| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 060262000 | Air passage bypasses combustion chamber | 17 |
| 20090188234 | SHARED FLOW THERMAL MANAGEMENT SYSTEM - A thermal management system includes at least two of a multiple of heat exchangers arranged in an at least partial-series relationship. | 07-30-2009 |
| 20090217643 | Gas discharge device for a vehicle engine - The present invention provides gas discharge technique useful on vehicle engines. The technique comprises an s-shaped duct having an ejector formed therein to mix hot exhaust gas with a relatively cooler airflow entrained by the ejector. The s-shaped duct may be comprised of a number of segments that when integrated form the s-shape and ejector. A variety of cross sections including circular and rectangular are provided for the s-shaped duct. The segments may be composed of smaller subsections and may further have cooling slots. A nacelle is provided to at least partially enclose the s-shaped conduit and may have an inlet that is in fluid communication with the ejector of the s-shaped conduit. | 09-03-2009 |
| 20090235641 | T-Jet - The following is the design of a personal, strap-on the back flight device. It is my version of an air breathing jetpack; it is designed to give an individual pilot vertical take-off, horizontal flight and vertical landing capabilities. | 09-24-2009 |
| 20090235642 | VALVE SYSTEM FOR A GAS TURBINE ENGINE - A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath. | 09-24-2009 |
| 20090235643 | VALVE SYSTEM FOR A GAS TURBINE ENGINE - A valve system intermediate a secondary flow path and a primary flow path to selectively communicate secondary airflow into the primary gas flow path and control airflow injected from a higher pressure plenum into a lower pressure flowpath. | 09-24-2009 |
| 20100005780 | METHOD OF MANUFACTURING A CMC FLOW MIXER LOBED STRUCTURE FOR A GAS TURBINE AEROENGINE - A method of fabricating a lobed structure for a gas turbine flow mixer having an annular upstream portion extended downstream by a portion forming a multilobed skirt, the method comprising:
| 01-14-2010 |
| 20100031631 | GAS TURBINE COMPRISING A GUIDE RING AND A MIXER - A gas turbine is disclosed. The gas turbine includes a rotor which is driven by a turbine, a stator, struts that are fixed to the stator downstream from the turbine and that configure a guide ring for deflecting the rotational flow of hot gas, and a mixer arranged on the downstream end of the hot gas channel. The guide ring and the mixer are structurally and fluidically combined, the struts of the guide ring being connected to the wall structure of the mixer in the region of their radially outer ends. | 02-11-2010 |
| 20100218483 | CONTROLLED FAN STREAM FLOW BYPASS - A turbine engine component of a turbofan engine fitted with a bypass air valve includes at least one turbine engine component having a surface with at least one aperture, said turbine engine component located from between a bypass fan duct and a turbine exhaust nozzle of the turbofan engine; a bypass air valve includes a liner concentrically disposed about the turbine engine component and parallel to a centerline of the turbofan engine, the liner has a surface including at least one aperture and at least one impermeable region, and means for actuating the liner about the turbine engine component; and a flow transfer location comprising an area proximate to a turbine exhaust stream flow. | 09-02-2010 |
| 20100326048 | GAS TURBINE ENGINE DUCT HAVING A COUPLED FLUID VOLUME - A fluid tank structure is disclosed that is integrated within a bypass duct of a turbofan gas turbine engine. The fluid tank structure includes a hollow interior that is closed off when the fluid tank structure is coupled with an inner wall of the bypass duct. The inner wall forms one wall of a fluid tank volume enclosed by the fluid tank structure. | 12-30-2010 |
| 20110126512 | TURBOFAN GAS TURBINE ENGINE AERODYNAMIC MIXER - A mixer for a turbofan engine includes a centerbody and a mixer nozzle. The mixer nozzle surrounds at least a portion of the centerbody and is spaced apart to define a core flow path between the mixer nozzle and the centerbody. The mixer nozzle is configured, when bypass air flows through the turbofan engine, to direct at least a portion of the bypass air to impinge on the centerbody. The mixer nozzle includes a plurality of circumferentially spaced mixer lobes that extend axially in a rearward direction and have a cross-section shape defined by a set of equations. | 06-02-2011 |
| 20120227375 | GAS TURBINE ENGINE - A cowl cavity is defined between a core engine cowl and an engine inner casing. A core fairing is provided at the forward end of the engine inner casing and has a radially outer surface which, with the cowl, provides an airwashed surface of a bypass duct. A plenum is provided within the cowl cavity and receives air from bypass flow through openings in the surface. Conduits extend from a partition defining the plenum to distribute air to locations and components within the engine. | 09-13-2012 |
| 20130232949 | PRESSURE BALANCED AREA CONTROL OF HIGH ASPECT RATIO NOZZLE - An apparatus for minimizing signature of an engine having an exhaust gas path and a fan gas path includes a casing, and a liner disposed within the casing. The exhaust gas path passes within the liner and the fan gas path passes between the liner and the casing. The casing is disconnected into an upstream portion and a downstream portion, each portion in registration with the liner, such that there are minimum pressure imbalances caused by the exhaust gas path, the fan gas path and ambient on the downstream portion. | 09-12-2013 |
| 20140075919 | TEC MIXER WITH VARIABLE THICKNESSES - A mixer of a bypass turbine aeroengine according to one embodiment, includes circumferential inner and outer flow surfaces in a wavy configuration to form a plurality of lobes of the mixer. The mixer has an upstream end portion of sheet metal with a first thickness and a downstream end portion of sheet metal with a second thickness less than the first thickness. | 03-20-2014 |
| 20140230404 | GAS TURBINE ENGINE EXHAUST MIXER - An exhaust mixer for a gas turbine engine includes an annular wall having upstream end adapted to be fastened to an engine case and a downstream end forming a plurality of inner and outer mixer lobes. A support member interconnects at least a number of the inner lobes, and includes a circumferentially extending stiffener ring located radially inwardly from the inner lobes and a series of circumferentially spaced apart mixer struts radially extending from the inner lobes to the stiffener ring. The mixer struts have a radial length at least equal to a width of a main gas path defined between the inner lobes and the exhaust cone such that the mixer struts extend entirely through the main gas path. The stiffener ring being fixed solely to the mixer struts such as to float with respect to the exhaust cone and permit relative movement therebetween. | 08-21-2014 |
| 20150107225 | VARIABLE IMMERSION LOBE MIXER FOR TURBOFAN JET ENGINE EXHAUST AND METHOD OF FABRICATING THE SAME - A method of fabricating a mixer for a gas turbine engine is provided. The method includes forming a forward end and an aft end of the mixer, and forming an annularly undulating contour that defines a plurality of core immersion lobes and a plurality of bypass immersion lobes between the forward end and the aft end. The plurality of bypass immersion lobes includes a first bypass immersion lobe and a second bypass immersion lobe. The first bypass immersion lobe has a first crown contour line extending from the forward end to the aft end of the mixer, and the second bypass immersion lobe has a second crown contour line extending from the forward end to the aft end of the mixer. The first crown contour line is different than the second crown contour line. | 04-23-2015 |
| 20160102634 | REVERSE FLOW SINGLE SPOOL CORE GAS TURBINE ENGINE - 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. | 04-14-2016 |
| 20160153397 | Gas Turbine Engine System for Modulating Flow of Fan By-Pass Air and Core Engine Air | 06-02-2016 |
