Abstract: A cooling manifold has a plurality of pieces. The pieces extend in a circumferential direction to abutting flanges. The flanges are secured together at circumferential ends of each piece. Cooling channels are formed in between inner and outer walls. Air inlets are formed in the pieces with the air inlets delivering air in the interior. There are fingers on an outer periphery. The fingers are aligned within an air outlet. The air can be delivered into the inlet, cool the interior, and leave through the outlet extending to a main conduit. The main conduit is secured directly to an outer periphery of the cooling manifold.

Abstract: A gas turbine engine includes an engine static structure housing that includes a compressor section and a turbine section. A combustor section is arranged axially between the compressor section and the turbine section. A core nacelle encloses the engine static structure to provide a core compartment. An oil tank is arranged in the core compartment and is axially aligned with the compressor section. A heat exchanger is secured to the oil tank and arranged in the core compartment.

Abstract: An environmental defense shield includes symmetric airfoil-shaped vanes contributing to and positioned around a plenum space and positioned in front of a turbine engine. An annular band stiffener is set into the vanes, which projects forward in a diminishing size, the vanes merging together to create or attaching to a solid nose. The environmental defense shield serves to protect the engine from debris while also smoothing airflow into the engine.

Abstract: A gas turbine comprises a compressor module, with a lower pressure compressor section including a plurality of stages, with at least one of the plurality of stages being an integrally bladed rotor. A higher pressure compressor section includes a plurality of stages with at least one of the plurality of stages being an integrally bladed rotor. A fan drive turbine shaft drives a fan rotor through a gear reduction. The fan rotor delivers a portion of air into a bypass duct, and a portion of air into the compressor module. A bypass ratio defined by the volume of air delivered into the bypass duct compared to the volume of air delivered into the compressor module is greater than or equal to about 6.0.

Abstract: In a perforated coaxial jet burner implemented by a lot of air-fuel coaxial jets, a swirl plate (33) as an end face of the burner on the combustion chamber's side has a lot of air holes for supplying unburned premixed gas of fuel and air to the combustion chamber. Grooves (36) are formed downstream of the air holes of the swirl plate. Adhesion of flame to the swirl plate is inhibited by feeding part of the unburned premixed gas to the grooves. Further, the width of each remaining part (37) between adjacent grooves is set at several millimeters that is approximately equal to the flame quenching distance, by which adhesion of flame to the remaining parts is also prevented. With this configuration, both stable combustion and low NOx combustion can be achieved irrespective of the load condition.

Abstract: A cooling system according to an embodiment includes: a pin fin bank cooling circuit; and an air feed cavity for supplying cooling air to the pin fin bank cooling circuit; wherein the pin fin bank cooling circuit extends radially outward from and at least partially covers at least one central plenum of a multi-wall blade and a first set of near wall cooling channels of the multi-wall blade.

Abstract: A center body assembly for a fuel injector is disclosed. The center body assembly defines a primary liquid passage, a liquid gallery, swirl slots, and a prefilm passage. The primary liquid passage supplies liquid fuel to the liquid gallery. The liquid gallery includes a gallery discharge end and a gallery inlet end that aligns with the primary liquid passage. The liquid gallery tapers down from the gallery inlet end to the gallery discharge end. The swirl slots extend from the liquid gallery to the prefilm passage in a helical pattern. The prefilm passage includes an annular shape that turns inward towards an assembly axis of the center body assembly.

Abstract: A combustion liner for use in a combustor assembly is provided. The combustion liner includes a side wall that defines a combustion chamber having a main flow axis extending therethrough. The combustion chamber channels a flow of combustion gas therethrough along a mixer swirl flow axis oriented obliquely relative to the main flow axis. The combustion liner also includes film cooling holes defined within the side wall. The film cooling holes are configured to discharge a pressurized fluid jet into the combustion chamber, and arranged in at least a first row and a second row positioned a first distance from the first row. The film cooling holes are arranged such that, when mixed with the flow of combustion gas, the pressurized fluid jet discharged from film cooling holes in the first row is directed along a discharge flow axis misaligned from film cooling holes in the second row.

Abstract: A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a speed different than the turbine section such that both the turbine section and the fan section can rotate at closer to optimal speeds providing increased performance attributes and performance by desirable combinations of the disclosed features of the various components of the described and disclosed gas turbine engine.

Abstract: Methods and systems for cooling airflow may comprise a gas turbine engine that may comprise an engine core, comprising a chamber, and/or a heat exchanger. The chamber may comprise a chamber outboard surface and/or a chamber interior. The heat exchanger may be coupled to the chamber outboard surface and may comprise a heat exchanger base and/or a cooling tube. The cooling tube may be disposed inside the chamber interior and may comprise a tube entrance end coupled to the heat exchanger base, a tube exit end coupled to the heat exchanger base, and/or a tube body between the tube entrance end and the tube exit end, the tube body encompassing a tube interior.

Abstract: There is proposed a method for controlling a bistable shutoff vale for an aircraft engine, involving the steps of: emitting a closure signal on one of the control channels—emitting an open signal on the other control channel, the method being characterized in that—the amplitude of the close signal is designed to cause the vane to move into a first position constantly in a given period of time,—in the same period, the amplitude of the open signal fluctuates between a value designed to urge the valve toward a second position and a rest value, the actuating values of the two signals being adapted to one another in order to keep the vane immobile when the two signals received by the valve simultaneously adopt said actuation values.

Abstract: A gas turbine combustor 2 according to the present invention comprises: a combustion chamber 50 in which fuel is burned with air to generate combustion gas; a plurality of fuel nozzles 30 arranged in multiple concentric annular rows; a first plate 32 arranged downstream of the fuel nozzles 30 and having multiple concentric circular air hole rows made up of a plurality of air holes corresponding to the fuel nozzles 30; a second plate 33 arranged downstream of the first plate 32 and having multiple air hole rows corresponding to the air hole rows of the first plate 32; and a partition wall part 37 which partitions a space part 46 between the first plate 32 and the second plate 33 into rooms corresponding to the air hole rows.

Abstract: An engine and particle decelerator is provided herein. The engine having: an inlet opening for directing air towards a compressor of the engine; and a particle decelerator located between the inlet opening and the compressor such that air travelling towards the compressor from the inlet opening must travel through the particle decelerator and wherein an area of the particle decelerator is greater than an inlet and an outlet of the particle decelerator.

Abstract: A fan blade for a turbomachinery fan and methods for fabricating a fan blade for a turbomachinery fan are disclosed. The fan blade for a turbomachinery fan includes an airfoil having a leading edge and a trailing edge in a chordwise direction, a tip and a root in a spanwise direction, a suction side and a pressure side. The fan blade includes a sheath including a solid portion that covers the leading edge of the airfoil, a first wing attached to the suction side of the airfoil, and a second wing attached to the pressure side of the airfoil. Construction of the fan blade includes one or more hollow cavities between the suction side and the pressure side of the airfoil.

Abstract: A gas turbine engine including a compressor section and a turbine section, coupled by one or more shafts, is provided. The compressor section progressively compresses air and includes an aft-stage of rotor blades rotatable about an axial direction of the gas turbine engine. A cooling system is included with the gas turbine engine for cooling compressed air in or from the compressor section. The cooling system includes a fluid tank for storing a volume of cooling fluid and one or more fluid lines in fluid communication with the fluid tank. The one or more fluid lines include an outlet positioned adjacent to the aft-stage of rotor blades for injecting cooling fluid into the compressed air proximate the aft-stage of rotor blades.

Abstract: A thrust reverser for an aircraft engine assembly, comprising two sub-assemblies configured to surround an engine, each of the sub-assemblies comprising an inner fixed structure, a translating cowl, a torque box, a substantially vertical 12 o'clock bifurcation, and a substantially vertical 6 o'clock bifurcation. At least one of the sub-assemblies comprises at least one separating partition extending between the inner fixed structure, the translating cowl and the torque box of the sub-assembly, the separating partition being rigidly fixed to the inner fixed structure and the torque box, and connected to the translating cowl via a sliding connection.

Abstract: Apparatus, methods and computer programs are provided. In one example, an apparatus is a rocket motor, comprising: a casing having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension; and propellant, located inside the casing, arranged to generate a force in a direction that is substantially perpendicular to the length dimension of the casing.

Abstract: A gas turbine engine component includes a wall that provides exterior and interior surfaces. The interior surface faces an internal cooling passage of the gas turbine engine component. An aperture extends through the wall and interconnects the interior and exterior surfaces to one another and is configured to provide a cooling fluid from the cooling passage to the exterior surface. The aperture has first and second outlet holes overlapping one another at an intersection to provide opposing sharp corners at the intersection.

Abstract: A fuel nozzle assembly includes a center body having a pilot air passage and a pilot fuel passage defined therein. A pilot nozzle having a plurality of premix passages is disposed within a downstream end portion of the center body. Each premix passage includes an inlet that is in fluid communication with the pilot air passage, an outlet that is positioned axially downstream from the inlet and a fuel port that is in fluid communication with the pilot fuel passage. An outer sleeve is coaxially aligned with and radially spaced from the center body so as to define an annular passage therebetween. A strut extends radially outwardly from the center body to the outer sleeve. The fuel nozzle assembly further includes an inlet passage that is in fluid communication with the pilot air passage. The inlet passage extends through the outer sleeve, the strut and the center body.

Abstract: An atomizer of the invention includes a mixing chamber, spray fluid flow passages supplying a spray fluid to the mixing chamber, a spray medium flow passage supplying a spray medium to the mixing chamber, outlet holes spraying a fluid mixture of the spray fluid and the spray medium, and fluid mixture flow passages connecting the mixing chamber and the outlet holes to each other. The fluid mixture flowing through each of the fluid mixture flow passages joins with each other at a joining section and is sprayed from the outlet hole. The mixing chamber includes a middle ejection hole spraying the spray fluid to the mixing chamber. The middle ejection hole is arranged at a joining section of the spray fluid flow passages. The spray fluid flowing through the spray fluid flow passages joins with each other at the joining section and is sprayed from the middle ejection hole.