Abstract: A gas turbine engine component includes a first surface and a second surface. The component further includes a dilution hole defined by the first surface and the second surface. The component further includes a first effusion hole and a second effusion hole each having an inlet defined by the second surface and an outlet defined by the first surface such that the outlet of the first effusion hole is located nearer to the dilution hole than the outlet of the second effusion hole.

Abstract: A combustor includes: a casing having an air chamber filled with air; at least one mixing passage forming member forming at least one mixing passage which is connected at an inlet side to the air chamber and at an outlet side to a combustion chamber and having an inlet at the inlet side of the at least one mixing passage so as to communicate with the air chamber; and at least one fuel nozzle disposed inside the air chamber and having a fuel injection hole, positioned upstream of the inlet of the at least one mixing passage forming member, for injecting fuel downstream.

Abstract: A fuel injector for a gas turbine engine, the injector comprising a main fuel flow circuit and a pilot fuel flow circuit. The pilot fuel flow circuit comprises at least one convoluted pilot fuel flow passage extending continuously from a pilot fuel feed at an upstream end of the injector to a pilot fuel spray nozzle inlet. The at least one pilot fuel flow passage comprises at least two axially-extending portions extending between the upstream end of the injector towards a downstream end of the injector and joined by an elbow portion proximal the downstream end of the injector. In use, pilot fuel can flow in opposing axial directions within the at least two axially-extending portions, the axially-extending portions being in thermal communication with the main fuel flow circuit.

Abstract: A gas turbine engine article includes an article wall that defines leading and trailing ends and first and second sides that join the leading and trailing ends. The article wall defines a cavity. A cooling passage network is embedded in the article wall between inner and outer portions of the article wall. The cooling passage network has an inlet orifice through the inner portion of the article wall to receive cooling air from the cavity, a plurality of sub-passages that extend axially from the at least one inlet orifice, at least one outlet orifice through the outer portion of the airfoil wall, and trip strips for mixing cooling air in the cooling passage network.

Abstract: Sealing arrangements and turbomachines are provided. A sealing arrangement for a turbomachine includes a transition duct having an upstream end and a downstream end. The transition duct includes an aft frame that circumferentially surrounds the downstream end of the transition duct. A stage one nozzle spaced apart from the aft frame and defining a gap therebetween. A sealing assembly extends across the gap. The sealing assembly includes a first seal link magnetically coupled to the aft frame. The sealing assembly further includes a second seal link magnetically coupled to the first seal link and the stage one nozzle.

Abstract: A combustor has an axial fuel staging system to allow a fuel-air mixture to be injected from two axially spaced stages using an injector for injection of a secondary fuel-air mixture. The combustor includes a liner defining a combustion chamber; a transition piece coupled to a rear end of the liner; a flow sleeve defining an annular channel by surrounding the liner and the transition piece; and at least one injector disposed on a circumferential position of the flow sleeve to inject a mixture of fuel and air into the combustion chamber. Each of the at least one injector includes an injection pipe extending radially and passing through both the flow sleeve and either of the liner and the transition piece; a plate coupled to the injection pipe; and a plurality of mixing passages formed through the plate. The combustor improves fuel-air mixing and prevents flash back.

Abstract: An aircraft includes an environmental control system having a bleed air inlet, a gas turbine engine having a low pressure bleed air supply and a high pressure bleed air supply, and a turbo air cycle machine. The turbo air cycle machine includes a compressor fluidly coupled to the low pressure bleed air supply, as well as a turbine fluidly coupled to at least one of the low pressure bleed air supply or the high pressure bleed air supply.

Abstract: Methods and apparatus to operate a gas turbine engine with hydrogen gas are disclosed. An example combustor nozzle apparatus of a gas turbine engine includes a first circuit to transport a blend of hydrogen gas, inert gas, and/or other combustible gas from a supply to a gas turbine combustor, the blend of hydrogen gas, inert gas, and/or other combustible gas including between 100% hydrogen gas, 100% inert gas, or 100% other combustible gas, a second circuit to transport water from the supply to the gas turbine combustor, and a nozzle tip. The nozzle tip includes a first outlet in connection with the second circuit to provide the water to the gas turbine combustor, and a second outlet in connection with the first circuit, the second outlet concentrically positioned within the first outlet to provide the blend of hydrogen gas, inert gas, and/or other combustible gas to the gas turbine combustor.

Abstract: A gas turbine engine comprising: an engine core comprising a compressor; a compressor bleed valve in communication with the compressor and configured to release bleed air from the compressor; at least one component provided at the inlet of the engine core having a de-icing conduit, configured to receive the bleed air; and a flow controller, configured to provide bleed air to the de-icing conduit of the at least one component in response to either or both of a requirement to de-ice the component and a requirement to release bleed air from the compressor to optimise operation of the core.

Abstract: The present disclosure provides a thrust reverser comprising a stationary structure defining an annular body with a centerline, a first reverser door pivotally coupled to the stationary structure by a pair of first reverser door hinges, a first reverser door hinge axis extending through the first reverser door hinges and positioned at a first angle relative to a centerline, and a second reverser door pivotally coupled to the stationary structure by a pair of second reverser door hinges, a second hinge line axis extending through the second reverser door hinges and positioned at a second angle relative to the centerline.

Abstract: A power plant with a multi-stage intercooled compressor, a combustion chamber, a turbine which is arranged downstream of the combustion chamber, a compressor air line which connects the compressor to the combustion chamber, and a first heat exchanger which is connected into the compressor air line and into an exhaust gas line branching off from the turbine. A first compressor air expander is arranged in the compressor air line between the first heat exchanger and the combustion chamber, and the power plant includes a device for regasifying liquid natural gas, having a natural gas line, wherein a heat exchanger device is connected into the natural gas line between two compressor stages of the compressor.

Abstract: An injection system for a turbine engine annular combustion chamber includes a support configured to support and to center a fuel injector head. The support includes a frustoconical surface connected at its downstream end of smallest diameter to an upstream end of a cylindrical surface. The system further includes a bowl configured to mix air and fuel arranged downstream of the support and at least one axial swirl inducer extending at least in part around the support. Each swirl inducer includes vanes delimiting between them substantially axial channels for the passage of an air flow. The channels open at their upstream ends on said frustoconical surface.

Abstract: An airfoil assembly includes an airfoil that has an exterior wall that defines an interior cavity. The exterior wall extends between a leading end and a trailing end and an open inboard end and an open outboard end. The exterior wall is formed of a high temperature-resistant material selected from refractory metal-based alloys, ceramic-based material or combinations thereof. A support frame extends in the interior cavity and protrudes from the interior cavity through at least one of the open inboard end and the open outboard end.

Abstract: An aircraft can include a turbine engine including a rotor and a stator, and also including a compressor, a combustor, and a turbine in axial flow arrangement. The aircraft can further include an anti-icing system with a magnetic field generator having a rotating portion and a non-rotating portion, as well as an array of carbon nanotubes thermally coupled to an exposed surface.

Abstract: A flight vehicle has an engine that includes air inlet, an isolator (or diffuser) downstream of the air inlet, and a combustor downstream of the isolator. The isolator includes a bulged region that has at least one dimension, perpendicular to the direction of the air flow from the inlet to the combustor, that is at a local maximum, larger than comparable isolator dimensions both upstream and downstream of the bulged region. The bulged region stabilizes shocks within the isolator, and facilitates flow mixing. The flow diversion of high energy flow around the outermost walls of the bulged section into the center of the flow at the aft end of the isolator, increases mixing of the flow, and results in a more consistent flow profile entering the combustor over a wide range of flight conditions (Mach, altitude, angle-of-attack, yaw) and throttle settings.

Abstract: A system includes an internal manifold ring defining at least one circumferentially extending flow channel. A plurality of feed arms extends outward from the manifold ring. A circumferentially segmented outer ring can be supported from the feed arms, outboard of the feed arms. Each feed arm can include a plurality of branches extending therefrom, wherein each branch is in fluid communication with the manifold ring through a respective one of the feed arms. A plurality of nozzles can be included, each nozzle connected to a respective one of the branches, wherein the system is devoid of nozzles radially inward from the manifold ring.

Abstract: A gas turbo machine having a combustion chamber. The combustion chamber includes a chamber bottom having an upstream wall and a downstream wall extending between inner and outer annular walls. Each upstream and downstream walls each having openings for mounting fuel injection devices for injecting fuel through said openings. The downstream wall is sectorised.

Abstract: The jet engine comprising a ramjet air path extending from an intake, into a combustion chamber, and out an exhaust nozzle, a fuel inlet leading into the combustion chamber, an oxidizer inlet leading into the combustion chamber and a partition being operable to selectively close the ramjet air path upstream of the combustion chamber to allow operation of the jet engine in rocket mode and open the ramjet air path to allow operation of the jet engine in ramjet mode.

Abstract: One aspect the present subject matter is directed to a nozzle segment including a stator component having an airfoil. The airfoil includes a leading edge portion, a trailing edge portion, a pressure side wall and a suction side wall and a plurality of film holes in fluid communication with the radial cooling channel. A strut is disposed within the radial cooling channel and defines an inner radial cooling passage within the radial cooling channel. The strut defines a plurality of apertures that provide for fluid communication from the inner radial cooling passage to the radial cooling channel and the plurality of film holes provide for bore cooling of the airfoil of at least one of the pressure side wall or the suction side wall and provide for film cooling of the trailing edge portion of the airfoil between about fifty percent and one hundred percent of the chord length.

Abstract: A hydraulic and pneumatic control circuit for a turbojet, a main hydraulic line having an oil/fuel heat exchanger with a function of transferring heat from the oil flowing in an oil circuit of the turbojet to the fuel flowing in the main hydraulic line, the circuit having a first hydraulic line for feeding fuel to a combustion chamber of the turbojet, a second hydraulic line for feeding fuel to one or more actuators for controlling variable geometry equipment, each actuator being fed with fuel via an electrohydraulic servovalve, a pneumatic line for feeding air to a pneumatic control member for bleed valves of a compressor and a blade tip clearance control valve of a turbine of the turbojet, and a fuel/air heat exchanger positioned on the second hydraulic line upstream from the hydraulic servovalve and on the pneumatic line upstream from the pneumatic control member.