air-fuel mixture.

Abstract: An aircraft propulsion system is provided that includes a turbine engine and an exhaust gas condenser. The exhaust gas condenser includes a housing, a nozzle, inlet and outlet air scoops, and an exhaust gas conduit bank. The housing has an interior cavity, top and bottom sides, and first and second lateral sides. The exhaust gas conduit bank is disposed in the interior cavity of the housing. The exhaust gas conduit bank includes a plurality of exhaust gas conduits and bypass air passages that are disposed in an alternating configuration. The exhaust gas conduits extend axially between the forward and aft ends, are open at the forward end and at the nozzle. The bypass air passages receive bypass air from the inlet air scoop and direct the bypass air laterally across the condenser to the outlet air scoop.

Abstract: The subject matter of this specification can be embodied in, among other things, an engine assembly includes a nacelle configured to at least partially surround an engine, and a thrust reverser coupled to the nacelle, the thrust reverser having a thrust-reversing element movable relative to the nacelle between a stowed position and a deployed position, a hydraulic actuator operably coupled to move the thrust-reversing element between the stowed position and the deployed position, and a fluid control system configured to operate the hydraulic actuator, the fluid control system having an electrohydraulic servo valve operable to selectively route fluid between a pressurized fluid source, the hydraulic actuator, and a fluid return reservoir, and a bypass fluid line providing fluid communication between the hydraulic actuator and the fluid return reservoir independent of the electrohydraulic servo valve.

Abstract: An apparatus is provided for an aircraft propulsion system. This apparatus includes a variable area nozzle apparatus. The variable area nozzle apparatus includes a plurality of panels, a plurality of inter-panel members and an actuation system. The panels are arranged circumferentially about an axial centerline. The panels include a first panel and a second panel. The inter-panel members are arranged circumferentially about the axial centerline. The inter-panel members include a first inter-panel member between and connected to the first panel and the second panel. The first inter-panel member is configured from or otherwise includes an elastomeric material. The actuation system is configured to move the panels and the inter-panel members between a restricted flow arrangement and an unrestricted flow arrangement. The actuation system includes a first linkage and a second linkage. The first linkage is coupled to the first panel. The second linkage is coupled to the second panel.

Abstract: A nacelle system having a translating structure is disclosed. In various embodiments, the nacelle system includes a fixed structure; a thrust reverser having a translating sleeve configured to translate relative to the fixed structure and in response to a first dual-circuit hydraulic system; and a variable area fan nozzle having a translating nozzle configured to translate relative to the translating sleeve and in response to a second dual-circuit hydraulic system.

Abstract: A latch includes a handle, a hook, a first link, and a second link. The hook is rotatably mounted to the handle at a first rotational axis. The hook includes a hook body. The hook body includes a first lateral portion and a second lateral portion. The first lateral portion is laterally spaced from the second lateral portion. The first link is laterally between the first lateral portion and the second lateral portion. The first link is rotatably mounted to the first lateral portion and the second lateral portion at the first rotational axis. The second link is laterally between the first lateral portion and the second lateral portion. The second link is rotatably mounted to the first link at a second rotational axis. The second link is configured for translation relative to the first lateral portion and the second lateral portion.

Abstract: An engine nacelle assembly includes a nacelle that defines a flow path between an intake end and an exhaust end, and a nozzle that is disposed at the exhaust end of the nacelle about a central axis. The nozzle includes opposing thrust reverser openings and thrust reverser doors. A cross-sectional area transverse to the central axis proximate a forward side of each of the opposing thrust reverser openings is oval shaped with an offset radius within offset arc segments that are aligned with the opposing thrust reverser openings and a first radius outside of the offset arc segment. The first radius is less than the offset radius.

Abstract: A surge control system includes a rotor system with at least one compressor section and at least one turbine section operably coupled to a shaft. The surge control system also includes sensors configured to collect sensor data from the rotor system, an electric motor operably coupled to the rotor system, and a controller. The controller is operable to detect surge event from the sensor data, determine an amount of power to apply to the rotor system, and increase the amount of power provided to the rotor system to recover from the surge event.

Abstract: A latch mechanism is disclosed. In various embodiments, the latch mechanism includes a T-bolt body and a T-bolt comprising a longitudinal rod extending in a longitudinal direction from the T-bolt body and a transverse rod disposed at an end of the longitudinal rod. A safety keeper defining a load slot may be configured to receive the transverse rod whereby the T-bolt is configured to secure the T-bolt body with respect to the safety keeper.

Abstract: An assembly for a nacelle, which includes a fixed cowl, a movable cowl that is movable between a closed position and an open position, and including an upper edge and a lower edge and, between the two edges, a reinforcing rib. The fixed cowl includes, an upper reinforcing rib and a lower reinforcing rib, a hinge including a fixed part attached to the upper reinforcing rib of the fixed cowl, a movable part attached to the reinforcing rib of the movable cowl, and at least two locks, each one including a first part secured to the lower reinforcing rib and a second part secured to the reinforcing rib of the movable cowl. An assembly of this kind serves to provide access to the inside of the nacelle while ensuring that the cowls remain sufficiently rigid.

Abstract: A system for controlling blade clearances within a gas turbine engine includes a rotor disk and a rotor blade coupled to the rotor disk. Additionally, the system includes an outer turbine component positioned outward of the rotor blade such that a clearance is defined between the rotor blade and the outer turbine component. Furthermore, the system includes a heat exchanger configured to receive a flow of cooling air bled from the gas turbine engine and transfer heat from the received flow of the cooling air to a flow of coolant to generate cooled cooling air. Moreover, the system includes a valve configured to control the flow of the coolant to the heat exchanger. In this respect, the cooled cooling air is supplied to at least one of the rotor disk or the rotor blade to adjust the clearance between the rotor blade and the outer turbine component.

Abstract: A rocket engine system having a heating system configured to heat an oxidizer; a combustion section having a flow path from an upstream inlet section through a restricted throat section to a downstream outlet section, the combustion section configured to accelerate the heated oxidizer as an oxidizer stream within the flow path in response to flow dynamics to supersonic speed; and a fuel system configured to introduce a fuel into the flow path to mix supersonically with the heated oxidizer to define a combined fuel and oxidizer stream at a first supersonic speed. The combined fuel and oxidizer stream undergoes a deflagration to denotation transition in the combustion section defined by an oblique shock wave and a normal shock wave that interact to achieve a standing detonation wave generally at an upstream portion of the restricted throat section configured such that combustion exits the downstream outlet section to provide thrust.

Abstract: An annular or semi-annular acoustic attenuation structure includes an inner skin and an outer skin, defining between them an annular or semi-annular volume, and a plurality of partitions extending in the space in a radial direction perpendicular to the inner and outer skins and in a circumferential direction. The inner and outer skins and the partitions are made of composite material including a fibrous reinforcement densified by a matrix. The fibrous reinforcement of each partition includes yarns or fibers extending into the fibrous reinforcement of at least one of the inner or outer skins.

Abstract: A thrust reverser system for jet aircraft comprising an exhaust tailpipe mounted to the turbine engine aft turbine flange and two clamshell doors, actuators and a locking system to prevent inadvertent deployment of the clamshell doors in-flight. Two improved design clamshell doors configurations, either one or a combination of both, mounted on either side of the top and bottom of the exhaust tailpipe, fitted with two patented design actuators mounted one on each side of the external sides of the tailpipe between the clamshell doors and the tailpipe, possibly in a depression in the tailpipe called blister, assuming them to be hydraulic actuators for discussion purposes. The actuators drive the clamshell doors using improved floating linkages loosely pivoted to the exterior of the exhaust tailpipe.