Abstract: A cyrogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid.

Abstract: Turbojet engine nacelle (1) comprising an air intake structure (4) able to channel a flow of air towards a fan of the turbojet engine, and a central structure (5) intended to surround said fan and to which the air intake structure is attached in such a way as to ensure aerodynamic continuity, the central structure comprising, on the one hand, a casing intended to surround the fan and, on the other hand, an outer structure, characterized in that the air intake structure comprises a longitudinal outer panel (40) incorporating an air intake lip (4a) and which may extend so far as to incorporate at least part of the outer structure of the central section and beyond, said longitudinal outer panel being translationally mobile between an operating position in which the outer panel ensures the aerodynamic continuity of the outside of the nacelle (1), and a maintenance position in which the outer panel (40) is set away from the outer structure of the central section and the air intake lip (4a) is set away from the inn

Abstract: The invention relates to a field of feeding reaction engines, and in particular a turbopump 9 for feeding at least one combustion chamber 2 with a first propellant, the turbopump comprising a pump 9a, a turbine 9b to the pump 9a in order to drive it, and a variable and/or disengageable braking device 11.

Abstract: A method for flushing a section of a fuel system of a burner of a gas turbine is provided. The method includes the steps of providing a feed line for feeding fuel from a fuel source to the burner, providing a first feed point close to the burner and a second feed point remote from the burner in the feed line in each case for feeding a medium into the feed line, feeding the medium via the first feed point, and discharging, via the burner, the fuel which is in the feed line between the first feed point and the burner, and also feeding the medium via the second feed point, and discharging, via the first feed point, the fuel which is in the feed line between the second feed point and the first feed point.

Abstract: A turbomachine includes a compressor, a combustor operatively connected to the compressor, an end cover mounted to the combustor, and an injection nozzle assembly operatively connected to the combustor. The injection nozzle assembly includes a cap member having a first surface that extends to a second surface. The cap member further includes a plurality of openings. A plurality of bundled mini-tube assemblies are detachably mounted in the plurality of openings in the cap member. Each of the plurality of bundled mini-tube assemblies includes a main body section having a first end section and a second end section. A fluid plenum is arranged within the main body section. A plurality of tubes extend between the first and second end sections. Each of the plurality of tubes is fluidly connected to the fluid plenum.

Abstract: A cooling system for a transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have a multi-panel outer wall formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that at least one cooling chamber is formed between the inner and intermediate panels. The transition duct may also include an outer panel. The inner, intermediate and outer panels may include one or more metering holes for passing cooling fluids between cooling chambers for cooling the panels. The intermediate and outer panels may be secured with an attachment system coupling the panels to the inner panel such that the intermediate and outer panels may move in-plane.

Abstract: A combustor with a liner where each of the walls has a respective circumferential row of dilution holes defined therethrough adjacent a junction between the primary zone and the dilution zone. In the primary zone, the inner surface of each of the walls is covered by at least one heat shield attached thereto and spaced apart therefrom to allow air circulation between the inner surface and the at least one heat shield, the walls each having a plurality of cooling holes defined therethrough having a smaller diameter than that of the dilution holes. In the dilution zone, the inner surface of each of the walls is free of heat shields, and the walls each have a plurality of effusion cooling holes defined therethrough and having a smaller diameter than that of the dilution holes.

Abstract: A system for mounting a fuel injector to an engine casing with an aperture includes a fuel injector having a flange for mounting to the casing at the aperture. The flange allows it to pass through the aperture. An intermediate ring mediates the mounting of the flange to the casing at the aperture. The intermediate ring is inside the casing at the aperture and defines an opening from which the fuel injector extends into the engine with the flange positioned inward of the ring. The flange is dismountably sealed to an inner side of the ring, and the ring is dismountably sealed to an inner side of the casing to mount the fuel injector to the casing. On dismounting the flange from the ring, the fuel injector can be displaced to allow the ring, to be moved away from the aperture. The fuel injector can then be withdrawn.

Abstract: The present invention relates to a mechanical system and method that modulates airflow in an aircraft inlet diffuser that is used in conjunction with an aircraft engine that integrates both a center turbine engine and a high Mach engine such as a constant volume combustor (CVC) arrangement or ramjet arrangement with intakes formed co-centrically about the turbine. The modulation system uses two articulating components, a movable air flow duct and an articulating cone. The air flow duct, in a first position, is in exclusive air flow communication with the circular intake face of the turbine in a first position to receive air from the intake diffuser. In this configuration the expandable cone is in a retracted position and does not redirect airflow and allows the aircraft to operates in low speed mode as only the turbo jet receives airflow. In a transition speed mode, the air flow duct is retracted to allow air flow to both the turbo jet and the CVC.

Abstract: A system for cooling a generator mounted in the tail-cone of an engine. The system comprises a fairing, which receives through an inlet thereof air from a bypass duct and directs the bypass air towards a cavity of the tail-cone for cooling the generator. The bypass air is then expelled through an outlet of a support strut positioned in fluid communication with the tail-cone cavity. The fairing inlet and the strut outlet are both positioned in a plane substantially perpendicular to a longitudinal plane of the engine. In this manner, circulation of the bypass air through the fairing, the tail-cone cavity, and the strut may be achieved. The bypass air directed through the fairing further enables cooling of service lines accommodated in the fairing. A lobe mixer is further used to direct the fairing and shield the latter from core exhaust.

Abstract: A cooling system for a transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine is disclosed. The transition duct may have a multi-panel outer wall formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that at least one cooling chamber is formed between the inner and intermediate panels. The transition duct may also include an outer panel. The inner, intermediate and outer panels may include one or more metering holes for passing cooling fluids between cooling chambers for cooling the panels. The intermediate and outer panels may be secured with an attachment system coupling the panels to the inner panel such that the intermediate and outer panels may move in-plane.

Abstract: A turbine engine system includes a turbine engine shaft, a first rotor, a second rotor and an engine monitoring system. The shaft includes a shaft bore formed by a shaft wall. The shaft bore extends along an axial centerline through the shaft between a first shaft end and a second shaft end. The first rotor is connected to the shaft at the first shaft end, and the second rotor is connected to the shaft at the second shaft end. The engine monitoring system includes a sensor connected to the second rotor, a transmitter arranged at the first shaft end, and a plurality of conduit assemblies. Each conduit assembly includes a conduit that extends axially within the shaft bore and is connected to the shaft wall. A first of the conduit assemblies also includes a wire that extends through a respective conduit and connects the sensor and the transmitter.

Abstract: A gas turbine is provided having a secondary combustion chamber and a first guide vane row of a low-pressure turbine, the row being arranged directly downstream of the chamber. The radially outer boundary of the secondary combustion chamber is formed by at least one outer wall segment, which is secured on at least one support element arranged radially outwardly. The flow path of the hot gases is bounded radially outwardly, in the region of the guide vane row, by an outer platform which is secured at least indirectly on at least one guide vane support. A substantially radially extending gap-shaped cavity having a width in the range of 1-25 mm in the axial direction in the inlet region is arranged between the wall segment and the outer platform. At least one step element, which reduces the width by at least 10% in at least one step, extending substantially perpendicularly to the direction of flow of the hot gas in the cavity, is arranged in the inlet region.

Abstract: An aircraft engine thrust reverser cascade assembly includes a plurality of circumferentially spaced cascade segments, each cascade segment including a plurality of spaced vanes, including an aft-most vane, and rails defining a series of cells or air passages therebetween. The cascade assembly also includes an aft cascade ring removably attached to the aft ends of the cascade segments. A flow deflection shelf is mounted to each of the cascade segments and includes a deflector portion that at least partially extends forward of a group of cells of the cascade segments along which the flow deflection shelf is mounted. The deflector portion is configured to at least partially redirect at least a portion of a volume of air forwardly as the air outwardly passes through the group of cells of the cascade segments.

Abstract: A gas turbine includes a combustor chamber that houses a combustor unit configured to include a combustor that burns fuel to generate combustion gas for rotating a rotor, a turbine unit chamber that houses a turbine-unit rotor blade and a disk that rotate upon reception of the combustion gas, a combustor casing that forms the combustor chamber, and a casing that is configured to include the combustor casing in which a divided portion on a surface orthogonal to a rotation axis of the rotor is not formed in the combustor casing, but is formed in a portion on a downstream side of flow of the combustion gas lower than the combustor casing.

Abstract: A burner insert for a gas turbine combustion chamber is provided. The burner insert includes a burner insert wall including a cold side and a hot side, an edge delimiting the burner insert wall. The edge includes an edge bar extending at least partially circumferentially and projecting beyond the cold side. A burner opening for inserting a burner is formed in the burner insert wall.

Abstract: A fuel injection system comprises a combustor liner, an air tube, a fuel manifold and a fuel injector. The combustor liner defines primary and secondary combustion zones. The air tube extends through the combustor liner along an axis to an exit aperture inside the combustor. The fuel manifold is positioned proximate the combustor line to deliver fuel through the air tube to the primary and secondary combustion zones. The fuel injector is coupled to the fuel manifold and positioned to inject the fuel along an axis of the air tube, such that the fuel has a substantially symmetric distribution at the exit aperture.

Abstract: The invention relates to a method for controlling an electric motor actuating a mobile hood provided in a thrust reverser for a turboreactor, said method being characterized in that it comprises steps of: determining the operating state of the electric motor; interrupting the supply of the electric motor if said motor is not in operation during a defined period of time; reactivating the electric motor after an idle period and repeating the previous steps or definitively stopping the motor if the steps have already been repeated a pre-defined number of times. The invention also relates to a method for managing the electrical supply of a motor provided in a device arranged in the vicinity of a turboreactor.

Abstract: A method, system and apparatus for generating thrust. The method, system and apparatus can include a fuel that may be accelerated into one or more openings on a rotating wheel. The rotating wheel, which may be balanced, may then have a temporary increase in mass in one location on the wheel. Additionally, the fuel that is accelerated into one or more openings in the wheel may be decelerated and may exert an outward force on the wheel. This generation of force can be repeated and increased to provide thrust.

Abstract: An aircraft engine assembly including an attachment mechanism of a turbojet onto a rigid structure of a mounting strut, including a first, second, and third forward engine attachment for taking up thrust loads brought onto the fan frame, and arranged such that the third attachment passes through a first diametrical plane of the turbojet, the first and second attachments being disposed on one side and another of the first plane. The first and second forward engine attachments are respectively brought to the fan frame at two points, situated beyond the second diametrical plane of the turbojet, orthogonal to the first diametrical plane, with respect to the third attachment.