| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060798000 | Having ease of assembly or disassembly feature | 11 |
| 20100115965 | TORQUE LOAD TRANSFER ATTACHMENT HARDWARE - A bushing for assembling an exhaust duct liner with an exhaust duct in a gas turbine engine comprises a body, a bushing opening and a first tab. The body is insertable into a duct opening in an exhaust duct. The bushing opening extends through the body to receive a shaft of a liner fastener. The first tab protrudes from the body to extend into an interior of the duct to prevent rotation of a head of the liner fastener. | 05-13-2010 |
| 20110120143 | JET ENGINE NACELLE INTENDED TO EQUIP AN AIRCRAFT - The invention relates to a jet engine nacelle intended to equip an aircraft, formed from at least first and second half-shells and equipped with at least one locking system designed to lock one of the half-shells with respect to the other, the locking system being actuable between a locked state and an unlocked state by means of an actuating device ( | 05-26-2011 |
| 20120247123 | SECURING SYSTEM - A securing system is provided for securing a component to a gas turbine engine casing. The system has a core engine casing formed from a front casing portion and a rear casing portion. The casing portions have respective casing flanges at which the casing portions are detachably joined together. The system further has a component flange extending from a component. The component flange is configured so that, when the casing portions are joined together, the component flange is trapped between the casing flanges. One of the casing flanges projects inwards from the casing to form an airtight seal with the component flange on the inside of the casing. The other of the casing flanges projects outwards from the casing to form an airtight seal with the component flange on the outside of the casing. | 10-04-2012 |
| 20120260670 | APPARATUS TO SEAL WITH A TURBINE BLADE STAGE IN A GAS TURBINE - Disclosed is an apparatus configured to seal with a turbine blade stage of a gas turbine. The apparatus includes an outer shroud coupled to an inner shroud and configured to circumferentially surround the turbine blade stage. The inner shroud is configured to circumferentially surround the turbine blade stage to seal with the turbine blade stage and includes an attachment element configured to be inserted into the outer shroud to couple the inner shroud to the outer shroud. | 10-18-2012 |
| 20130091864 | QUICK DISENGAGING FIELD JOINT FOR EXHAUST SYSTEM COMPONENTS OF GAS TURBINE ENGINES - A quick disengaging field joint connects a first component of an exhaust system of a gas turbine engine to a second component of the exhaust system. The field joint includes a pair of opposed stepped liners connected via exterior-facing connecting flanges. The field joints can be disassembled entirely from outside the exhaust housing without requiring access to the interior of the exhaust housing. | 04-18-2013 |
| 20140060079 | AEROENGINE ARRANGEMENT - A core engine for a gas turbine engine, the core engine having a rotational axis, a casing and further comprises a gearbox assembly, a radial drive arrangement, pipes and connections and a connection structure; the radial drive arrangement drivingly connects the core engine to the gearbox assembly; the pipes and connections extend from the gearbox assembly to the core engine and the gearbox assembly is radially spaced from the casing by connection structure. A method of assembling the gas turbine engine includes the steps of axially translating the core engine and the fan casing so that the gearbox is adjacent the fan casing and attaching the gearbox assembly to the fan casing. | 03-06-2014 |
| 20140318150 | REMOVABLE SWIRLER ASSEMBLY FOR A COMBUSTION LINER - A novel and improved swirler assembly for use in a combustion liner of a gas turbine engine is disclosed. The swirler assembly is removable and sized to provide an improved assembly, repairability, and fit-up in the combustion liner. The swirler assembly includes a premix tube, premix swirler and a plurality of axially spaced mounting blocks for receiving removable fasteners for securing the swirler assembly to the combustor. | 10-30-2014 |
| 20140338363 | TURBINE GEAR ASSEMBLY SUPPORT HAVING SYMMETRICAL REMOVAL FEATURES - A gas turbine engine includes a fan including a plurality of fan blades rotatable about an axis, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a geared architecture arranged in a housing and driven by the turbine section for rotating the fan about the axis, and a support member that supports the geared architecture within the gas turbine engine. The support member includes an inner portion and an outer portion. One of the inner and outer portions is configured to be coupled to the geared architecture and the other of the inner and outer portions is configured to be coupled to the housing. A plurality of removal features each have at least one engaging surface to facilitate a pulling force on the support member in a direction parallel to the axis. The engaging surfaces on each of the removal features are oriented relative to each other to resist any bending moment on the support member during application of the pulling force. A gear system and method are also disclosed. | 11-20-2014 |
| 20150354826 | ANNULAR COMBUSTION CHAMBER OF A GAS TURBINE AND GAS TURBINE WITH SUCH A COMBUSTION CHAMBER - The invention relates to an annular combustion chamber of a gas turbine having a machine axis. The combustion chamber includes at least two zones. A first zone receives the fuel/air mixture of a plurality of burners. A second zone guides the hot gases being produced by the burners from the first zone to an entrance of a turbine section of said gas turbine. An annular transition liner includes a plurality of liner segments located at the transition between said first zone and second zone. Each of the liner segments includes with respect to the axial hot gas flow a leading edge, a trailing edge, and two sidewalls, and is provided with axial mounting means at the leading and trailing edges, such that the liner segment can be installed in axial direction and is axially fixed by respective segments of the neighboring first zone. Local spacer ribs are provided at the leading edge of the liner segments in order to establish a gap of minimum width between the liner segments and the fixing segments of the neighboring first zone. | 12-10-2015 |
| 20160010869 | COMBUSTION CHAMBER OF A GAS TURBINE WITH BOLTED COMBUSTION CHAMBER HEAD | 01-14-2016 |
| 20170234165 | GAS TURBINE ENGINE PACKAGE AND CORRESPONDING METHOD | 08-17-2017 |
