Abstract: A method for cooling a rotatable nozzle includes rotating a curved seal about a seal land while maintaining contact therewith. Cooling air is directed through a first diffusion hole in the curved seal to cool the nozzle if the rotatable curved seal is in a first position where higher heat is encountered. Cool air is directed through a second diffusion hole in the curved seal to cool the nozzle if the rotatable curved seal is in a first position where higher heat is encountered and if in a second position where relatively lower heat is encountered.

Abstract: A fuel plenum for a fuel nozzle assembly includes a gaseous fuel conduit, a conduit passage, and a liquid fuel conduit. Said gaseous fuel conduit received at a first end of said fuel plenum. Said fuel plenum is configured to distribute gaseous fuel received from said gaseous fuel conduit. Said conduit passage extends from the first end to a second end of said fuel plenum. Said conduit passage is at least partially defined by at least one interior wall of said fuel plenum. Said liquid fuel conduit defined by an outer wall and a portion of said liquid fuel conduit extending through said conduit passage. Said liquid fuel conduit outer wall is offset from said at least one interior wall.

Abstract: A cooling arrangement (100), including: a substrate having a pocket (24) defined by a rib side surface (28) and a bottom surface (26); a feature (112) formed in the rib side surface; and an impingement plate (102). In an installed configuration (116) a resilience of the impingement plate enables an interference between the impingement plate and the feature that locks the impingement plate in an installed position (118). An elastic compression of the impingement plate from the installed configuration eliminates the interference.

Abstract: An anti-ice valve, a turbine engine including an anti-ice valve assembly and a method of operating an anti-ice valve where an anti-ice valve has a housing and a valve element configured to control a flow of hot bleed air through the housing, and a muscle air passage extending through the housing, a cooling air passage extending through the housing, and a heat exchanger located within the housing and having heat transfer surfaces in thermal communication with the muscle air passage and the cooling air passage.

Abstract: A gas turbine engine including a secondary air system with interconnected fluid passages defining at least one flow path between a common source of pressurized air and a common outlet. Some of the fluid passages deliver the pressurized air to components of the gas turbine engine. The fluid passages include a common fluid passage through which all circulation of the pressurized air to the common outlet passes. The common fluid passage has a section including a venturi configured for controlling a flow of the pressurized air from the source to the outlet. In one embodiment, the venturi is provided in a common inlet or common outlet passage. A method of pressurizing a secondary air system is also discussed.

Abstract: A flexure for a metering valve has a flexure body defining a center axis to be aligned with a fuel nozzle. The flexure body includes an outermost peripheral surface surrounding the center axis. A retaining feature is formed in the flexure body and is configured to mount the flexure body to a metering valve component. The retaining feature is located radially between the center axis and the outmost peripheral surface. A metering valve is also disclosed.

Abstract: Bleed air systems for use with aircraft and related methods are disclosed. An example apparatus includes a compressor having a compressor inlet and a compressor outlet. The compressor is to be driven by a drive shaft extending from an engine of an aircraft. The example apparatus also includes a first passageway to fluidly couple a first low-pressure bleed air port from the engine to the compressor inlet and a second passageway to fluidly couple the compressor outlet to a system of the aircraft.

Abstract: The gas turbine system (GT) includes a combustor (2) having a fuel injection nozzle assembly (4) for jetting hydrogen gas (H) and pure water (W), a reservoir (12) for pooling the pure water (W) to be supplied to the combustor (2), a gas compressing device (10) for boosting the hydrogen gas (H) to be supplied to the combustor (2), a fuel supply passage (6) for guiding the boosted hydrogen gas (H) towards the combustor (2), and a pressurizing passage (16) communicating between the reservoir (12) and the fuel supply passage (6) for pressurizing the pure water (W) by means of the boosted hydrogen gas (H).

Abstract: A turbine engine is provided that includes a turbo-compressor and a combustor section. The turbo-compressor includes a compressor section and a turbine section. The combustor section is fluidly coupled between the compressor section and the turbine section. The compressor section includes a first number of stages. The turbine section includes a second number of stages that is different than the first number.

Abstract: A turbine engine includes a main shaft bearing compartment seal. The seal includes at least an approximately circular seal portion and a seal carrier disposed about the approximately circular seal portion. A plurality of anti-rotation pins maintain the seal carrier in position relative to a housing and are received in an anti-rotation slot of the seal carrier.

Abstract: A ram air turbine is provided that utilizes a one-piece strut. The strut includes an integral gearbox section and an integral drive section. Within the strut, a turbine shaft and a bevel gear engages a driveshaft and a pinion gear, which transfers rotation from the turbine to a generator. The strut may be machined from a single piece of metal, such as aluminum.

Abstract: A bypass housing receives a fan and defines a front end. An airflow path delivers air into an inlet duct over a limited circumferential extent of the bypass housing. An airflow path passes across a low pressure compressor rotor. An airflow path passes through a core engine, which includes a high pressure compressor rotor, a combustor, and a high pressure turbine rotor. Products of combustion downstream of the high pressure turbine rotor pass into an intermediate duct and then across a low pressure turbine rotor. The low pressure turbine rotor is positioned closer to the front end of the engine than is the high pressure turbine rotor. The low pressure turbine rotor is positioned axially intermediate the low pressure compressor rotor and the fan. The low pressure turbine rotor drives both the fan and the low pressure turbine rotor. An aircraft is also disclosed.

Abstract: A mid-turbine frame for a gas turbine engine ducts gases between a high pressure turbine and a low pressure turbine. The mid-turbine frame may include an outer flowpath ring, an inner flowpath ring, and a plurality of vanes extending therebetween. The outer flowpath ring comprises a unitary structure, while the inner flowpath ring and the plurality of vanes comprises a plurality of segments.

Abstract: An exhaust nozzle for a turbine engine includes multiple daisy style corrugations arranged circumferentially about the exhaust nozzle. Each of the daisy style corrugations has a radially inner base portion and multiple lobes protruding radially outward from the base portion.

Abstract: The present application provides a fuel nozzle for a gas turbine engine using a primary fuel and a secondary fuel. The fuel nozzle may include a number of primary fuel injection ports for the primary fuel, a water passage, a number of secondary fuel injection ports, and a secondary fuel evaporator system for atomizing the secondary fuel.

Abstract: A nozzle liner for a rotatable nozzle includes a seal land and a rotatable seal for moving with the nozzle. The seal has a first diffusion hole for distributing cooling air if the rotatable seal is in a first position and a second diffusion hole for distributing cooling air if the rotatable seal is in a first position and if in a second position.

Abstract: The invention relates to a damper for reducing pulsations in a gas turbine, which includes an enclosure, a main neck extending from the enclosure, a spacer plate disposed in the enclosure to separate the enclosure into a first cavity and a second cavity and an inner neck with a first end and a second end, extending through the spacer plate to interconnect the first cavity and the second cavity. The first end of the inner neck remains in the first cavity and the second end remains in the second cavity. A flow deflecting member is disposed proximate the second end of the inner neck to deflect a flow passing through the inner neck. With the solution of the present invention, as a damper according to embodiments of the present invention operates, flow field hence damping characteristic in the second cavity constant regardless the adjustment of the spacer plate in the enclosure.

Abstract: The invention relates to a method for starting an aircraft turboshaft engine, said turboshaft engine comprising a combustion chamber, a compressor shaft on which a compressor wheel is mounted to feed compressed air to said combustion chamber, at least one starter connected to said shaft so as to provide it with a specified starting torque for driving it in rotation. The method comprises accelerating the compressor shaft during a first start-up phase, then stabilizing the rotational speed of the compressor shaft during a second start-up phase. During acceleration of the compressor shaft, the rotational speed of the shaft is regulated such that the acceleration of the shaft remains substantially constant.

Abstract: A single actuator variable area fan nozzle (“VAFN”) system is provided. In contrast to typical system, the single actuator VAFN may comprise a single actuator, a linkage and a VAFN panel. The single actuator may be configured to actuate the VAFN panel between a stowed position and a deployed position. In various embodiments, the linkage may travel a path during deployment and/or retraction that is between the loft lines of the nacelle.

Abstract: An exhaust nozzle for a gas turbine engine may include a plurality of flap trains in the exhaust stream of the gas turbine engine. The flap trains are operable to selectively control three separate flow paths of gas that traverse the engine. A first stream of is the core airflow. The second stream of air is peeled off of the first stream to form a low pressure fan bypass air stream. The third stream of air traverses along the engine casing and is passed over a flap assembly to aid in cooling. The flaps are operable converge/diverge to control the multiple streams of air.