子分类：
| 序号 | 专利名 | 申请号 | 申请日 | 公开（公告）号 | 公开（公告）日 | 发明人 |
|---|---|---|---|---|---|---|
| 201 | AIRCRAFT DRIVE DEVICE AND AIRCRAFT PROVIDED WITH A DRIVE DEVICE | US15850350 | 2017-12-21 | US20180170565A1 | 2018-06-21 | Jürgen STEINWANDEL; Michael HOFMANN; Michael JUDAS |
| An aircraft drive device (10) for creating propulsion and/or lift of an aircraft, the drive device comprising a rotating shaft (12), a rotating shaft bearing (14, 16), a rotating shaft drive machine (18) for rotationally driving the rotating shaft (12), and a housing (20), wherein the rotating shaft bearing (14,16) and the rotating shaft drive device (18) are arranged in an interior of the housing (44) and the rotating shaft (12) protrudes to the outside of the housing (20) through a housing opening (42) of the housing (20), wherein in the region of the housing opening (42) an overvoltage arrester device (52) is arranged which is designed for diverting an overvoltage present at the rotating shaft (12). | ||||||
| 202 | AIRCRAFT FIRE SEAL STRUCTURE AND AIRCRAFT | US15831503 | 2017-12-05 | US20180163631A1 | 2018-06-14 | Akira Takeuchi |
| A fire seal structure prevents flame from coming out of a fire-prevention region of an aircraft. The fire seal structure includes: a plurality of walls including: a first wall provided on a first partitioning member; and a second wall provided on a second partitioning member. One of the plurality of walls is a spring wall that functions as a spring. The first and the second partitioning members define the fire-prevention region. The plurality of walls forms a labyrinth-shaped gap between the first and the second partitioning members. Each of the plurality of walls contains a refractory material and includes a front end part. When the first and the second partitioning members are stationary with respect to each other, the front end part is not in contact with another member, and the spring wall is disposed closest to a facing member, among the plurality of walls. | ||||||
| 203 | SEALING COOLING INNER FIXED STRUCTURE | US15372052 | 2016-12-07 | US20180156131A1 | 2018-06-07 | Timothy Olson |
| An inner fixed structure (IFS) seal arrangement may comprise an IFS comprising an inner skin, and outer skin, and a cellular core located between the inner skin and the outer skin, an IFS seal, a seal retainer configured to retain the IFS seal, and a cooling flow channel disposed between the outer skin and the seal retainer. | ||||||
| 204 | ACOUSTIC ATTENUATION STRUCTURE WITH A PLURALITY OF ATTENUATION DEGREES FOR A PROPULSION ASSEMBLY OF AN AIRCRAFT | US15884507 | 2018-01-31 | US20180148187A1 | 2018-05-31 | Laurent Georges VALLEROY; Marc VERSAEVEL; Bertrand DESJOYEAUX; Patrick GONIDEC |
| The present disclosure particularly relates to an acoustic attenuation structure for a propulsion assembly of an aircraft. The acoustic attenuation structure includes an acoustically reflective wall and a sandwich panel. The sandwich panel includes a honeycomb structure surrounded by two acoustically porous skins, a rear skin and a skin. The acoustically reflective wall and the sandwich panel are arranged in such a way as to be separated by a layer of air. | ||||||
| 205 | Detecting damage in a composite panel without removing overlying insulation | US14553229 | 2014-11-25 | US09964521B2 | 2018-05-08 | Song Chiou; Jared Victor; Vijay V. Pujar |
| Systems and methods for detecting damage in composite components are disclosed. A first attachment feature and a second attachment feature may be coupled to a composite component. A transmitting device may be coupled to the first attachment feature. A receiving device may be coupled to the second attachment feature. A signal may be transmitted from the transmitting device, through the first attachment feature, through the composite component, and through the second attachment feature to the receiving device. The signal may be analyzed to detect damage in the composite component. | ||||||
| 206 | Combined inlet laminar and thrust reverser cascade efflux flow control system | US14684503 | 2015-04-13 | US09951719B2 | 2018-04-24 | Nigel David Sawyers-Abbott |
| A nacelle assembly for a gas turbine engine according to an example of the present disclosure includes, among other things, a fan nacelle bounding a bypass flow path. The fan nacelle includes a first nacelle section and a second nacelle section. The second nacelle section includes a moveable portion movable relative to a forward portion to define a secondary flow passage. The first nacelle section includes an inlet lip. A thrust reverser is configured to selectively communicate a portion of bypass airflow between the bypass flow path and the secondary flow passage. A pump is configured to selectively communicate airflow between the inlet lip and the secondary flow passage. A method of flow distribution for a gas turbine engine is also disclosed. | ||||||
| 207 | ACCESSORY GEARBOX FOR A GAS TURBINE ENGINE | US15722421 | 2017-10-02 | US20180094717A1 | 2018-04-05 | Uwe MINKUS |
| A gas turbine engine arrangement includes an accessory gearbox which is mounted so as to be aligned in an axial direction along the engine. The accessory gearbox may be recessed at least partly into a casing of the engine. | ||||||
| 208 | AIRCRAFT ENGINE ILLUMINATION AND DIAGNOSTIC SYSTEM | US15706514 | 2017-09-15 | US20180079533A1 | 2018-03-22 | Mark E. SUCHEZKY; Neil H. CRAFT; Gregg G. WILLIAMS |
| An aircraft engine incorporates one or more externally-visible illuminators that are illuminated responsive to an activation signal. At least one of a quantity, a location, a pattern, a sequence, a color, or a color pattern of the one or more externally-visible illuminators, a control of at least one color of the one or more externally-visible illuminators, or a control of at least one intensity of the one or more externally-visible illuminators, is uniquely associated with the manufacturer of the aircraft engine, so as to provide for distinguishing the manufacturer of the aircraft engine while the aircraft engine is in operation in an aircraft. | ||||||
| 209 | Multi-zone active laminar flow control system for an aircraft propulsion system | US14713621 | 2015-05-15 | US09908620B2 | 2018-03-06 | Keith T. Brown; Stuart J. Byrne; Steven M. Kestler |
| A nacelle is provided for an aircraft propulsion system. The nacelle may include an outer barrel and an active laminar flow control system. The active laminar flow control system may include a plurality of suction sources and a plurality of arrays of perforations in the outer barrel. The active laminar flow control system may be configured with a plurality of zones. Each of the zones may include a respective one of the suction sources which is fluidly coupled with a respective one of the arrays of perforations in the outer barrel. | ||||||
| 210 | Core case heating for gas turbine engines | US14969774 | 2015-12-15 | US09896964B2 | 2018-02-20 | Steven Clarkson; Daniel K. Van Ness, II; Paul Thomas Rembish |
| A case for a gas turbine engine includes a core body. The core body defines a longitudinally extending core flow path, a laterally extending bleed air duct coupling the core flow path in fluid communication with the external environment, and a structure-supporting member spanning the bleed air duct. A heating element is connected to the core body and is in thermal communication with the structure-supporting member. | ||||||
| 211 | INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN | US15234067 | 2016-08-11 | US20180043997A1 | 2018-02-15 | Kishore Ramakrishnan; Shourya Prakash Otta |
| The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members mounted at one or more predetermined locations around a circumference of the fan shaft of the fan. More specifically, the predetermined location(s) has a swirl distortion exceeding a predetermined threshold. Further, the inlet assembly includes at least one airflow modifying element configured within the inlet so as to reduce swirl distortion entering the fan. | ||||||
| 212 | INLET ASSEMBLY FOR AN AIRCRAFT AFT FAN | US15234055 | 2016-08-11 | US20180043996A1 | 2018-02-15 | Kishore Ramakrishnan |
| The present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an aft fan mounted to a fuselage of an aircraft. Further, the inlet assembly includes a plurality of structural members, such as struts or strakes, mounted at predetermined locations around a circumference of the fan shaft of the fan at the inlet. The predetermined location(s) may be determined as a function of swirl distortion entering the inlet. As such, the structural member(s) are configured to reduce swirl distortion of the airflow entering the fan. In some embodiments, the inlet assembly may also include inlet guide vanes. In alternative embodiments, the inlet assembly may be absent of inlet guide vanes. | ||||||
| 213 | SLIDING FASTENER SYSTEMS TO ACCOMMODATE DIFFERENTIAL THERMAL GROWTH | US15209472 | 2016-07-13 | US20180016019A1 | 2018-01-18 | Jihad Ramlaoui |
| Sliding fastener systems can include a fastener configured to fixedly attach to a first component having a first thermal expansion coefficient, a tray configured to fixedly attach to a second component having a second thermal expansion coefficient different from the first thermal expansion coefficient, the tray defining a sliding surface, and at least a portion of a locking element configured to engage with the fastener, the locking element having a base with a contact surface that movably contacts the sliding surface of the tray when the tray, the fastener, and the locking element fasten the first component and the second component together. The fastener and the locking element are configured to move with the first component and the tray is configured to move relative to the fastener and the locking element with the second component when there is a differential thermal expansion between the first and second components. | ||||||
| 214 | SANDWICH PANEL DISBOND REPAIR | US15208448 | 2016-07-12 | US20180015710A1 | 2018-01-18 | Andrew L. Joslyn; John Oum; Samadi Yoeuth |
| A structural panel for an aircraft nacelle may comprise a first skin, a second skin, and a core sandwiched between them. A disbond may be present between the first skin and the core. A plurality of layers of blueprint film adhesive may be positioned over the disbond. The structural panel may be vacuum bagged, and the blueprint film adhesive may be heated. The blueprint film adhesive may flow through perforations in the first skin and bond the first skin to the core. | ||||||
| 215 | Laser Projected Engine Hazard Zone Systems And Methods | US15202640 | 2016-07-06 | US20180009547A1 | 2018-01-11 | Ethan J. Brewer |
| An aircraft engine hazard zone projection system is described that includes an engine having an engine inlet and an engine outlet, and engine housing, and a light-emitting system connected to the engine housing. The light-emitting system is configured to project light on a ground below the engine housing so as to form at least one predetermined hazard zone surrounding the engine. The at least one predetermined hazard zone identifies at least one of an area subject to an engine inlet suction force or an area subject to an engine outlet exhaust force. | ||||||
| 216 | GAS TURBINE ENGINE AND METHOD TO COOL A GAS TURBINE ENGINE CASE ASSEMBLY | US15185820 | 2016-06-17 | US20170363003A1 | 2017-12-21 | Joseph D. Evetts; William J. Riordan; Federico Papa |
| A method of cooling a gas turbine engine case assembly includes moving a fan air valve that is operatively connected to a pre-cooler having a bypass inlet that is configured to receive bypass air that bypasses a gas turbine engine core to facilitate a provision of bypass air through a fan air valve inlet to the bypass inlet to a first open position, in response to a core compartment temperature being greater than a target core compartment temperature. The method further includes bleeding the bypass air through a bypass outlet of the pre-cooler into a core compartment. | ||||||
| 217 | THERMAL INSULATION BLANKET AND THERMAL INSULATION BLANKET ASSEMBLY | US15176513 | 2016-06-08 | US20170356343A1 | 2017-12-14 | Andrew Michael Roach; David Patrick Calder; Graham Frank Howarth |
| A thermal insulation blanket assembly having a thermal insulation blanket including an aerogel insulation material having a first surface and a second surface that is oppositely-disposed from the first surface, a backing covering the second surface of the aerogel insulation material, and a skin layer covering the first surface of the aerogel insulation material. | ||||||
| 218 | AIRCRAFT WITH INJECTION COOLING SYSTEM AND INJECTION COOLING SYSTEM | US15104479 | 2013-12-23 | US20170342905A1 | 2017-11-30 | Qiang PANG; Guohua ZHONG; Yao LI |
| An aircraft with a turbofan engine assembly having at least one compressor, a nacelle surrounding the turbine engine and defining an annular bypass duct between the nacelle and the turbine engine, a thrust reverser having at least one moveable control surface, a thrust reverser locking system configured to selectively lock the thrust reverser and an injection cooling system | ||||||
| 219 | Thermal insulation support | US13441993 | 2012-04-09 | US09783285B2 | 2017-10-10 | Thomas Joseph Connelly |
| A thermal insulation support includes a body including a first edge portion and a second edge portion extending generally along a longitudinal axis of the body. A first flange extending a first distance from the first edge portion of the body is coupled to a first wall. A second flange extending a second distance from the second edge portion of the body is coupled to a second wall. | ||||||
| 220 | Method for making contoured acoustic structures | US15220043 | 2016-07-26 | US09779715B1 | 2017-10-03 | Matthew Seldal |
| An existing acoustic honeycomb panel having a radius of curvature is cut into segments that have longitudinal and lateral sides that extend between the edges of the honeycomb. The segments are bonded together along their longitudinal or lateral sides to form a segmented acoustic honeycomb in which the radius of curvature is different from the radius of curvature of the original acoustic honeycomb panel. The shape of the longitudinal or lateral sides of the segments and the thickness of the adhesive bond can be controlled to provide segmented acoustic honeycomb panels that are tightly curved and which are suitable for use in demanding noise damping applications, such as jet engine nacelles. | ||||||
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