Abstract: A diffuser assembly is provided for a turbine engine. This diffuser assembly includes a diffuser module with a combustor plenum and a mixing chamber. The diffuser module is configured to receive first and second airflows into the mixing chamber and direct a mixed airflow out of the mixing chamber. The diffuser module includes a mixer configured to mix the first and the second airflows together within the mixing chamber to provide the mixed airflow.

Abstract: A combustor is provided in which a fuel and working fluid can be injected in an annulus. In one form the fuel and working fluid is circumferentially flowed within the annulus and traverses the annulus in an axial direction from one side to another side where the flow exits. The working fluid and air can be co-axially admitted to the combustor and in one form the working fluid can be swirled about the fuel dispensed from a fuel injector. The combustor can provide for a rich burning zone. In one embodiment the combustor is configured as an inter-turbine combustor having an outlet at one axial side of the combustor. A lean burn region can be created within a flow path of the turbine.

Abstract: A method of tailoring a combustor flow for a gas turbine engine includes controlling an airflow into a swirler to be generally uniform and controlling an airflow into a quench zone to provide a desired pattern factor.

Abstract: In order to stably use a catalyst for pyrolysis and supply a reformed fuel, the fuel supply system includes a fuel reforming section which pyrolyzes a hydrocarbon system fuel by the heat of the combustion chamber to generate the reformed fuel. The fuel reforming section includes a preheat vaporization section provided on the combustion chamber, and a decomposition reaction section that is provided on the preheat vaporization section and includes the catalyst for pyrolysis. The preheat vaporization section heats the fuel, the decomposition reaction section pyrolyzes the heated fuel to generate the reformed fuel, and the fuel reforming section supplies the reformed fuel to the combustion chamber. The reforming catalyst includes a zeolitic catalyst.

Abstract: A sealing arrangement has a turbine static structure with contact surfaces, a bearing compartment with contact surfaces, and a cavity between the turbine static structure and the bearing compartment. There are also two seals, wherein each seal is configured to contact the turbine static structure and the bearing compartment.

Abstract: A method for controlling fuel flow to a gas turbine engine during starting includes monitoring acceleration of the gas turbine engine to determine actual acceleration value, and calculating a fuel flow rate for a setpoint acceleration using the actual acceleration value as a factor. The method further includes commanding the calculated fuel flow for the setpoint acceleration to the gas turbine engine.

Abstract: A multi-lobe exhaust mixer has an annular body composed of a plurality of circumferentially adjacent lobe segments. The lobe segments may be made of a ceramic matrix composite material to reduce the weight of the mixer and ensure proper behavior when exposed to high thermal gradients. Each lobe segment may have partial lobes at circumferentially opposed ends thereof and at least one complete lobe therebetween. The partial lobes of the circumferentially adjacent lobe segments combining to conjointly form complete lobes at the junction between the circumferentially adjacent lobe segments. The partial lobes may be nested into each other to dampen vibrations.

Abstract: Example folding door thrust reversers for aircraft engines are disclosed herein. An example apparatus includes a nacelle of a turbofan engine, where a fan duct is defined between the nacelle and a core of the turbofan engine. The example apparatus includes an opening in the nacelle between an outside of the nacelle and the fan duct and an inner door and an outer door disposed within the opening and pivotably coupled to the nacelle along aft edges thereof. The example inner and outer doors are pivotable between a first position in which the inner door and the outer door are disposed within the opening and oriented substantially parallel to each other, and a second position in which the inner door is disposed in the fan duct and oriented substantially perpendicular to an outer surface of the core and the outer door extends outward from the nacelle.

Abstract: A turbine exhaust case (TEC) of a turbofan aeroengine includes a mixer for mixing exhaust gases with a bypass air stream, the TEC comprising an annular hub and the mixer surrounding the hub, and a plurality of deswirling struts circumferentially spaced apart with respect to a central axis of the TEC and located entirely within an axial length of the mixer. The mixer defines a trailing edge having one or more upstream-most locations thereof where the mixing of the exhausted gases and the bypass air stream begins to take place. The deswirling struts each extend radially across the annular exhaust gas duct and interconnect the mixer and the hub, defining a trailing edge positioned upstream of and axially spaced away from the one or more upstream-most locations of the trailing edge of the mixer.

Abstract: The object of the invention is to suppress development of a secondary current whirl in the vicinity of a rotor blade front edge of a gas turbine even when a cooling refrigerant is mixed in from the front edge of an end wall section of the rotor blade.

Abstract: The present invention relates to a support structure (16) for a gas turbine engine (1). The support structure (16) has an axial extension in an axial direction (A) and a circumferential extension in a circumferential direction (C). Moreover, the support structure (16) comprises a plurality of tubular members (18, 20) of a first material type arranged in sequence in the circumferential direction (C). Each tubular member (18, 20) at least partially delimits a flow guiding passage extending at least partially in the axial direction (A). The support structure (16) comprises a leading portion (22) and a trailing portion (24) in the axial direction (A). Furthermore, the support structure (16) comprises a leading edge member (26) of a second material type, the leading edge member (26) being located at the leading portion (22). At least two of the tubular members (18, 20) are fixedly attached to a leading edge member (26).

Abstract: A gas turbine engine has a core engine and a nacelle having a cowl translatable along the axis of the engine from a stowed position to a deployed position, A thrust reverser cascade extends circumferentially about the engine axis. To obviate the need for additional blanking cascades the cowl has a forwardly extending tongue that in the deployed position of the cowl is at the same axial position as the thrust reverser cascade.

Abstract: A gas turbine combustor floating collar for mounting an igniter or fuel nozzle to a combustor is provided with an outer periphery in a non-circular shape, for example having at least one section thereof formed with a flat surface, such as a square or triangular shape. The outer periphery of the floating collar is complementary to and completely surrounded by an inner periphery surface of a recess of a boss affixed on the combustor, thereby preventing substantial rotation of the floating collar with respect to the boss.

Abstract: An electrical raft assembly for a gas turbine engine is provided. The raft assembly comprises a rigid electrical raft formed of a rigid material that includes an electrical system comprising electrical conductors embedded in the rigid material. The raft assembly further comprises an engine component that is mounted to the electrical raft. The electrical raft includes one or more integral cooling passages which guide a coolant fluid through the raft to cool the engine component.

Abstract: A turbine combustor section has a flow conditioner including a plurality of conduits arranged to convey pressurized air to an air chamber for entrance into a plurality of fuel nozzles. Each conduit includes an inlet configured to receive the pressurized air from an annular passage and an outlet configured to deliver the pressurized air to the air chamber. A cross-sectional area of each conduit varies between the inlet and the outlet so as to reduce a pressure drop across the flow conditioner.

Abstract: A support structure in a gas turbine combustor end cap (24) including a bracket (60) with a first leg (61) and a second leg (62) forming a generally trapezoidal geometry. Each leg has a first end (61A, 62A) attached to an inner concentric ring (46), and a second end (61B, 62B) attached to a crossbar (65). The crossbar is attached to an outer concentric ring (48). A circular array of such brackets interconnects the two concentric rings (46, 48). Each leg has at least one curved middle portion (63, 64), such as an arcuate or sinusoidal curve at a midpoint on the length of each leg. This shape provides flexibility in a radial direction that accommodates differential thermal expansion of the concentric rings while providing a rigid connection in an axial direction.

Abstract: The invention relates to a mixing device and a turbofan engine having such a mixing device 30 for mixing a first gas flow 40 with a second gas flow 50 in a turbofan engine 20, having an actuating device 95 and walls 60, which bound a channel 65 for the first gas flow 40 and a channel 70 lying radially outside for the second gas flow 50, the actuating device 95 comprising a coupling element 110 that is coupled to the walls (60), the actuating device 95 being designed to pivot the walls 60 between a first position and a second position disposed radially outside relative to the first position, the actuating device 95 comprising an adjusting ring 105 that can be rotated between a first rotating position and a second rotating position in the peripheral direction and that is joined to the coupling element 110.

Abstract: A heat exchange system includes a tubular fan air inlet portion and a tubular cooled air outlet portion connected to a first end of a tubular mid portion. The heat exchange system further includes a tubular hot air inlet portion and a tubular recycled fan air outlet portion connected a second end of the mid portion. Still further, the heat exchange system includes an integrally-formed, compliant heat exchanger tube extending between the hot air inlet portion and the cooled air outlet portion within the mid portion to define a heat exchanger first flow passage within the heat exchanger tube and a second flow passage outside of the heat exchanger tube but within the tubular mid portion. Methods for fabricating such heat exchange systems are also provided.

Abstract: A gas turbine engine assembly is connected to a pylon for mounting the gas turbine engine to an aircraft. The assembly has a frame supporting at least one accessory independently of the gas turbine engine. Frame is attached to the pylon at forward and rearward engine mounting locations. The frame includes at least one hollow tube and at least one hollow tube is fluid tight. The one hollow tube is evacuated or contains pressurised fluid and a pressure sensor is provided to detect a change in pressure in the at least one hollow tube to determine if there is a leak in the at least one hollow tube and hence if the frame is damaged. This ensures that the frame may be repaired or replaced before there is a loss of operation of one or more of the accessories which may result in a failure of the gas turbine engine.

Abstract: A precooler for an aircraft engine system includes a precooler core and a precooler inlet to direct a compressor bleed flow into the precooler core to cool the compressor bleed flow. The precooler further includes a precooler outlet to direct the compressor bleed flow from the precooler to a selected component of the aircraft engine system and a precooler bleed port through which a portion of the compressor bleed flow is diverted to a secondary component of the aircraft engine system. The precooler bleed port is oriented such that flow entering the precooler bleed port must substantially reverse direction from a direction of the compressor bleed flow through the precooler.