Abstract: A conjoined grommet assembly for a combustor wall assembly of a gas turbine engine has a first grommet defining at least in-part a first dilution hole, and a second grommet defining at least in-part a second dilution hole. The first and second dilution holes are spaced closely together such that the first grommet is in contact with the second grommet.

Abstract: A blocker door assembly which may be for a cooling system that may be applied to a gas turbine engine includes a plurality of blocker doors circumferentially spaced about an engine axis. Each blocker door is constructed and arranged to move in a circumferential direction to, at least in-part, control air flow through a passage in an adjacent fixture. A sync-ring is concentrically located about the engine axis, disposed in an annular first duct in direct communication with each passage, and engaged to each one of the plurality of blocker doors for simultaneous operation. The sync-ring is aero-dynamically shaped to reduce surrounding airflow resistance.

Abstract: The present invention discloses a novel apparatus and methods for augmenting the power of a gas turbine engine, improving gas turbine engine operation, and reducing the response time necessary to meet changing demands of a power plant. Improvements in power augmentation and engine operation include additional heated compressed air injection, steam injection, water recovery, exhaust tempering, fuel heating, and stored heated air injection.

Abstract: For gains in terms of aerodynamic performance levels, an aircraft engine assembly includes a turbomachine, a fastening strut for the turbomachine, and at least one foldable ventilator cover which surrounds the turbomachine and which includes: a first cover sector which includes a first end portion which is mounted so as to be articulated to the fastening strut, along a first articulation axis, and a second cover sector which includes a first end portion which is mounted so as to be articulated to a second end portion of the first cover sector, along a second articulation axis parallel with the first articulation axis. The second end portion is mounted so as to be guided at one side and the other thereof by a thrust inverter cover of the engine assembly and an air inlet structure of this assembly, respectively.

Abstract: A variable area fan nozzle assembly for a turbofan engine includes a nacelle having an aft edge and a translating thrust reverser sleeve with a trailing edge. The thrust reverser sleeve is movably disposed aft of the nacelle's aft edge and is movable between a forward position and an aft position. A translating fan nozzle having a forward edge is movably disposed behind the trailing edge, and is movable between a stowed position and a deployed position. An upstream bypass flow exit is defined between the trailing edge and the forward edge when the fan nozzle is in the deployed position. An extendable actuation system is configured to move the fan nozzle between the stowed position and the deployed position.

Abstract: A mid-turbine frame is provided. The mid-turbine frame may comprise an inner case having an annular surface with an interface section disposed on the annular surface. The interface section may include an interface feature. A balancing section may be disposed on the annular surface defining an opening and disposed circumferentially adjacent the interface section.

Abstract: The present invention discloses a novel apparatus and methods for controlling an air injection system for augmenting the power of a gas turbine engine, improving gas turbine engine operation, and reducing the response time necessary to meet changing demands of a power plant. Improvements in control of the air injection system include ways directed towards preheating the air injection system, including using an gas turbine components, such as an inlet bleed heat system to aid in the preheating process.

Abstract: In various embodiments, a lubricating shaft assembly may comprise a shaft, a front seal, a front seat, an inner race, a spacer, a lubricating fitting and a nut. The front seat may define a first portion of a fluid conduit. The front seat may be part of a first load path. The inner race may define a second portion of the fluid conduit. The inner race may be installed about the shaft. The inner race may be part of the first load path. The spacer may define an internal diameter of a third portion of the fluid conduit. The lubricating fitting may be installed about at least a portion of the spacer. The lubricating fitting may define an outer diameter of the third portion of the fluid conduit. The lubricating fitting may be outside the first load path.

Abstract: A gas turbine engine combustor configured to cool a first vane platform by directing airflow through an aft combustor rail is disclosed. In various embodiments, a gas turbine engine combustor may comprise an aft combustor rail connected between an outer wall of an outer liner and an aft heat shield panel, and a plurality of holes in the aft combustor rail. The aft combustor rail may be positioned near an outer vane platform of a first vane stage of a turbine. Further, the plurality of holes may be designed to focus pass-through airflow onto the outer vane platform. In various embodiments, each hole of the plurality of holes may be angled between 20°-90° relative to the aft combustor rail. Additionally, the plurality of holes may be angled to converge the airflow to a focused area, such as a leading edge vane airfoil of the first vane stage.

Abstract: An airflow control system for a gas turbine system according to an embodiment includes: a compressor component of a gas turbine system for generating an excess flow of air; a mixing area for receiving an exhaust gas stream produced by the gas turbine system; an air extraction system for extracting at least a portion of the excess flow of air generated by the compressor component of the gas turbine system to provide bypass air; and diverting the bypass air into the mixing area to reduce a temperature of the exhaust gas stream; and an exhaust processing system for processing the reduced temperature exhaust gas stream.

Abstract: An airflow control system for a gas turbine according to an embodiment includes: an airflow generation system for attachment to a rotatable shaft of a gas turbine system, the airflow generation system drawing in an excess flow of air through an air intake section; a mixing area for receiving an exhaust gas stream of the gas turbine system; an air extraction system for: extracting at least a portion of the excess flow of air generated by the airflow generation system to provide bypass air; and diverting the bypass air into the mixing area to reduce a temperature of the exhaust gas stream; and an exhaust processing system for processing the reduced temperature exhaust gas stream.

Abstract: A nacelle may include a pylon and a thrust reverser having an inner fixed structure. A locking mechanism may automatically engage and prevent relative movement between thrust reverser halves. A bumper may be coupled to the inner fixed structure. The locking mechanism may limit deflections between the thrust reverser and the pylon in response to a burst duct. A locking mechanism in the inner fixed structure may include a pressure relief door and an arrestor which hooks onto a retaining bar in a pylon bracket. The locking mechanism may allow the thrust reverser halves to be opened for access to the engine.

Abstract: The application relates to a method and a device for suppressing ice formation on intake structures of a compressor, particularly the compressor of a gas turbine. The technical aim of the present invention is to provide a method and a device for suppressing the formation of ice on said structures, which avoid the disadvantages of known solutions, such as a reduction of the performance of the gas turbine, and have a simple and broad applicability. According to the present invention the mechanical vibratory energy of said structures during operation is converted into electrical energy by a piezoelectric element, firmly applied to said structure, and in a connected electrical circuit the generated electrical energy is then converted into thermal energy by an ohmic resistor and this thermal energy is conducted to at least a portion of the structure for suppressing ice formation. Excess energy may be transmitted by a transmitter to other circuits in adjacent structures.

Abstract: A turbine engine assembly having a turbine core with a compressor section, a combustion section, a turbine section, and a nozzle section and a liquid natural gas (LNG) fuel system having a LNG reservoir, a vaporizer heat exchanger, a first liquid supply line operably coupling the LNG reservoir to an input of the vaporizer heat exchanger, a gas supply line operably coupling an output of the vaporizer heat exchanger to the combustion section, a second liquid supply line operably coupling the LNG reservoir to the gas supply line and a thermostatic expansion valve (TEV) and a dual fuel aircraft control system.

Abstract: This invention relates to electrical power systems, including generating capacity of a gas turbine, and more specifically to augmentation of power output of gas turbine systems, that is useful for providing additional electrical power during periods of peak electrical power demand.

Abstract: A surface heat exchanger is provided which utilizes forward and aft brackets to retain the heat exchanger in position. The surface heat exchanger includes a plurality of core cooling channels as well as fins which are disposed for air flow through the gas turbine engine. The brackets include a low-friction wear material as well as an isolator sheet which provides some spring force on the heat exchanger.

Abstract: A system includes a multi-tube fuel nozzle of a turbine combustor. The multi-tube fuel nozzle includes a support structure defining an interior volume configured to receive an air flow; a plurality of mixing tubes disposed within the interior volume, wherein each of the plurality of mixing tubes comprises a respective fuel injector; and an outer annular wall configured to direct an air flow from an annulus between a liner and a flow sleeve of the turbine combustor at least partially radially inward into the interior volume through an air inlet and toward the plurality of mixing tubes, wherein the outer annular wall at least partially defines an air flow passage extending from the annulus to the interior volume.

Abstract: The present disclosure relates to a seal for an aircraft turbojet engine nacelle thrust reverser, the seal having body having a cross section with respect to a longitudinal axis of the seal that is substantially circular, and a base which is intended to be mounted in a seal support. The seal is notable in that the base of said seal has apertures spaced from each other.

Abstract: A thrust reverser system deployable from a core cowl is provided. The thrust reverser system may comprise one or more doors that are configured to overlap to reduce leakage and increate reverse thrust while deployed. The doors may be configured to vanes that are configured to split and/or direct airflow.

Abstract: A gas turbine includes a compressor and a combustor; a SOFC having a cathode and an anode; a first compression air supply line supplying compression air to the combustor; a second compression air supply line supplying compression air to the cathode; an exhaust air supply line supplying exhaust air discharged from the cathode to the combustor; a first fuel gas supply line supplying a fuel gas to the combustor; a second fuel gas supply line supplying a fuel gas to the anode; a fuel gas supply ratio change unit capable of changing a supply ratio of the fuel gas supplied to the combustor and the fuel gas supplied to the anode; an exhaust fuel gas supply line supplying an exhaust fuel gas discharged from the anode to the combustor; and a controller performing open-close control of the control valves and according to an operation state of the SOFC.