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Небесная энциклопедия

Космические корабли и станции, автоматические КА и методы их проектирования, бортовые комплексы управления, системы и средства жизнеобеспечения, особенности технологии производства ракетно-космических систем

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Мониторинг СМИ

Мониторинг СМИ и социальных сетей. Сканирование интернета, новостных сайтов, специализированных контентных площадок на базе мессенджеров. Гибкие настройки фильтров и первоначальных источников.

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Поддерживает ввод нескольких поисковых фраз (по одной на строку). При поиске обеспечивает поддержку морфологии русского и английского языка
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Применить Всего найдено 6272. Отображено 100.
02-02-2012 дата публикации

Fuselage structure made of composite material

Номер: US20120025022A1
Автор: Gabriele Dreher, Steffen Kunze, Thomas Thiess, Umberto Girard, Wolfgang Buchs
Принадлежит: EUROCOPTER DEUTSCHLAND GMBH

A fuselage structure, particularly an aircraft door ( 1 ) of composite material comprising at least one panel ( 2 ) and at least one beam ( 3 ) mounted to each other and the panel ( 2 ) with the panel ( 2 ) being formed of at least one group of composite layers ( 5, 6, 20 ). The at least one beam ( 3 ) is provided at least at one of its respective ends ( 10, 11 ) with a flange ( 13 ) suitable for adhesive engagement with the at least one panel ( 2 ). The at least one group of composite layers ( 5, 6, 20 ) of the panel ( 2 ) is in form locking engagement with this flange ( 13 ) of the beam ( 3 ).

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Номер записи: 1
23-02-2012 дата публикации

Composite stringer with web transition

Номер: US20120045609A1
Автор: John D. Morris, Kenneth L. Brook, Michael A. Lee, Niraj Patel
Принадлежит: Boeing Co

A stringer comprises a base portion and first and second webs extending outwardly from the base portion. The orientation of at least one of the first and second webs may transition from a first angle to a second angle within an angle transition zone.

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Номер записи: 2
28-06-2012 дата публикации

Method of designing natural laminar flow wing for reynolds numbers equivalent to actual supersonic aircraft

Номер: US20120166148A1
Автор: Hiroaki Ishikawa, Kenji Yoshida, Kisa Matsushima, Yoshine Ueda
Принадлежит: Japan Aerospace Exploration Agency JAXA

In designing supersonic aircrafts, a method of designing a natural laminar flow wing is provided which reduces friction drag by delaying boundary layer transition under flight conditions of actual aircrafts. A target Cp distribution on wing upper surface, suited to natural laminarization in which boundary layer transition is delayed rearward in desired Reynolds number states, is defined by a functional type having as coefficients parameters depending on each spanwise station, a sensitivity analysis employing a transition analysis method is applied to the parameters, and a search is performed for the optimum combination of parameters to delay transition rearward.

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Номер записи: 3
23-08-2012 дата публикации

Decompression arrangement for an aircraft

Номер: US20120214393A1
Автор: Jürgen Vogt, Katja Strobel
Принадлежит: AIRBUS OPERATIONS GMBH

A decompression arrangement includes first and second cabin lining elements each with an edge region, the edge region of the second being arranged at a smaller distance from an aircraft outer skin than the edge region of the first, an air discharge opening arranged between the first edge region and second edge region and which is adapted to discharge air from an inner region of the cabin delimited by the cabin lining elements into a region of the aircraft lying between the cabin lining elements and the aircraft outer skin, and a decompression element integrated in the second cabin lining element having a flap pivotable about an axis from a closed position, in which it closes a pressure equalising opening formed in the decompression element, into a first open position, in which it opens the pressure equalising opening if a first predetermined differential pressure acts on the decompression element.

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Номер записи: 4
20-09-2012 дата публикации

Device for detecting shocks on a structure

Номер: US20120234110A1
Автор: Jerome Dubost
Принадлежит: AIRBUS OPERATIONS SAS

A detection device ( 10 ) for shocks on a part ( 20 ) comprises a base ( 11 ) intended to be fixed by a lower face ( 111 ) onto a surface of the part ( 20 ) where the occurrence of a shock is to be detected and comprises one or more detectors ( 12 ) fixed at the base ( 11 ) and protruding with respect to the base, a detector ( 12 ) being deformed in a persistent way with a magnitude equal to or greater than a threshold magnitude As when it is subjected to a shock with an energy equal to or greater than a threshold energy Es. Under the effect of such a shock with energy Es or greater, a detector ( 12 ) is simply deformed or broken such that the visual inspection allows the occurrence of the shock to be detected.

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Номер записи: 5
18-10-2012 дата публикации

Multi-Element Airfoil System

Номер: US20120261517A1
Автор: David P. Lockard, Martin J. McKenney, Mehdi R. Khorrami, Raymond D. Atherley, Reggie T. Kidd, Travis L. Turner
Принадлежит: National Aeronautics and Space Administration NASA

A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.

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Номер записи: 6
18-10-2012 дата публикации

Active gurney flap

Номер: US20120261519A1
Автор: Paul R. Brewer, Reg R. Raval, Steven Shorcott
Принадлежит: Claverham Ltd

A gurney flap assembly has an actuator, a flexible or hinged body, the body flexing from a retracted to a deployed position in reaction to motion of the actuator, and a first seal extending along a first edge of the flexible body that flexes from the stowed position. Linear motion of the actuator output is transposed to the gurney flap thereby moving it from a retracted position into the airstream. This trailing edge device will improve airfoil lift.

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Номер записи: 7
17-01-2013 дата публикации

Rotary mandrel tool support

Номер: US20130014372A1
Автор: Jeffrey L. Miller
Принадлежит: Boeing Co

A method for creating a layup of reinforcing fibers comprises mounting a face sheet to a spindle of a mandrel tool support, the face sheet having a layup surface for the reinforcing fibers, and counterbalancing the mounted face sheet for stiffness and center of balance.

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Номер записи: 8
07-03-2013 дата публикации

Aircraft tail surface with a leading edge section of undulated shape

Номер: US20130056585A1
Автор: Raul Carlos Llamas Sandín
Принадлежит: AIRBUS OPERATIONS SL

An aircraft tail surface ( 11 ), such as a horizontal tail plane or a vertical tail plane, comprising a leading edge ( 14 ) having in at least a section along the tail span an undulated shape formed by a continuous series of smooth protrusions ( 17 ) and recesses ( 19 ) so that, in icing conditions, the ice accretion is produced only on the peaks ( 18 ) of said protrusions ( 17 ) and on the bottoms ( 20 ) of said recesses ( 19 ), thereby creating a channeled airflow and an arrangement of air vortices which impart energy to the airflow in the aerofoil boundary layer which delay the airflow separation that causes the stall, thus reducing the detrimental effects of ice accretion in its aerodynamic performance.

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Номер записи: 9
04-04-2013 дата публикации

Access Door Assembly and Method of Making the Same

Номер: US20130082143A1
Автор: Marc Storozuk
Принадлежит: Boeing Co

In another embodiment of the disclosure, there is provided an access door assembly for joining to a structure. The access door assembly has an access door with at least one access door nonlinear edge. The access door assembly further has a support structure with at least one support structure nonlinear edge. The access door assembly further has a doubler element attached to an interior side of the support structure. The support structure nonlinear edge is designed to interlace with the access door nonlinear edge to form an access door assembly for joining to a structure, the access door assembly having an interlaced nonlinear edge interface. A diameter of the doubler element of the access door assembly is preferably reduced as compared to a diameter of a doubler element of an access door assembly having a linear or circular edge, such that the reduced diameter preferably results in an overall reduced weight of the access door assembly and the structure to which the access door assembly is joined.

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Номер записи: 10
04-04-2013 дата публикации

Aircraft thermal insulation

Номер: US20130082144A1
Автор: Matthew Lang, Ruben James Joseph Fernandes
Принадлежит: BAE SYSTEMS plc

The invention provides an arrangement and methods for thermally insulating aircraft, particularly but not exclusively for when the aircraft is operating in extremely hot or cold conditions, and describes an aircraft skin construction including a foam-stiffened CFC sandwich panel forming part of the aircraft outer skin mounted to an underlying load bearing aircraft structure, wherein the panel at the mounting to the structure includes two outer layers of CFC material with an inner layer of foam material sandwiched therebetween.

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Номер записи: 11
23-05-2013 дата публикации

Aircraft assembly and method for producing an aircraft assembly

Номер: US20130125354A1
Автор: Axel Soeffker, Hermann Benthien
Принадлежит: AIRBUS OPERATIONS GMBH

An aircraft assembly is provided having a first aircraft component and a second aircraft component, at least one of the aircraft components being composed of a fibre-reinforced composite material, a connecting device connecting the first aircraft component to the second aircraft component, the connecting device including a first and a second section having a carrier element connected to the first and second aircraft components respectively, the first section having a first surface facing the first aircraft component and a second surface facing away from the first aircraft components, the second section having a first surface facing the second aircraft component and a second surface facing away from the second aircraft component. The first and second sections include a plurality of hook and loop elements extending from the second surface of the carrier element and cooperate to produce a hook and loop connection between the first and second aircraft components.

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Номер записи: 12
13-06-2013 дата публикации

Airfoil for rotor blade with reduced pitching moment

Номер: US20130149160A1
Автор: Ashish Bagai
Принадлежит: Sikorsky Aircraft Corp

A rotor blade for a rotary wing aircraft includes a root region extending from a rotor head to about 15% to 20% of a blade radius, a main region extending from a radial extent of the root region to about 80% to 95% of the blade radius, and a tip region extending from a radial extent of the main region to a blade tip. At least a portion of one of the root region, the main region and the tip region includes an airfoil profile section defined by a scaled set of coordinates in which a set of y/c coordinates listed in Table I are scaled by a selected factor.

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Номер записи: 13
18-07-2013 дата публикации

Device and method for assembling two sections of aircraft fuselage

Номер: US20130181092A1
Автор: Alain Depeige, Bruno Cacciaguerra, Helene Cazeneuve, Jean-Claude Lacombe, Marc-Antoine Castanet
Принадлежит: AIRBUS OPERATIONS SAS

Device for assembling two aircraft fuselage sections each comprising a skin and longitudinal stiffeners, the device comprising fishplates fishing couples of mutually aligned longitudinal stiffeners belonging respectively to said sections, each fishplate comprising two fishplate pieces each comprising at least one longitudinal sole plate intended to be fixed to a longitudinal stiffener, and a transverse support head, and at least one demountable connecting member able to mutually clamp the respective support heads of said fishplate pieces to allow a transmission of longitudinal forces between these fishplate pieces. Set of fuselage sections assembled by means of a device of the type described above. Method of assembling two fuselage sections by means of a device of the type described above.

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Номер записи: 14
15-08-2013 дата публикации

Auxiliary flap arrangement and aerodynamic body comprising such an auxiliary flap

Номер: US20130206918A1
Автор: Carsten Weber, Markus Fischer
Принадлежит: AIRBUS OPERATIONS GMBH

The invention relates to an auxiliary flap arrangement for modifying the profile of an aerodynamic body, in particular the trailing edge of the aerodynamic body. The auxiliary flap arrangement here exhibits at least two auxiliary flaps that can be adjusted by an adjustment device while coupled to each other.

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Номер записи: 15
21-11-2013 дата публикации

Reinforced Stiffeners and Method for Making the Same

Номер: US20130309443A1
Автор: Christopher G. Harris, Jeffrey L. Marcoe, John D. Morris, Karl B. Lee, Lyle R. Deobald, Mark C. Iden
Принадлежит: Boeing Co

A composite stiffener is fabricated using preforms of laminated, unidirectional composite tape. The stiffener includes a void that is reinforced by a filler wrapped with a structural adhesive. The surfaces of the preforms surrounding the void include a layer of composite fabric which is bonded to the filler by the adhesive, thereby increasing the toughness of stiffeners around the void and improving pull-off strength of the stiffener.

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Номер записи: 16
28-11-2013 дата публикации

Skin-stiffener transition assembly, method of manufacture and application of said skin-stiffener transition assembly

Номер: US20130316131A1
Автор: Walter Oefner
Принадлежит: EUROCOPTER DEUTSCHLAND GMBH

A skin-stiffener transition assembly ( 1 ) with a skin ( 4 ) and two L-shaped fiber-reinforced fabric preforms ( 2, 3 ) for a flange and a composite gusset filler ( 5 ) integrated between said skin ( 4 ) and said two L-shaped fiber-reinforced fabric preforms ( 2, 3 ). The skin ( 4 ) comprises two separate skin layers ( 7, 8 ) and the flange comprises a flange layer ( 6 ), each of said skin layers ( 7, 8 ) and said flange layer ( 6 ) being provided with cut outs ( 9, 10 ) along one side. Said skin layers ( 7, 8 ) and the flange layer ( 6 ) are attached to the composite gusset filler ( 5 ) with at least one of its cut-outs ( 9, 10 ) to one side of the essentially polygonal cross section of the composite gusset filler ( 5 ) and with at least one of its adjacent cut-outs ( 9, 10 ) to another side of the essentially polygonal cross section of the composite gusset filler ( 5 ). The present invention relates as well to a method of manufacture of said skin-stiffener transition assembly ( 1 ) and to applications of said skin-stiffener transition assembly ( 1 ).

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Номер записи: 17
02-01-2014 дата публикации

Joining Composite Fuselage Sections Along Window Belts

Номер: US20140001311A1
Автор: Bernhard Dopker, Henri Michael Ostroff, Wallace Chi-Hua Chan
Принадлежит: Boeing Co

A fuselage having a longitudinal window belt has a composite outer skin including upper and lower composite skin sections. The skin includes at least one window opening located at the window belt. The upper and lower skin sections are joined together by a longitudinal splice joint located at the window belt.

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Номер записи: 18
09-01-2014 дата публикации

Nanoreinforced films and laminates for aerospace structures

Номер: US20140011414A1
Автор: Anthony M. Mazany, Teresa May Kruckenberg, Vijay V. Pujar
Принадлежит: Goodrich Corp, Rohr Inc

A composite laminate for use on an external part of an aerospace vehicle has improved ultraviolet resistance and resistance to microcracking from thermal cycling. The laminate comprises a nanoreinforcement film, a support veil, and a composite layer. The laminate also can have a lightning strike protection layer and an external paint and primer. The nanoreinforcement film can comprise carbon nanomaterial and a polymer resin, and the composite layer has one or more layers of a reinforcement and a polymer resin. The carbon nanomaterial can be carbon nanofibers, and the nanoreinforcement film can have an areal weight of less than about 100 g/m2. The carbon nanomaterial can also comprise carbon nanofibers and carbon nanotubes.

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Номер записи: 19
23-01-2014 дата публикации

Method of inspecting impacts observed in fan casings

Номер: US20140020485A1
Автор: Julien Tran, Richard Lavignotte
Принадлежит: SNECMA SAS

A method for inspecting impacts present on an internal face of a fan casing, the method including: spotting a first impact present on the internal face of the fan casing; delimiting an inspection area containing the first impact; spotting the various impacts present in the delimited inspection area, the various spotted impacts forming a set of impacts to be considered; measuring, for each impact that is to be considered, the depth of length of the impact; for each impact to be considered, determining a harmfulness value, using at least one chart relating the depth and length of each impact to be considered to a level of harmfulness; determining, for the inspection area containing the first impact, a total harmfulness value by adding together the harmfulness level determined for each impact to be considered.

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Номер записи: 20
20-02-2014 дата публикации

High lift system for an aircraft and method for influencing the high lift characteristics of an aircraft

Номер: US20140048655A1
Автор: Bernhard Schlipf, Xavier Hue
Принадлежит: AIRBUS OPERATIONS GMBH

A high lift system for an aircraft includes a basic body, a flap which is movably mounted on the basic body and has a flap edge, and a retaining element. The high lift system is set up to form a gap between the flap edge and the basic body. The retaining element is mounted on a region of the flap close to the flap edge and extends towards the basic body to restrict the distance between the flap edge and the basic body. The retaining element is preferably configured as a linear attachment means. Consequently, a gap dimension between a flap and a basic body can be influenced to restrict flexing effects during loading of the flap and of the basic body.

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Номер записи: 21
20-03-2014 дата публикации

Discretely Tailored Multi-Zone Bondline for Fail-Safe Structural Repair

Номер: US20140076481A1
Автор: Eugene A. Dan-Jumbo, Everett A. Westerman, Russell L. Keller
Принадлежит: Boeing Co

A repair patch for reworking an inconsistent area of a composite structure includes a patch body adapted to cover the inconsistent area and having a first patch region, a second patch region outside the first patch region and a first separation zone between the first patch region and the second patch region, with the first patch region, the first separation zone and the second patch region having increasing interlaminar fracture toughness from a center to an edge of the patch body; and a layer of adhesive for bonding the patch body to the composite structure.

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Номер записи: 22
02-01-2020 дата публикации

CORE STRUCTURES FOR COMPOSITE PANELS, AIRCRAFT INCLUDING THE SAME, AND RELATED METHODS

Номер: US20200001565A1
Автор: Serencsits William L.
Принадлежит:

Core structures for composite panels, aircraft including the same, and related methods. A core structure includes a core body defining a first body side, a second body side, and a plurality of core cells. Each core cell includes at least one cell wall and defines a tubular cell void. Each core cell includes a bulk region and a transition region. Each cell wall is flared within the transition region to increase a surface area of the first body side relative to an average transverse cross-sectional area of the core body. A composite panel includes a core structure with at least one laminate ply coupled to the first body side of a core body. A method of manufacturing a composite panel includes forming a core body via an additive manufacturing process and operatively attaching at least one laminate ply to the core body. 1. A core structure for a composite panel , the core structure comprising:a core body defining a first body side, an opposed second body side, and a plurality of core cells extending between the first body side and the second body side, wherein each core cell:(i) includes at least one cell wall extending parallel to a cell axis and between a first end of the core cell and a second end of the core cell, wherein the first end is defined on the first body side and the second end is defined on the second body side;(ii) defines a tubular cell void that extends parallel to the cell axis at least partially between the first end and the second end of the core cell; and(iii) includes a bulk region and a transition region, the transition region extending between the bulk region and the first end of the core cell,wherein each cell wall has a bulk cell wall thickness, as measured in a direction perpendicular to the cell axis, within the bulk region, and each cell wall has a transition cell wall thickness, as measured in the direction perpendicular to the cell axis, within the transition region, the transition cell wall thickness varying across the transition region ...

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Номер записи: 23
07-01-2021 дата публикации

DEFORMING FOIL STRUCTURE FOR BRIDGING CURVED FLUID-DYNAMIC SURFACE

Номер: US20210001972A1
Автор: Chen Yong, FEREIDOONI Amin, MANGIONE Jaye, MARCHWICA Jan, WICKRAMASINGHE Viresh
Принадлежит: NATIONAL RESEARCH COUNCIL OF CANADA

A bridging structure for a deforming foil, such as a morphing wing, that provides a fluid-dynamic surface throughout foil deformation that forms a curved fluid-dynamic surface with a relatively low drag. A high extent of foil deformation can be provided, with lower actuation force, providing a fluid-dynamic surface with a simple or complex curve in one direction, by providing a set of rail-mounted members that are joined at one end to a deforming sheet. By coupling the members with high elongation, resilient bodies, adjacent members can support each other, while permitting extension, and accommodating curvature. 1. A bridging structure extensible to provide a curved fluid-dynamic suction or pressure surface for a deforming foil , the bridging structure comprising:a deformable sheet having an elastic elongation in a flexed direction of at least 5% of a length of the sheet in the flexed direction; a thickness having an elastic elongation less than ⅕th that of the flexed direction; and a second sheet direction perpendicular to the flexed direction;an array of at least four members running in a longitudinal direction, each member having: a top surface affixed to the elastomeric sheet in a direction perpendicular to the deforming direction; a pair of facing side-walls, each side-wall facing a sidewall of an adjacent member, except for outward facing side-walls of end members; a bottom surface opposite the top surface; and two ends;a set of resilient bodies interconnecting adjacent members, the resilient bodies deformable to extend spacings of the members in the deforming direction to increase an extent of the bridging structure in the flexed direction by at least 5% more than a closed pose of the bridging structure;a curved guide rail running perpendicular to the members; anda mounting mechanism for each member, coupling the member to the guide rail, the coupling permitting an angular pivot of at least 5° about an axis in the span direction during translation along the ...

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Номер записи: 24
05-01-2017 дата публикации

Passenger Aircraft

Номер: US20170001702A1
Автор: Pfingsten Kai-Christoph, Strotmann Arno
Принадлежит:

A passenger aircraft () comprising a pressurised cabin, which comprises a top volume () and a passenger deck (), the passenger deck () comprising at least one passenger compartment () separated from the rest of the passenger volume () by walls (), is characterised in that a burstable wall () that extends below the ceiling () of the passenger volume and defines the passenger compartment () is arranged so as to be adjacent to the side wall (), the space () between the burstable wall () and the side wall () being open towards the top volume () and forming a common pressure chamber therewith. 1. A passenger aircraft , comprising:a pressurised cabin having a length,wherein the pressurized cabin is enclosed laterally and vertically by a side wall, a top volume; and', 'a passenger deck; and', 'a lower deck,, 'wherein at least a portion of the length of the pressurized cabin compriseswherein the top volume and the passenger deck are separated vertically by a passenger deck ceiling,wherein the passenger deck and the lower deck are separated vertically by a passenger deck floor,wherein the top volume extends laterally from a first section of the side wall to a second section of the side wall,wherein the passenger deck extends laterally from the first section of the side wall to the second section of the side wall,wherein the lower deck extends laterally from the first section of the side wall to the second section of the side wall,wherein the top volume has a first volume,wherein the passenger deck has a passenger deck volume, 'at least one passenger compartment,', 'wherein the passenger deck compriseswherein the at least one passenger compartment has a corresponding at least one passenger compartment volume,wherein each passenger compartment of the at least one passenger compartment is separated from the rest of the passenger deck by a corresponding one or more walls of a corresponding at least one one or more walls,wherein each passenger compartment of the at least one ...

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Номер записи: 25
05-01-2017 дата публикации

UPPER JOINTS BETWEEN OUTBOARD WING BOXES AND WING CENTER SECTIONS OF AIRCRAFT WING ASSEMBLIES

Номер: US20170001707A1
Автор: CHARLES Jordan Daniel, Hamada Jason
Принадлежит:

An upper joint of a wing assembly of an aircraft includes an outboard upper wing panel, a center upper wing panel, a rib, and an upper joint assembly operatively interconnecting the outboard upper wing panel, the center upper wing panel, and the rib. For a substantial fore/aft span of the upper joint, a first subset of a plurality of center upper stringers is operatively coupled to the upper joint assembly, and a second subset of the plurality of center upper stringers terminates without being directly coupled to the upper joint assembly. 1. An upper joint for a wing assembly of an aircraft , the upper joint comprising:an outboard upper wing panel of an outboard wing box, wherein the outboard upper wing panel includes an outboard upper skin and a plurality of outboard upper stringers operatively coupled to the outboard upper skin;a center upper wing panel of a wing center section, wherein the center upper wing panel includes a center upper skin and a plurality of center upper stringers operatively coupled to the center upper skin;a rib that is located between the outboard wing box and the wing center section; andan upper joint assembly operatively interconnecting the outboard upper wing panel, the center upper wing panel, and the rib; the plurality of outboard upper stringers comprises hat-shaped stringers;', 'a first subset of the plurality of center upper stringers is operatively coupled to the upper joint assembly, wherein the first subset of the plurality of center upper stringers comprises hat-shaped stringers; and', 'a second subset of the plurality of center upper stringers terminates without being directly coupled to the upper joint assembly, wherein the second subset of the plurality of center upper stringers comprises blade stringers, Z-stringers, or I-stringers., 'wherein for a first substantial fore/aft span of the upper joint2. The upper joint of claim 1 , wherein the upper joint assembly includes:an outboard flange operatively coupled to the outboard ...

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Номер записи: 26
05-01-2017 дата публикации

Fore flap disposed on the wing of an aircraft

Номер: US20170001712A1
Автор: Bernhard Schlipf, Xavier Hue
Принадлежит: AIRBUS OPERATIONS GMBH

A leading edge slat arranged on the aerofoil of an aircraft. The leading edge slat is provided on the front of the main wing. The leading edge slat has a partially extended setting, with its trailing edge flat against the wing, and a further extended setting, with its trailing edge spaced apart from the nose of the wing to open a gap feeding high- energy air from the lower surface of the slat to the upper surface of the wing. The leading edge slat includes a body and a trailing edge facing the main wing, which can be bent around the spanwise direction of the slat relative to the body, and on which the trailing edge of the slat is provided, and which by means of a device generating a contact force is loaded for making contact between the trailing edge of the slat and the profile nose of the wing.

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Номер записи: 27
02-01-2020 дата публикации

SHORT TAKE OFF AND LANDING AIRCRAFT WITH ADJUSTABLE VORTICES DEVICE

Номер: US20200001982A1
Автор: Utt Larry
Принадлежит:

An improved aircraft system is provided. The improved aircraft system comprises an adjustable vortices device that may be attached to an aircraft to create various vortices effects, which increase take-off weight and improve low-speed handling of the aircraft. The adjustable vortices device comprises a linear actuator, a pivot mechanism, and a vortex generator. The pivot mechanism is operably connected to the linear actuator in a way such that the translational energy of the linear actuator causes the pivot mechanism to rotate about a central axis. The vortex generator is moveably attached to a surface of the aircraft and coupled to the pivot mechanism in a way such that rotating the pivot mechanism causes the vortex generator to rotate about a central axis, which alters the angle the vortex generators move through the air. 1. An aircraft system comprising:a fuselage, [ an internal structure,', 'an upper surface coupled to said internal structure, and', 'a lower surface coupled to said internal structure,', 'wherein said streamlined airfoil-contoured body is arranged such that said fluid moves at a higher average velocity over said upper surface and at a lower average velocity over said lower surface,, 'a streamlined airfoil-contoured body that generates lift when propelled through a fluid at an angle of incidence of at least zero, wherein said streamlined airfoil-contoured body is defined by,'}, a linear actuator configured to convert rotational energy into translational energy,', 'wherein said translational energy of said linear actuator causes said pivot mechanism to rotate about a central axis, and', 'a pivot mechanism operably connected to said linear actuator,'}, 'wherein rotating said pivot mechanism causes said vortex generator to rotate about said central axis,', 'a vortex generator coupled to said pivot mechanism,'}], 'an adjustable vortices device operably connected to said streamlined airfoil-contoured body, wherein said adjustable vortices device ...

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Номер записи: 28
03-01-2019 дата публикации

LOW STALL OR MINIMUM CONTROL SPEED AIRCRAFT

Номер: US20190002076A1
Автор: Bartlett Paul, Markley John K., Russell Jason
Принадлежит: JP AEROSPACE GROUP, INC.

A low stall or minimum control speed aircraft comprising a fuselage that has vertically flat sides; wings with high a lift airfoil profile of constant chord section set at zero degree planform sweep, twin booms having inner vertically flat surfaces, twin vertical stabilizers, a flying horizontal stabilizer; preferably twin engines having propellers and wherein each engine preferably has a thrust-line that is inclined nose-up to a maximum of +8 degrees, and is parallel to the wing chord underneath wing mounts and landing gear doors that provide surfaces for channeling propeller wash in a rearward direction; all working in concert so that the airplane has an extremely low stall speed and minimum control speed. The engines may be diesel, hydrogen fuel cell, electric fuel cell, diesel-electric, gas turbine or combinations thereof. The propellers may be counter-rotating. 1. A low stall speed aircraft , comprising:a fuselage having a first side and a second side, each of said first side and said second side comprising a vertically oriented flat surface;a first boom and a second boom, each comprising an inner vertically oriented flat surface, a forward end, and an after end, and further comprising at least one first propeller in said forward end of said first boom, and at least one second propeller in said forward end of said second boom, providing thrust in a forward direction;a first wing and a second wing, each of said first wing and said second wing comprising an inner portion having a wing root, an outer portion, a leading edge, and a trailing edge;a first vertical stabilizer and a second vertical stabilizer, each of said first and second horizontal stabilizer comprising a lower end and an upper end; anda flying horizontal stabilizer comprising a lower surface;wherein said first wing root is attached to said first side of said fuselage, and said second wing root is attached to said second side of said fuselage; andwherein said first boom, at its forward end, extends ...

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Номер записи: 29
03-01-2019 дата публикации

LEADING EDGE SKIN STRUCTURE

Номер: US20190002081A1
Автор: EVANS Rodney, HOLMES Vernon John
Принадлежит:

A leading edge skin panel for an aerodynamic structure of an aircraft. The skin panel includes attachment components for attaching the leading edge skin panel to the structure. A primary attachment component is configured to substantially prevent spanwise relative movement between the leading edge skin panel and the structure. The remaining attachment component are configured to permit a predetermined amount of spanwise relative movement between the leading edge skin panel and the structure. 1. A leading edge skin panel for an aerodynamic structure of an aircraft , comprisinga plurality of attachment components for attaching the leading edge skin panel to the aerodynamic structure,wherein a primary attachment component of the plurality of attachment components is configured to suppress spanwise relative movement between the leading edge skin panel and the aerodynamic structure, andthe or each remaining attachment component of the plurality of attachment components is configured to permit a predetermined amount of spanwise relative movement between the leading edge skin panel and the aerodynamic structure.2. The leading edge skin panel according to claim 1 , wherein the attachment components are each located a uniform distance from a leading edge of the leading edge skin panel claim 1 , and are spaced apart from each other along a spanwise direction.3. The leading edge skin panel according to claim 2 , wherein each attachment component is a distance from the leading edge which is less than half of a distance between of each attachment component and a trailing edge of the leading edge skin panel.4. The leading edge skin panel according to claim 1 , comprising an outer aerodynamic upper surface extending in a chordwise and spanwise direction claim 1 , and an inner surface; wherein the plurality of attachment components are provided on the inner surface.5. The leading edge skin panel according to claim 4 , wherein each of the plurality of attachment components is bonded ...

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Номер записи: 30
03-01-2019 дата публикации

Aircraft

Номер: US20190002098A1
Автор: Liu Angfeng
Принадлежит:

The embodiments of the present invention provide a navigator, comprising a gyro flying device and a cover that seals and encloses the gyro flying device. The gyro flying device is connected to the cover by a retaining mechanism. The gyro flying device comprises: a gyrorotor having an axisymmetric structure and rotatable around a central axis thereof; and a driving mechanism coaxially mounted with the gyrorotor to drive the gyrorotor to rotate around the central axis thereof, thereby manipulating rise and fall of the navigator. The retaining mechanism is further disposed to adjust an inclination angle of the gyro flying device, so as to adjust a flying direction of the navigator. The navigator has the advantages of quiet, safe, frictionless, extensive uses, etc. 1. A navigator comprising:a gyro flying device and a cover that seals and encloses the gyro flying device, the gyro flying device being connected to the cover by a retaining mechanism, the gyro flying device comprises:a gyrorotor having an axisymmetric structure and rotatable around a central axis thereof; anda driving mechanism coaxially mounted with the gyrorotor to drive the gyrorotor to rotate around the central axis thereof, thereby manipulating rise and fall of the navigator,wherein the retaining mechanism is further disposed to adjust an inclination angle of the gyro flying device, so as to adjust a flying direction of the navigator.2. The navigator according to claim 1 , wherein the navigator further comprises a vacuum maintaining system connected to the cover to maintain an interior of the cover in a vacuum state.3. The navigator according to claim 2 , wherein the retaining mechanism is connected to the gyro flying device through a bearing.4. The navigator according to claim 3 , wherein the retaining mechanism comprises a plurality of telescopic adjustment levers to achieve an adjustment of the inclination angle of the gyro flying device.5. The navigator according to claim 4 , wherein the navigator ...

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Номер записи: 31
01-01-2015 дата публикации

PANEL MEMBER FOR AN AIRFRAME

Номер: US20150004348A1
Автор: BULLINGER Oliver, VULTAGGIO Claudio, Zuardy Ichwan
Принадлежит:

The disclosure relates to a composite panel member for an airframe of an aircraft or spacecraft, the composite panel member having a laminated or sandwich structure including: a first outer layer extending over a first side of the panel member; a second outer layer extending over a second side of the panel member; a core layer between the first and second outer layers; and at least one support element configured as an electrical conductor and provided within the core layer between the first and second outer layers. In this regard, the at least one support element extends within the core layer substantially parallel to the first and second outer layers. 1. A composite panel member for an airframe of an aircraft or spacecraft , the composite panel member having a laminated or sandwich structure comprising:a first outer layer extending over a first side of the panel member;a second outer layer extending over a second side of the panel member;a core layer between the first and second outer layers; andat least one support element configured as an electrical conductor and provided in the core layer between the first and second outer layers;wherein the at least one support element extends within the core layer substantially parallel to the first and second outer layers.2. The panel member according to claim 1 , wherein the at least one support element is elongate and extends substantially continuously within the core in the direction substantially parallel to the first and second outer layers.3. The panel member according to claim 1 , wherein claim 1 , in a plane of a cross-section taken through the laminated or sandwich structure from the first side to the second side of the panel member claim 1 , the at least one support element spans a full width of the core layer from the first outer layer to the second outer layer to form a barrier through the core layer.4. The panel member according to claim 1 , wherein the at least one support element is adapted for electrical ...

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Номер записи: 32
07-01-2021 дата публикации

PROFILED STRUCTURE AND ASSOCIATED TURBOMACHINE

Номер: US20210003074A1
Автор: GEA AGUILERA Fernando, GRUBER Mathieu Simon Paul, Riou Georges Jean Xavier
Принадлежит: SAFRAN AIRCRAFT ENGINES

An airflow profiled structure having a profiled leading edge. The profiled leading edge having, along a leading edge line, a serrated profile line with a succession of teeth and depressions. The airflow profiled structure also includes a porous acoustically absorbent region located towards the bottom of the depressions. 1. A profiled air flow structure comprising:a body;porous acoustically absorbent regions;an upstream leading edge and/or a downstream trailing edge; andalong the upstream leading edge and/or the downstream trailing edge line, a serrated profile line showing a succession of teeth and depressions,wherein the porous acoustically absorbent regions locally form bottoms for the depressions where the porous acoustically absorbent regions occupy a part of the body to define, together with the body, the serrated profile line at the upstream leading edge and/or the downstream trailing edge.2. The profiled structure according to further comprising: along the chord, the serrated profile line has a maximum amplitude, h, and', {'br': None, 'i': 'd=h/', '10, within 30%.'}, 'the porous acoustically absorbent region has a geometric centre located at a distance d downstream of the upstream leading edge or upstream of the downstream trailing edge, at the bottom of the depressions such that], 'between upstream and downstream, a chord in which3. The profiled structure according to further comprising: along the upstream leading edge or the downstream trailing edge, the serrated profile line has a distance between two consecutive tooth tips,', 'along the chord, the serrated profile line has a maximum amplitude, h, and', along the upstream leading edge and/or the downstream trailing edge, two limits separated by a distance a such that a is equal to one third of said distance between two consecutive tooth tips, to within 30%,', 'along the chord, two limits separated by a distance b such that b=h/3, within 30%., 'the porous acoustically absorbent region has], 'between ...

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Номер записи: 33
13-01-2022 дата публикации

Active Lift Control Device and Method

Номер: US20220009618A1
Автор: John A. Cooney, Jr., Neal E. Fine
Принадлежит: Arctura Inc

A lift control device actively controls the lift force on a lifting surface. The device has a protuberance near a trailing edge of its lifting surface, which causes flow to separate from the lifting surface, generating regions of low pressure and high pressure which combine to increase the lift force on the lifting surface. The device further includes a means to keep the flow attached around the protuberance or to modify the position of the protuberance in response to a command from a central controller, so as to provide an active control of the lift between a maximum value and a minimum value.

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Номер записи: 34
14-01-2016 дата публикации

ELEVON CONTROL SYSTEM

Номер: US20160009370A1
Автор: Miralles Carlos Thomas, Olson Nathan, Plumb Robert Nickerson, Tao Tony Shuo
Принадлежит:

A system comprising an aerial vehicle or an unmanned aerial vehicle (UAV) configured to control pitch, roll, and/or yaw via airfoils having resiliently mounted trailing edges opposed by fuselage-house deflecting actuator horns. Embodiments include one or more rudder elements which may be rotatably attached and actuated by an effector member disposed within the fuselage housing and extendible in part to engage the one or more rudder elements. 1. An aerial vehicle , comprising:a fuselage having an exterior surface;a first pair of airfoils, a proximal portion of each of the first pair of airfoils rotatably attached at a forward pivot point at the exterior surface; anda second pair of airfoils, a proximal portion of each of the second pair of airfoils rotatably attached at an aft pivot point on the exterior surface;wherein the first pair of airfoils are configured to be rotatable to a retracted position that is aligned with a retracted position of the second pair of airfoils, with wing tips of the first pair of airfoils located adjacent to the aft pivot point when the first pair of airfoils is in its retracted position, and wherein wing tips of the second pair of airfoils are located adjacent to the forward pivot point when the second pair of airfoils is in its retracted position.2. The aerial vehicle of claim 1 , wherein the proximal portion of each of the first pair of airfoils is rotatably attached at a first coaxial pivot point.3. The aerial vehicle of claim 2 , wherein the proximal portion of each of the second pair of airfoils is rotatably attached at a second coaxial pivot point.4. The aerial vehicle of claim 1 , wherein each of the first pair of airfoils has a retracted position disposed along the exterior surface.5. The aerial vehicle of claim 4 , wherein each of the second pair of airfoils has a retracted position disposed along the exterior surface.6. The aerial vehicle of claim 5 , wherein the second pair of airfoils further comprise:an aft port airfoil and ...

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Номер записи: 35
14-01-2016 дата публикации

Stable Low Aspect Ratio Flying Wing

Номер: US20160009391A1
Автор: Friesel Eric Walter
Принадлежит:

A low aspect ratio flying wing provides aerodynamic stability throughout the flight envelope with improved aerodynamic efficiency. Insufficient stability and reduced aerodynamic efficiency typical of low aspect ratio flying wings is improved through wing design and proper application and placement of horizontal stabilizers and boundary layer control. Lateral asymmetric boundary layer manipulation is employed to alter flying wing orientation in flight. Lateral extension and retraction of the main structure wing optimizes efficiency. This novel flying wing is not found in literature or “prior art” and provides improvement in aerodynamic stability and efficiency over previous designs. Given the large amount of research, literature, patents and activity in the field since the 1930's and the absence of a practical design indicates the non-obvious nature of these disclosures. In addition, those skilled in the art teach away from present disclosures failing to realize the better than predicted advantages. 1. A flying wing comprising:An aspect ratio greater than 0.5 and less than 1.5; andA main structure of a wing, said main structure wing with a center of gravity located forward of 50 percent chord, said main structure wing aspect ratio greater than 0.5 and less than 1.5, said main structure wing is primary lifting airfoil; andHorizontal stabilizers mirrored about a central plane, said central plane is defined by the longitudinal and vertical axis of the main structure wing, said horizontal stabilizers extending outboard of the lateral extents of the main structure wing by a distance greater than 10 percent of the main structure wing longest chord length, said horizontal stabilizers offset vertically from the main structure wing chord line by a distance greater than 10 percent of the main structure wing longest chord length, and said horizontal stabilizers positioned longitudinally aft of the main structure wing 50 percent chord point.2. A flying wing of further comprising ...

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Номер записи: 36
14-01-2016 дата публикации

Adjustable Retaining Structure for a Cradle Fixture

Номер: US20160009421A1
Автор: Alfredo Jose Gerosa, Branko Sarh, Gregory Gudzinski, Harinder S. Oberoi, Kevin Marion Barrick, Quang T. Do, Richard Griffith Reese, Iv, Yuanxin Charles Hu
Принадлежит: Boeing Co

A method and apparatus for adjusting an adjustable retaining structure. The adjustable retaining structure may be rotated, passively, about a spherical interface as a panel applies a load to the adjustable retaining structure.

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Номер записи: 37
14-01-2021 дата публикации

METHOD OF FUSING THERMOPLASTIC COMPOSITE STRUCTURES

Номер: US20210009251A1
Автор: Wadsworth Mark Anthony
Принадлежит: SPIRIT AEROSYSTEMS, INC.

A method for fusing thermoplastic composite structures includes placing a substructure on an inner surface of a skin that is laid up on a shaping surface of a tool configured to maintain the shape of an outer mold line. The method further includes applying at least one insulation layer over a flange of the substructure and over exposed portions of the inner surface of the skin not in contact with the substructure, and applying a vacuum bag to at least partly enclose the skin and the substructure. The method yet still further includes applying heat to the shaping surface to fuse the substructure to the skin such that the skin exceeds its melting point and at least a portion of a raised segment of the substructure does not exceed its melting point. 1. A system for fusing composite structures , the system comprising: a substructure comprising a flange and a raised segment, the flange comprising a faying surface;', 'a skin comprising a first side and a second side opposite the first side, the second side of the skin comprising: (A) another faying surface contacting the faying surface of the flange, and (B) an exposed portion not contacted by the substructure;, 'a composite part comprising—'}an insulation layer covering the flange opposite the faying surface of the flange and covering the exposed portion of the skin, the insulation layer being at least two-tenths of an inch (0.2″) thick under ambient pressure; a shaping surface in a shape of an outer mold line and receiving the first side of the skin;', 'a heating element supplying heat to the shaping surface; and, 'a tool comprising—'}a vacuum bag covering the composite part such that the composite part is at least partially enclosed by the combination of the vacuum bag and the shaping surface of the tool.2. The system of claim 1 , wherein the vacuum bag comprises a first side adjacent the composite part and a second side opposite the first side claim 1 , the system further comprising a coolant fluid in contact with the ...

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Номер записи: 38
14-01-2021 дата публикации

ELLIPTICAL WING TIP AND METHOD OF FABRICATING SAME

Номер: US20210009256A1
Автор: AHLSTROM Eric A.
Принадлежит:

A wingtip of a lifting surface of an aeronautical vehicle, the lifting surface having a span, a leading edge, a trailing edge, an upper surface and a lower surface, the wingtip being in a range of five percent to fifteen percent of an end portion of the span of the lifting surface, the wingtip including: an elliptical shape between the leading and trailing edges, the elliptical shape tapering in a direction towards an outer edge of the wing tip, wherein the tapering occurs in a plurality of geometric parameters of the lifting surface including spanwise chord distribution between the leading and trailing edges, spanwise mean camber distribution between the leading including and trailing edges, spanwise maximum thickness between the upper and lower surfaces, and spanwise twist of a mean average of the spanwise chord distribution of the wingtip. 1. A wingtip of an airfoil of an aircraft , the airfoil having a span , a leading edge , a trailing edge , an upper surface and a lower surface , the wingtip being in a range of five percent to fifteen percent of an end portion of the span of the airfoil , the wingtip comprising: an elliptical shape between the leading and trailing edges , the elliptical shape tapering in a direction towards an outer edge of the wing tip , wherein the tapering of the wingtip occurs in four geometric parameters including spanwise chord distribution between the leading and trailing edges , spanwise mean camber distribution between the leading and trailing edges , spanwise maximum thickness between the upper and lower surfaces , and spanwise twist of a mean average of the spanwise chord distribution of the wingtip; andwherein the chord length is tapered in a range of 0.45 to 0.50 of the initial chord at 100% span, the mean camber distribution tapers to zero at 100% span, a thickness to chord ratio tapers to less than one percent at 100% span, and twist of a mean chord distribution of the span of the wingtip is in a range of negative one degree and ...

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Номер записи: 39
10-01-2019 дата публикации

ENCLOSURE COOLING FOR THERMAL MANAGEMENT OF UNMANNED AERIAL VEHICLES

Номер: US20190009878A1
Автор: Anderson Jon James, Chiu Chinchuan, Le Don, Sahu Vivek, WANG Peng, Wang Shujuan
Принадлежит:

Arrangements described herein relate to apparatuses, systems, and methods for a housing of an unmanned aerial vehicle (UAV), the housing includes but is not limited to a metallic porous material having a shape of an enclosure of the UAV, and a phase change material (PCM) provided in at least a portion of the metallic porous material. The metallic porous material and the PCM are configured to passively cool the UAV. 1. A housing of an unmanned aerial vehicle (UAV) , comprising:a metallic porous material comprising a first porous portion and a second porous portion, the metallic porous material configured to enclose one or more components of the UAV; anda phase change material (PCM), wherein pores of the first porous portion are free of the PCM and pores of the second porous portion are filled with the PCM,wherein the metallic porous material and the PCM are configured to passively cool the UAV.2. The housing of claim 1 , wherein the metallic porous material are configured to enclose one or more heat-generating components of the UAV.3. (canceled)4. (canceled)5. (canceled)6. The housing of claim 1 , whereinthe first porous portion of the metallic porous material is configured to form an exterior surface of the housing; andthe second porous portion of the metallic porous material is configured to form an interior surface of the housing.7. The housing of claim 1 , wherein pores of the first porous portion of the metallic porous material are configured to receive ambient air for cooling by convection.8. (canceled)9. The housing of claim 6 , whereinthe second porous portion of the metallic porous material is configured to face toward one or more heat-generating components of the UAV.10. The housing of claim 6 , wherein the second porous portion of the metallic porous material is configured to contact one or more heat-generating components of the UAV.11. (canceled)12. The housing of claim 1 , further comprising:a support structure configured to support the metallic porous ...

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Номер записи: 40
10-01-2019 дата публикации

High efficiency stall proof airfoil and means of control

Номер: US20190009890A1
Автор: George Paul Baker, Jr.
Принадлежит: Individual

A high-efficiency, stall-proof airfoil is an aircraft wing configuration whereby a motive force directly induces gaseous fluid flow across a lifting surface of the airfoil without requiring a movement of the wing through an air space. The airfoil is provided with means to control a pitch, a roll and a yaw motion and to control a position and stability of the aircraft. When not undergoing horizontal displacement, it provides highly efficient use of fuel resources, precluding the formation of drag and its incumbent power consumption. Air pressure at a bottom of the wing remains essentially ambient. Therefore, differential pressure between a lower surface of the wing and an upper surface of the wing maintains its maximum possible quantity. Virtually, all of the power consumed is utilized in a production of lift. Additionally, because lift is generated without regard to an angle-of-attack, forward speed, nor a configuration of a leading edge of the wing, the configuration is essentially stall proof.

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Номер записи: 41
09-01-2020 дата публикации

HIGH TEMPERATURE THERMOPLASTIC PRE-IMPREGNATED STRUCTURE FOR AIRCRAFT HEATED FLOOR PANEL

Номер: US20200010168A1
Автор: Botura Galdemir Cezar, Ching Nathaniel, Hu Jin, Slane Casey, Zhao Wenping
Принадлежит:

A heated floor panel assembly for aircraft includes structural layers made of a fiber matrix and a high temperature thermoplastic resin. The structural layers are within the heated floor panel assembly to protect the other assembly components from damage and absorb stress. The heated floor panel assembly further includes a heating layer with a heating element, an impact layer, and a core layer to take shear stress exerted on the assembly. 1. A floor panel assembly having a bottom surface and a top surface , the floor panel assembly comprising: structural layers comprising:', 'a first fiber matrix; and', 'a first high temperature thermoplastic resin infiltrating the first fiber matrix;, 'a first stack of structural layers adjacent the bottom surface, the first stack of'}a core layer, adjacent the first stack of structural layers, that absorbs shear stress; a second fiber matrix; and', 'a second high temperature thermoplastic resin infiltrating the second fiber matrix;', 'a heating layer between the core layer and the top surface; and', 'an impact layer between the heating layer and the top surface, wherein the first stack of structural layers, the core layer, the second stack of structural layers, the heating layer, and the impact layer are bonded together., 'a second stack of structural layers between the core layer and the top surface, the second stack of structural layers comprising2. The floor panel assembly of claim 1 , wherein the first fiber matrix and the second fiber matrix comprise fiber glass or carbon fibers.3. The floor panel assembly of claim 1 , wherein the first high temperature thermoplastic resin and second high temperature thermoplastic resin have a working temperature above 200 degrees Celsius.4. The floor panel assembly of claim 1 , wherein the first high temperature thermoplastic resin and second high temperature thermoplastic resin have a melting point between 150 and 400 degrees Celsius.5. The floor panel assembly of claim 1 , wherein the ...

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Номер записи: 42
09-01-2020 дата публикации

HEATED FLOOR PANEL WITH IMPACT LAYER

Номер: US20200010169A1
Автор: Ching Nathaniel, Hein Brandon, Hu Jin, Mullen James A., Owens George F., Slane Casey, Waina Daniel, Zhao Wenping
Принадлежит:

A heated floor panel assembly with an impact layer is made to absorb blunt force and protect a heating element inside the assembly. The assembly further includes a plurality of structural layer made of a reinforced polymer matrix, one or more core layer with a honeycomb or foam structure for absorbing shear stress, and a heating layer containing the heating element. The impact layer is a composite layer with a resin impregnated fiber matrix. 1. A floor panel assembly having a bottom surface and a top surface , the floor panel assembly comprising:a first stack of structural layers adjacent the bottom surface;a first core layer, adjacent the first stack of structural layers, that absorbs shear stress;a second stack of structural layers between the core layer and the top surface;a heating layer between the core layer and the top surface; and a fiber matrix; and', 'a thermoplastic resin infiltrating the fiber matrix;, 'an first impact layer between the heating layer and the top surface, the impact layer comprisingwherein the first stack of structural layers, the core layer, the second stack of structural layers, the heating layer, and the impact layer are bonded together.2. The floor panel assembly of claim 1 , wherein the first and second stacks of structural layers comprise a reinforced polymer matrix comprising a structural fiber matrix impregnated with a structural resin.3. The floor panel of claim 2 , wherein the structural fiber matrix is a carbon fiber or fiberglass.4. The floor panel of claim 2 , wherein the structural resin is polyether ether ketone claim 2 , polycarbonate claim 2 , polyphenylene sulfide claim 2 , polyetherimide claim 2 , epoxy claim 2 , phenolic claim 2 , bismaleimide claim 2 , benzoxazine claim 2 , or combinations thereof.5. The floor panel assembly of claim 1 , wherein the first core layer comprises a high density metallic material in a honeycomb structure selected from the group consisting of aluminum alloys claim 1 , stainless steel claim ...

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Номер записи: 43
15-01-2015 дата публикации

Vertical Takeoff and Landing (VTOL) Air Vehicle

Номер: US20150014475A1
Автор: Fisher Christopher Eugene, Ganzer David Wayne, Hibbs Bart Dean, King Joseph Frederick, Mukherjee Jason Sidharthadev, Parks William Martin, Taylor Dana J., Tokumaro Phillip T.
Принадлежит:

A flight control apparatus for fixed-wing aircraft includes a first port wing and first starboard wing, a first port swash plate coupled between a first port rotor and first port electric motor, the first port electric motor coupled to the first port wing, and a first starboard swash plate coupled between a first starboard rotor and first starboard electric motor, the first starboard electric motor coupled to the first starboard wing. 1. A flight control apparatus for a fixed-wing aircraft , comprising:a first port wing and a first starboard wing;a first port swash plate coupled between a first port rotor and a first port electric motor, the first port electric motor coupled to the first port wing; anda first starboard swash plate coupled between a first starboard rotor and a first starboard electric motor, the first starboard electric motor coupled to the first starboard wing.2. The apparatus of claim 1 , further comprising:a second port wing and second starboard wing;a second port swash plate coupled between a second port rotor and second port electric motor, the second port electric motor coupled to the second port wing; anda second starboard swash plate coupled between a second starboard rotor and second starboard electric motor, the second starboard electric motor coupled to the second starboard wing.3. The apparatus of claim 2 , further comprising:a horizontal stabilizer coupled to a fuselage and an elevator rotatably coupled to the horizontal stabilizer, the fuselage coupled between the first port wing and second starboard wing.4. The apparatus of claim 3 , further comprising:a port aileron rotatably disposed on a trailing edge of the first port wing; anda starboard aileron rotatably disposed on a trailing edge of the first starboard wing.5. The apparatus of claim 1 , further comprising:first and second landing gear attached to the first port wing and first starboard wing, respectively.6. The apparatus of claim 5 , further comprising:a third landing gear ...

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Номер записи: 44
03-02-2022 дата публикации

Composite Fabric Hat Stringers having Interleafed Tape Plies

Номер: US20220033049A1
Автор: Forouzan Behzadpour, Patrick Bickford Stickler
Принадлежит: Boeing Co

A composite hat stringer for stiffening a panel includes a plurality of composite fabric plies arranged to form a cap, a pair of flanges and a pair of webs respectively connecting the cap with the pair of flanges. The cap includes at least one 0° composite tape ply interleafed in the composite fabric plies within the cap.

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Номер записи: 45
03-02-2022 дата публикации

BEAD-STIFFENED MOVABLE SURFACES

Номер: US20220033061A1
Автор: Apdalhaliem Sahrudine, Ashmawi Waeil M., Behzadpour Forouzan
Принадлежит:

A movable surface of an aircraft has a front spar extending along a spanwise direction between opposing movable surface ends. The movable surface also includes a plurality of ribs defining a plurality of bays between adjacent pairs of the ribs. Each rib extends between the front spar and a trailing edge portion of the movable surface. The movable surface further includes an upper and a lower skin panels coupled to the ribs and the front spar. In addition, the bull surface includes a plurality of bead stiffeners coupled to an inner surface of at least one of the upper skin panel and the lower skin panel. The bead stiffeners within the bays are spaced apart from each other and are oriented non-parallel to the front spar and have a bead stiffener cap having opposing cap ends respectively locate proximate the front spar and the trailing edge portion. 1. A movable surface of an aircraft , comprising:a front spar extending along a spanwise direction between opposing movable surface ends;a plurality of ribs located at spaced intervals between the movable surface ends and defining a plurality of bays between adjacent pairs of the ribs, each rib extending between the front spar and a trailing edge portion of the movable surface;an upper skin panel and a lower skin panel coupled to the plurality of ribs and the front spar; anda plurality of bead stiffeners coupled to an inner surface of at least one of the upper skin panel and the lower skin panel, the bead stiffeners within the bays being spaced apart from each other and oriented non-parallel to the front spar and having a bead stiffener cap having opposing cap ends respectively locate proximate the front spar and the trailing edge portion.2. The movable surface of claim 1 , further comprising:a rear spar located proximate the trailing edge portion and extending between opposing movable surface ends;at least some of the plurality of ribs extending between the front spar and the rear spar;the upper skin panel and the lower ...

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Номер записи: 46
19-01-2017 дата публикации

TRUSSED STRUCTURE

Номер: US20170015401A1
Автор: Kaye Allan
Принадлежит:

A trussed structure comprising a frame and at least one strut, wherein the frame is of composite material and includes sockets which are integral with the frame. The invention also provides a process of making the trussed structures. The struts are typically of composite material and the trussed structures of the invention are particularly suitable for use in aircraft, for example, as wing ribs or floor beams. 1. (canceled)2. A trussed aircraft structure , comprising:a single member perimeter frame formed of layered and bonded composite material having at least two plies of fiber material bonded together, the perimeter frame having two long sides and two short sides, the two long sides comprising a first long side and opposing second long side separated by a space bounded by the perimeter frame;each of the first long side having formed therein at least two socket portions including a first socket portion and a second socket portion and the second long side having formed therein at least two further socket portions including a third socket portion and a fourth socket portion;the first socket portion being axially aligned with third socket portion and the second socket portion being axially aligned with the fourth socket portion;a first rectilinear strut having a first long axis, the first rectilinear strut extending through the first socket portion and the third socket portion, across the space separating the first long side and the second long side;a second rectilinear strut having a second long axis, the second rectilinear strut extending through the second socket portion and the fourth socket portion, across the space separating the first long side and the second long side;wherein the first long axis and the second long axis intersect at a location outside the perimeter frame when the first rectilinear strut is inserted in the first socket portion and the third socket portion and the second rectilinear strut are inserted into the in the second socket portion and ...

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Номер записи: 47
18-01-2018 дата публикации

Electromechanical hinge-line rotary actuator

Номер: US20180015999A1
Автор: Derick S. Balsiger, Nicholas R. Van De Veire
Принадлежит: Hamilton Sundstrand Corp

An electromechanical rotary actuator includes a drive member, a motor disposed inside and directly coupled to the drive member, and an output arm. The motor has a rotor configured toward an outside of the motor and directly coupled to an input of the drive member and a stator configured toward an inside of the motor and positioned inside the rotor. The output arm is disposed about the motor and is drivably connected to the drive member. The output arm defines an arcuate opening.

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Номер записи: 48
17-01-2019 дата публикации

Structural element of an aircraft part and method for manufacturing a structural element

Номер: US20190016064A1
Автор: Predrag Todorovic, Thomas Kubisch
Принадлежит: Rolls Royce Deutschland Ltd and Co KG

A structural element of an aircraft part, in particular an aircraft engine part, with at least partly a double-curvature shape, including a plurality of sets of fibers in textile fabric structure, wherein in at least one region of the structural element the number of fibers in one direction is reduced per area in a flattened out state of the textile fabric structure. It also relates to a method for manufacturing a structural element.

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Номер записи: 49
17-01-2019 дата публикации

STRINGER WITH PLANK PLY AND SKIN CONSTRUCTION FOR AIRCRAFT

Номер: US20190016436A1
Автор: Behzadpour Forouzan, Stickler Patrick B.
Принадлежит:

A composite assembly includes a composite skin which includes a plurality of composite plies and a composite stringer secured to the composite skin. A first composite plank ply is positioned between a first pair of composite plies of the plurality of composite plies of the composite skin; the first composite plank ply has a width dimension less than a width dimension of the plurality of composite plies of the composite skin; and the first composite plank ply extends along a length of the composite stringer. At least a portion of the composite stringer is positioned in overlying relationship with the first composite plank ply. 1. A composite assembly , comprising:a composite skin comprising a plurality of composite plies; a first composite plank ply is positioned between a first pair of composite plies of the plurality of composite plies of the composite skin;', 'the first composite plank ply has a width dimension less than a width dimension of the plurality of composite plies of the composite skin;', 'the first composite plank ply extends along a length of the composite stringer; and', 'at least a portion of the composite stringer is positioned in overlying relationship with the at least one composite plank ply., 'a composite stringer secured to the composite skin, wherein2. The composite assembly of claim 1 , wherein the composite stringer comprises one of an I-stringer claim 1 , hat stringer claim 1 , J-stringer claim 1 , blade stringer or Z-stringer.3. The composite assembly of claim 1 , wherein the composite skin extends along a length and width of a wing of an aircraft.4. The composite assembly of claim 3 , wherein the composite stringer extends in a direction along the length of the wing of the aircraft.5. The composite assembly of wherein the first composite plank ply comprises a plurality of unidirectional fibers extending along a direction of the length of the wing.6. The composite assembly of claim 3 , further comprising a second composite plank ply ...

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Номер записи: 50
17-01-2019 дата публикации

Fuselage and aircraft including an air distribution multifunctional substructure and assembly method

Номер: US20190016464A1
Автор: BRUNO Thomas, Jean-Mathieu Guimard, Michel Bermudez
Принадлежит: AIRBUS SAS

An aircraft fuselage including a fuselage skin, cross-members supporting a floor of the aircraft, wherein the fuselage includes a multifunctional substructure fixed to at least one of the cross-members in a lowered part of the cross-member. The multifunctional substructure comprises: at least one duct of an air distribution system having a substantially rectangular section at the location of the at least one cross-member and having lateral walls of the duct that are substantially vertical; a seat fixing track or a stiffener fixed on a first lateral wall of the duct of the multifunctional substructure, and; a seat fixing track or a stiffener fixed on a second lateral wall, opposite the first lateral wall, of the duct of the multifunctional substructure.

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Номер записи: 51
17-01-2019 дата публикации

Preceramic resin formulations, impregnated fibers comprising the preceramic resin formulations, composite materials, and related methods

Номер: US20190016640A1
Автор: Benjamin W.C. Garcia, David R. Nelson
Принадлежит: Northrop Grumman Innovation Systems LLC

A preceramic resin formulation comprising a polycarbosilane preceramic polymer, an organically modified silicon dioxide preceramic polymer, and, optionally, at least one filler. The preceramic resin formulation is formulated to exhibit a viscosity of from about 1,000 cP at about 25° C. to about 5,000 cP at a temperature of about 25° C. The at least one filler comprises first particles having an average mean diameter of less than about 1.0 μm and second particles having an average mean diameter of from about 1.5 μm to about 5 μm. Impregnated fibers comprising the preceramic resin formulation are also disclosed, as is a composite material comprising a reaction product of the polycarbosilane preceramic polymer, organically modified silicon dioxide preceramic polymer, and the at least one filler. Methods of forming a ceramic matrix composite are also disclosed.

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Номер записи: 52
16-01-2020 дата публикации

Methods of making hybrid laminate and molded composite structures

Номер: US20200016796A1
Автор: Adriana W. Blom, Edward M. Fisher, Jr.
Принадлежит: Boeing Co

Methods of making a composite structure comprise compression molding a fiber reinforced, thermoplastic component having a web and at least one flange integral with the web; laying up a fiber reinforced, thermoplastic cap; placing the fiber reinforced, thermoplastic cap on the flange; and joining the fiber reinforced, thermoplastic cap with the flange.

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Номер записи: 53
16-01-2020 дата публикации

METHOD AND SYSTEM FOR TUNING A TOW PLACEMENT SYSTEM

Номер: US20200016848A1
Автор: Hagman Thomas J.
Принадлежит:

A method and system for managing tow end placement. A tow is laid up over a tuning surface. A first position of a first tow end of the tow and a second position of a second tow end of the tow are measured. A first error between the first position of the first tow end and an expected position for the first tow end, and a second error between the second position of the second tow end and an expected position for the second tow end, are computed. A determination is made as to whether at least one of the first error or the second error is outside of selected tolerances. A start timing offset of a tow placement system is adjusted if the first error is outside of selected tolerances and a stop timing offset of the tow placement system is adjusted if the second error is outside of selected tolerances. 1. A method for calibrating a tow placement system , the method comprising:laying up a tow over a tuning surface;measuring a first position of a first tow end of the tow;measuring a second position of a second tow end of the tow;computing a first error between the first position of the first tow end and an expected position for the first tow end;computing a second error between the second position of the second tow end and an expected position for the second tow end;determining whether at least one of the first error or the second error is outside of selected tolerances; andadjusting a start timing offset of the tow placement system if the first error is outside of selected tolerances and a stop timing offset of the tow placement system if the second error is outside of selected tolerances.2. The method of claim 1 , wherein computing the first error comprises:computing a distance between the first position of the first tow end and a first marker on the tuning surface.3. The method of claim 2 , wherein computing the first error further comprises:computing a difference between a preselected distance and the distance computed between the first position of the first tow end and ...

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Номер записи: 54
16-01-2020 дата публикации

Method for manufacturing a one-piece reinforced structure and obtained structure

Номер: US20200017186A1
Автор: Manuel Torres Martinez
Принадлежит: Individual

A method for manufacturing a one-piece reinforced structure and obtained structure is provided, the method using base components made of partially-cured composite material and joining the base components together, applying a coating made of composite material on the base components, and applying heat on the assembly formed by the base components covered with the coating until a complete curing of the assembly is obtained, such that a one-piece reinforced structure made of composite material is obtained formed by the coating and the base components adhered to the coating, wherein the base components that form part of the very manufactured structure act as a mould during the manufacturing process, thus preventing the need to use moulds on which the composite material is deposited that must be subsequently removed from the final obtained structure.

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Номер записи: 55
16-01-2020 дата публикации

AIRCRAFT GENERATING LARGER LIFT BY REDUCTION OF FLUID RESISTANCE

Номер: US20200017198A1
Автор: ZHU Xiaoyi
Принадлежит:

The invention discloses a lift source for an aircraft comprising a fuselage and wings, wherein first channels are formed in the wings, a plurality of first inlets are formed in upper surfaces of the wings, a plurality of first pressure ports are formed in lower surfaces of the wings and are communicated with the first inlets via the first channels; and spoiler devices are arranged in the first channels and under the effect of the spoiler devices, form high-speed fluid layers on the upper surfaces of the wings, thereby generating a pressure difference from the lower surfaces of the wings which counteracts an external fluid pressure on the upper surfaces of the wings in the opposite direction, so a lift is generated by reduction of fluid resistance when fluid flows through the upper and lower surfaces of the wings, thereby developing a high-speed aircraft with a larger lift and thrust. 1. An aircraft , comprising a fuselage and wings , wherein first channels are formed in the wings , a plurality of first inlets are formed in upper surfaces of the wings and are communicated with the first channels , the first channels extend in a lengthwise direction of the wings from roots of sides close to the fuselage to tails of sides away from the fuselage and are communicated with an outside via first outlets , and the first channels are internally provided with spoiler devices used to extend fluid paths and form high-speed fluid layers together with the upper surfaces of the wings;a plurality of first pressure ports are formed in lower surfaces of the wings and are communicated with the first inlets in the upper surfaces of the wings via the first channels, and an air intake area of the first inlets is larger than that of the first pressure ports; a pressure difference is generated between the lower surfaces of the wings with a low external fluid flow speed and the high-speed fluid layers and counteracts an external fluid pressure on the upper surfaces of the wings in an ...

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Номер записи: 56
21-01-2021 дата публикации

FUEL RECIRCULATION SYSTEM

Номер: US20210016891A1
Автор: Dwier Thomas Edward, Wellemeyer Robert G.
Принадлежит:

Disclosed are systems and methods for maintaining bulk fuel temperatures in an aircraft. In one aspect, a recirculation system causes fuel to be delivered from a relatively low point near the feed hopper of each tank on the aircraft to one or more outboard locations of the wings. Once there, the fuel, due to gravity, flows back over the lower skin of the wing in channels back towards the fuselage, thus cooling the fuel. In other aspects, control systems are disclosed that coordinate the recirculation based on fuel levels in the tanks and fuel temperatures. The control systems also utilize a fuel scavenge system to maintain acceptable temperatures in the tanks. 1. An aircraft system comprising:a fuel tank;a delivery system configured to deliver fuel from an inboard location of the fuel tank to an outboard location of the fuel tank, the fuel tank being: (i) located in the wing; (ii) defined by internal structures of the wing; or (iii) both; andthe fuel delivery system being configured to deliver the fuel to at least one outlet located at an outboard section of the wing, the outlet being located to discharge such that inboard flow occurs due to gravity along at least a portion of the lower skin of the wing, thus reducing a temperature of fuel within the wing fuel tank.2. The system of wherein the fuel tank is substantially defined by internal surfaces of the wing.3. The system of wherein the fuel outlet is oriented to deliver fuel into at least one channel claim 1 , the at least one channel being at least partially defined by longitudinally-extending structural configurations rising from the internal surfaces of a lower skin of the wing.4. The system of wherein the at least one channel comprises three distinct channels including a central channel flanked by two outside channels claim 3 , all of the three channels being defined by the internal surfaces of the lower skin.5. The system of wherein the three channels are defined laterally by structural configurations ...

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Номер записи: 57
21-01-2021 дата публикации

FUEL RECIRCULATION SYSTEM

Номер: US20210016892A1
Автор: Dwier Thomas Edward, Wellemeyer Robert G.
Принадлежит:

Disclosed are systems and methods for maintaining bulk fuel temperatures in an aircraft. In one aspect, a recirculation system causes fuel to be delivered from a relatively low point near the feed hopper of each tank on the aircraft to one or more outboard locations of the wings. Once there, the fuel, due to gravity, flows back over the lower skin of the wing in channels back towards the fuselage, thus cooling the fuel. In other aspects, control systems are disclosed that coordinate the recirculation based on fuel levels in the tanks and fuel temperatures. The control systems also utilize a fuel scavenge system to maintain acceptable temperatures in the tanks. 1. A fuel system comprising:a fuel-temperature-control system for maintaining fuel temperatures in the fuel system, the fuel temperature-control system including a controller;a temperature-reading device located in the fuel system;a circulation-flow delivery subsystem, the delivery subsystem adapted to deliver fuel to a heat-dissipating medium associated with the fuel system;a return arrangement for returning fuel from the heat-dissipating medium;the controller being configured to activate the circulation-flow delivery subsystem upon the detection of a first reading made by the temperature-reading device, the first reading being reflective of a first fuel temperature which is either greater than or equal to a predetermined temperature maximum, the predetermined temperature maximum reflecting a fuel-cooling need.2. The fuel system of wherein the controller is further configured to deactivate the circulation-flow delivery subsystem detection of a second reading made by the temperature-reading device claim 1 , the second reading reflecting a second temperature which is either less than or equal to a predetermined temperature minimum claim 1 , the predetermined temperature minimum being reflective of a fuel-warming need.3. The fuel system of comprising:a scavenge system, the scavenge system adapted to draw fuel ...

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Номер записи: 58
21-01-2021 дата публикации

AIRCRAFT FLOORING AND METHODS OF MANUFACTURING THE SAME

Номер: US20210016899A1
Автор: BRUNO Joseph, GROSS Max
Принадлежит:

A floor panel for installation in an aircraft includes a plurality of thermoplastic C-shaped stringers, a consolidated thermoplastic deltoid filler, a thermoplastic upper facing sheet, and a thermoplastic lower facing sheet. The stringers are disposed in a parallel arrangement with one another. The deltoid filler is disposed within a longitudinally-extending notch defined by a pair of adjacent stringers. The upper facing sheet covers an upper surface of the stringers and the deltoid filler. The lower facing sheet covers a lower surface of the stringers and the deltoid filler. The stringers, the deltoid filler, and the upper and lower facing sheets are integrally consolidated forming a unitary construction. 1. A method of manufacturing a floor panel for installation in an aircraft , comprising:consolidating a plurality of Carbon-Glass/thermoplastic sheets to form a thermoplastic block having a desired thickness;cutting the thermoplastic block to form consolidated strips;machining each thermoplastic strip to form a deltoid filler;positioning a thermoplastic lower facing sheet on a base of a tool assembly;loading a plurality of thermoplastic C-shaped stringers on a plurality of mandrels;loading a plurality of stringer and mandrel combinations into the tool assembly to achieve a desired width;inserting the deltoid filler into a longitudinally-extending notch defined between adjacent stringers of the plurality of stringers;covering an upper surface of the plurality of stringers with a thermoplastic upper facing sheet to form a pre-consolidated floor panel; andapplying, via the tool assembly, a uniformly-distributed compressive force on an outer periphery of the pre-consolidated floor panel, thereby integrally consolidating the plurality of stringers, the deltoid filler, and the upper and lower facing sheets into a consolidated floor panel.2. The method according to claim 1 , wherein at least one of the upper facing sheet or the lower facing sheet extends to cover side ...

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Номер записи: 59
28-01-2016 дата публикации

FLAT-STOCK AERIAL VEHICLES AND METHODS OF USE

Номер: US20160023743A1
Автор: Barrett Ronald M., Bramlette Richard B., Honea Robert B.
Принадлежит:

A flat-stock aerial vehicle includes a body having a plurality of flat-stock sheets connected to one another, at least one motor, and at least three aerodynamic propulsors driven by the at least one motor. The aerodynamic propulsors can provide lifting thrust, pitch, yaw, and roll control in both helicopter-like hover flight and airplane-like translational flight. 1. A flat-stock aerial vehicle comprising: a first flat-stock sheet having a first forward edge and a first aft edge and having an aft slot therein extending forward from the first aft edge, and', 'a second flat-stock sheet having a second forward edge and a second aft edge and having a forward slot therein extending aft from the second forward edge, wherein the aft slot is configured to engage with the forward slot;, 'a body having a forward body edge, an aft body edge, and a longitudinal axis, the body includingat least one motor; andat least three aerodynamic propulsors positioned between the forward body edge and aft body edge, the at least three aerodynamic propulsors defining a forward wing portion and an aft wing portion, the at least three aerodynamic propulsors driven by the at least one motor, the at least three aerodynamic propulsors being configured to provide lifting thrust, pitch, yaw, and roll control to the vehicle.2. The flat-stock aerial vehicle of claim 1 , wherein at least a portion of the body is made from flat-stock which is planar or possesses simple curves.3. The flat-stock aerial vehicle of claim 1 , wherein at least a portion of the body is made from laminated flat-stock which is planar or possesses simple curves.4. The flat-stock aerial vehicle of claim 1 , wherein the body has at least one cutout to accommodate electronics claim 1 , motors claim 1 , propellers claim 1 , landing gear claim 1 , or lights.5. The flat-stock aerial vehicle of claim 1 , further comprising removable landing gear which extend laterally beyond the radial outermost position of the at least three ...

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Номер записи: 60
26-01-2017 дата публикации

PRESSURE BULKHEAD FOR AN AIRCRAFT FUSELAGE

Номер: US20170021908A1
Автор: Grase Karim, Jörn Paul
Принадлежит:

A pressure bulkhead () for an aircraft fuselage () having a sandwich structure () which defines a central axis () and which extends between a circumferential edge area (), wherein the sandwich structure () includes an inner cover layer (), an outer cover layer (), and a core layer (), which extends between the inner and outer cover layers () and connects them. The object of providing a pressure bulkhead () for an aircraft fuselage () which, even in the case of larger fuselage diameters, can particularly effectively absorb the occurring pressure forces and which can also be built with minimal weight, is achieved in that, between the inner and the outer cover layers (), in addition to the core layer (), a support structure () is provided, which is connected to the cover layers () and which extends from the inner cover layer () to the outer cover layer (), and in that the support structure () comprises at least one layer () which, viewed in a cross section () parallel to the central axis (), extends from a first section () of the edge area () to an opposite second section () of the edge area (). 1. A pressure bulkhead for an aircraft fuselage comprising: an inner cover layer which extends transversely to the central axis,', 'an outer cover layer which extends opposite the inner cover layer and transversely to the central axis, and', 'a core layer which extends between the inner and outer cover layers and connects them,', 'wherein between the inner and the outer cover layer, in addition to the core layer, a support structure is provided, which is connected to the inner and outer cover layers and which extends from the inner cover layer to the outer cover layer, and', 'the support structure comprises at least one layer which, viewed in a cross section parallel to the central axis, extends from a first section of the edge area to an opposite second section of the edge area., 'a sandwich structure which defines a central axis and which extends between a circumferential ...

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Номер записи: 61
26-01-2017 дата публикации

Method And Device For Automatic Management Of An Actuator Controlled By A Servo-Valve

Номер: US20170021913A1
Автор: Pierre Fabre
Принадлежит: AIRBUS OPERATIONS SAS

The device for automatic management of an actuator controlled by a servo-valve, includes a sensor for measuring the actual value of a parameter at the output of the actuator for a given control command, a computation unit for computing a theoretical value of said parameter, by applying the control command to a nominal performance model which models the operation of the actuator exhibiting nominal performance, a computation unit for computing the difference between the measured actual value of the parameter and the computed theoretical value of said parameter, a control unit for computing an adapted gain as a function of this difference, and a link for applying the adapted gain to the servo-valve for it to use it as gain value, so as to allow the actuator to restore nominal performance.

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Номер записи: 62
26-01-2017 дата публикации

Structural health monitoring sensory system integrated to a self-adapting morphing system

Номер: US20170021918A1
Автор: Fernando Dotta, Paulo Anchieta Da Silva, Ricardo Pinheiro Rulli, Tomaz Lazanha
Принадлежит: Embraer SA

A system and method for damage detection and for evaluating the real operation conditions for structural platforms using structural health monitoring is integrated to a system and method that permits for the platform to provide a flexible geometric control considering a self-adapting morphing which is capable of providing better operating structural platform performance.

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Номер записи: 63
25-01-2018 дата публикации

AIRCRAFT WITH A FUSELAGE AND A COMPOSITE TAIL BOOM

Номер: US20180022433A1
Автор: FINK Axel
Принадлежит: AIRBUS HELICOPTERS DEUTSCHLAND GMBH

An aircraft with a composite tail boom that comprises at least partly a tubular tail boom cone with an outer skin and an inner skin, wherein the inner skin delimits a hollow interior of the composite tail boom, wherein a plurality of rod-shaped stiffening elements and a plurality of ring-shaped stiffening elements are arranged between the outer skin and the inner skin, the plurality of rod-shaped stiffening elements being oriented in longitudinal direction of the composite tail boom and the plurality of ring-shaped stiffening elements being distributed along the longitudinal direction in the tubular tail boom cone. 1. A composite tail boom for an aircraft , the composite tail boom comprising at least partly a tubular tail boom cone with an outer skin and an inner skin , characterized in that the inner skin delimits a hollow interior of the composite tail boom , wherein a plurality of rod-shaped stiffening elements and a plurality of ring-shaped stiffening elements are arranged between the outer skin and the inner skin , the plurality of rod-shaped stiffening elements being oriented in longitudinal direction of the composite tail boom and the plurality of ring-shaped stiffening elements being distributed along the longitudinal direction in the tubular tail boom cone , and wherein at least one of the plurality of ring-shaped stiffening elements comprises an associated thickness in radial direction of the tubular tail boom cone that is smaller than an associated thickness of at least one of the plurality of rod-shaped stiffening elements in radial direction of the tubular tail boom cone.2. The composite tail boom of claim 1 ,wherein the associated thickness of the at least one of the plurality of ring-shaped stiffening elements is at least three times smaller than the associated thickness of the at least one of the plurality of rod-shaped stiffening elements.3. The composite tail boom of claim 1 ,wherein each two neighboring rod-shaped stiffening elements of the ...

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Номер записи: 64
10-02-2022 дата публикации

METHOD AND SYSTEM FOR ARRANGING SWARMING DRONES

Номер: US20220041279A1
Автор: ROWSE Philip Jonathan, WU Chengjun
Принадлежит:

The present invention provides a stackable drone and a drone swarm comprising at least two stackable drones. Each drone comprising: a fuselage comprising a first end and a second end; a mating structure arranged in the fuselage and configured to have an opening at the first end of the fuselage, the mating structure forming a mating recess on a first side of the fuselage, the mating recess having an opening at the first side of the fuselage for receiving a mating projection from a further stacking unmanned aerial vehicle. The stackable drones do not require a large area of ground for take-off and landing, require only a small space for storage and transportation. When landing, based on the conical or pyramidal structure, the drone may slide down by gravitational force into the mating recess of another drone thereunder without needs of high precision positioning or alignment system. 1. A stackable unmanned aerial vehicle , comprising:a fuselage comprising a first end and a second end;a mating structure arranged in the fuselage and configured to have an opening at the first end of the fuselage, the mating structure forming a mating recess on a first side of the fuselage, the mating recess having an opening at the first side of the fuselage for receiving a mating projection from a further stacking unmanned aerial vehicle;wherein the mating recess is configured to taper gradually from the first end to the second end such that cross-sectional area of the first end is greater than cross-sectional area of the second end and the gradual taper simultaneously forms a mating projection having a leading end with a cross-sectional area that is narrower than a base end having a larger cross-sectional area.2. The stackable unmanned aerial vehicle according to claim 1 , wherein the mating recess is configured to have a V-shaped longitudinal cross-section.3. The stackable unmanned aerial vehicle according to claim 2 , wherein the fuselage further comprises a connection element ...

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Номер записи: 65
25-01-2018 дата публикации

FOIL HINGE SYSTEM

Номер: US20180023618A1
Автор: Galeotti Giovanni
Принадлежит:

A foil hinge system for a hydrofoil or airfoil includes a main foil body substantially extending in a longitudinal direction L and a transverse direction T, a flap hinged to the main foil body arranged on a longitudinal end side of the main foil body and a flexible element connecting the main foil body and the flap Several rigid structures are arranged in such a manner on the flexible element that they provide a dead stop when the flap is in a maximum deflection with respect to the main foil body 1. A foil hinge system for a hydrofoil or airfoil , comprising:a main foil body, substantially extending in a longitudinal direction and a transverse direction,a flap hinged to the main foil body arranged on a longitudinal end side of the main foil body, anda flexible element connecting the main foil body and the flap,wherein several rigid structures are arranged in such a manner on the flexible element, that they provide a dead stop when the flap is in a maximum deflection with respect to the main foil body.2. The foil hinge system according to claim 1 , wherein the rigid structures are shaped and arranged on the flexible element such that a gap is provided between the adjacent rigid structures claim 1 , which enables movement of the flap relative to the main foil body.3. The foil hinge system according to claim 2 , wherein the rigid structures substantially extend in the transversal direction and the gap is provided in the longitudinal direction.4. The foil hinge system according to claim 1 , wherein the flexible element substantially extends in a longitudinal direction and transverse direction and connects the main foil body and flap in the longitudinal direction claim 1 , wherein the rigid structures are provided in a spanwise direction on both sides of the flexible element.5. The foil hinge system according to claim 1 , wherein the flexible element is made of a thin flexible composite sheet.6. The foil hinge system according to claim 1 , wherein the flexible element is ...

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Номер записи: 66
23-01-2020 дата публикации

METHOD FOR MANUFACTURING RESIN SHEET, RESIN SHEET, METHOD FOR MANUFACTURING STRUCTURAL BODY, STRUCTURAL BODY, AND AIRFRAME OF AIRCRAFT

Номер: US20200023932A1
Автор: ABE Toshio, ISHIKAWA Naomoto, Kamihara Nobuyuki, MURAOKA Mikio, TAKAYANAGI Toshiyuki
Принадлежит:

A method for manufacturing a resin sheet includes a coating step; a heating step; and a pressurizing step. In the coating step, linear metal nanomaterial is coated on a surface of a resin film having thermal plasticity. In the heating step, the resin film having the linear metal nanomaterial coated on the surface thereof is heated and softened. In the pressurizing step, the resin film having the linear metal nanomaterial coated on the surface thereof is pressurized to press the linear metal nanomaterial along a direction orthogonal to the surface on which the linear metal nanomaterial is coated. Thus, the coated linear metal nanomaterial penetrates the resin film to obtain the resin sheet containing the linear metal nanomaterial. 1. A method for manufacturing a resin sheet , comprising:an applying step of applying a linear metal nanomaterial to a surface of a resin film having thermoplasticity;a heating step of heating and softening the resin film to which the linear metal nanomaterial is applied; anda pressurizing step of pressurizing the resin film to which the linear metal nanomaterial is applied, along a direction orthogonal to the surface to which the linear metal nanomaterial is applied,wherein the applied linear metal nanomaterial is embedded in the resin film to form a resin sheet containing the linear metal nanomaterial.2. The method for manufacturing a resin sheet according to claim 1 ,wherein the linear metal nanomaterial is a nanofiber coated with a metal thin film.3. The method for manufacturing a resin sheet according to claim 1 ,wherein the linear metal nanomaterial is a nanocoil in which a metal thin film is formed in a coil shape.4. The method for manufacturing a resin sheet according to claim 1 ,wherein the linear metal nanomaterial is formed in a network shape, andin the applying step, the linear metal nanomaterial formed in a network shape is transferred to the surface of the resin film.5. The method for manufacturing a resin sheet according to ...

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Номер записи: 67
23-01-2020 дата публикации

COMPOSITE FUSELAGE ASSEMBLY AND METHODS TO FORM THE ASSEMBLY

Номер: US20200023934A1
Автор: ARANA HIDALGO Alberto, CANAS RIOS Maria Almudena, GARCIA NIETO Carlos, INIESTA LOZANO Fernando, MARTINO-GONZALEZ Esteban, MENARD Edouard, SÁNCHEZ PÉREZ Melania, Vázquez Castro Jesús Javier
Принадлежит:

A method for manufacturing a composite rear section assembly having a continuous skin solution including obtaining a vertical tail plane tooling that has intermediate tooling for tooling intermediate preforms of a vertical tail plane (), obtaining a fuselage barrel tooling (), (), () and () the fuselage barrel tooling having a cut-out (), (), () and longitudinal cavities (), (), () and (), attaching the vertical tail plane tooling to the fuselage barrel tooling by the cut-out of the fuselage barrel tooling, performing a composite skin lay-up () over the fuselage barrel tooling and the vertical tail plane tooling to obtain a continuous skin, curing the composite skin lay-up, the vertical tail plane tooling and the fuselage barrel tooling and demolding the tools to obtain the composite assembly with a continuous skin (), (). 1. A method for manufacturing a combined vertical tail plane and fuselage barrel assembly having a continuous skin solution , the method comprising:assembling a vertical tail plane tooling that includes an intermediate tooling, wherein the vertical tail plane tooling is configured to form intermediate preforms for a vertical tail plane;assembling a fuselage barrel tooling, wherein the fuselage barrel tooling includes a cut-out and longitudinal cavities; attaching the vertical tail plane tooling to the fuselage barrel tooling at the cut-out of the fuselage barrel tooling;performing a composite skin lay-up over the fuselage barrel tooling and the vertical tail plane tooling to obtain a continuous skin over outer surfaces of both the vertical tail plane tooling and the fuselage barrel tooling;curing the composite skin lay-up while over the vertical tail plane tooling and the fuselage barrel tooling; anddemoulding the vertical tail plane tooling and the fuselage barrel tooling from the cured composite skin lay-up to form the combined vertical tail plane and fuselage barrel assembly having the continuous skin.2. The method according to claim 1 , ...

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Номер записи: 68
23-01-2020 дата публикации

DOOR FRAME STABILIZATION

Номер: US20200023935A1
Автор: Malcolm Justin David Boynton, Martins Eder A., Nordman Paul S.
Принадлежит:

A stabilization assembly for a door frame secured to a fuselage skin of an aircraft and positioned about an opening defined in the fuselage skin including a strap, wherein a first end portion of the strap is secured to the fuselage skin with a first fastener which extends through the first end portion of the strap and through at least a portion of the fuselage skin. A second end portion of the strap is coupled to the door frame and a third portion of the strap extends between the first and second end portions of the strap and spaced apart from the fuselage skin. 1. A stabilization assembly for a door frame secured to a fuselage skin of an aircraft and positioned about an opening defined in the fuselage skin , comprising:a strap, wherein:a first end portion of the strap is secured to the fuselage skin with a first fastener which extends through the first end portion of the strap and through at least a portion of the fuselage skin;a second end portion of the strap is coupled to the door frame; anda third portion of the strap extends between the first and second end portions of the strap and spaced apart from the fuselage skin.2. The stabilization assembly of claim 1 , wherein the strap has a unitary construction.3. The stabilization assembly of claim 2 , wherein the strap is constructed of one of a composite laminate material or a metal material.4. The stabilization assembly of claim 1 , wherein the door frame is constructed of one of a composite laminate material or a metal material.5. The stabilization assembly of claim 1 , wherein:the first end portion of the strap is positioned overlying a portion of a stringer with the portion of the stringer positioned between the first end portion of the strap and the fuselage skin; andthe first fastener extends through the portion of the stringer.6. The stabilization assembly of claim 5 , wherein the portion of the stringer comprises a flange of the stringer.7. The stabilization assembly of claim 5 , further includes a plate ...

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Номер записи: 69
23-01-2020 дата публикации

CONNECTION ASSEMBLY FOR TRANSMITTING LOADS BETWEEN TWO WING ELEMENTS

Номер: US20200023939A1
Автор: KREY Dennis, LORENZ Florian, Schlipf Bernhard
Принадлежит:

A C-shaped connection assembly transmits loads in a load plane between a first and a second wing element. The connection assembly comprises a first and a second L-shaped load-bearing device. Each load-bearing device comprises a joint region and two legs extending parallel to the load plane and away from the joint region towards respective end regions. One leg of the first load-bearing device extends parallel to one leg of the second load bearing device. These legs are connected to one another. Two coupling portions which connect the connection assembly to the second wing element are formed in the respective joint regions of the load-bearing devices. Two further coupling portions which connect the connection assembly to the first wing element are formed in respective free end region of the load-bearing device and the joint region of the second load-bearing device. 1. A connection assembly for transmitting loads in a load plane between a first wing element and a second wing element ,wherein the connection assembly is C-shaped and comprises a first load-bearing device and a second load-bearing device,wherein each of the first and second load-bearing devices is L-shaped and comprises a first leg, a second leg and a joint region,wherein the first leg extends parallel to the load plane and away from the joint region towards a first end region and wherein the second leg extends parallel to the load plane and away from the joint region towards a second end region,wherein the second legs of the first and the second load-bearing devices extend in parallel to one another,wherein a first coupling portion for connecting the connection assembly to the second wing element is formed in the joint region of the first load-bearing device,wherein a second coupling portion for connecting the connection assembly to the second wing element is formed in the second end region of the first load-bearing device and the joint region of the second load-bearing device,wherein a third coupling ...

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Номер записи: 70
23-01-2020 дата публикации

Aerodynamic surface of an aircraft

Номер: US20200023945A1
Автор: Sergei Nikolaevich Nizov
Принадлежит: Individual

An aerodynamic surface of an aircraft comprises a main part having a leading and a trailing edges and having an airfoil section. The aerodynamic surface also having at least two vortex generators in the form of teeth having edges along the length thereof. The teeth are mounted on the leading edge of the main part so as to be capable of generating two vortex cores on one tooth. The edges of a tooth adjoin the leading edge of the main part of the aerodynamic surface. The radius of an edge of each tooth along the length of the vortex generator is five times less than the radius of the leading edge of the main part. The main part of the aerodynamic surface has a cambered airfoil section, wherein the teeth are mounted with a deflection towards the smallest degree of curvature of the airfoil section of the main part. The invention is intended for reducing an aerodynamic drag at low angles of attack while maintaining an increased load hearing capacity of the aerodynamic surface by generating vortex cores adjoining one of the sides thereof.

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Номер записи: 71
23-01-2020 дата публикации

Propelling System with Variable Aerodynamic Controls

Номер: US20200023952A1
Автор: Hesse Thomas Norman
Принадлежит:

A propelling system with variable aerodynamic controls is a system used to generate and control the flight forces of an aircraft. The system includes a stator, a rotor, a plurality of propelling units, and a control system. The stator serves as the stationary connection to the aircraft. The rotor revolves the propelling units about a central rotation axis. The control system enables the control of the propelling units. The propelling units generate the flight forces for the aircraft in the desired direction. In addition, each of the propelling units include a blade body, a shaft channel, a spar shaft, and at least one aileron assembly. The shaft channel receives the spar shaft within the blade body. The spar shaft connects the blade body to the rotor. The blade body passively corrects its angle of attack and supports the aileron assembly. The aileron assembly adjusts the pitch of the blade body. 1. A propelling system with variable aerodynamic controls comprises:a stator;a rotor;a plurality of propelling units;a control system;each of the plurality of propelling units comprising a blade body, a shaft channel, a spar shaft, and at least one aileron assembly;the blade body comprising a trailing edge;the rotor being rotatably mounted to the stator;the plurality of propelling units being radially positioned around a central rotation axis of the rotor;the rotor being terminally connected to the spar shaft for each of the plurality of propelling units;the spar shaft for each of the plurality of propelling units being positioned perpendicular to the central rotation axis of the rotor;the shaft channel traversing into the blade body;the shaft channel being positioned perpendicular to the central rotation axis of the rotor;the spar shaft being positioned within the shaft channel;the blade body being rotatably mounted about the spar shaft;the at least one aileron assembly being operatively integrated into the blade body, adjacent to the trailing edge, wherein the at least one ...

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Номер записи: 72
23-01-2020 дата публикации

METHOD OF DETERMINING AN INITIAL LEADING EDGE CIRCLE OF AIRFOILS OF A BLADE AND OF IMPROVING THE BLADE IN ORDER TO INCREASE ITS NEGATIVE STALL ANGLE OF ATTACK

Номер: US20200023953A1
Автор: Eglin Paul, FUKARI Raphael
Принадлежит: AIRBUS HELICOPTERS

A method of determining an initial leading edge circle for airfoils of a blade and of improving a blade, and also an improved blade and a advancement propeller including the improved blade. The radius of the initial leading edge circle of each airfoil of the blade is determined and then increased, and its leading edge is moved away from a pressure side half-airfoil towards a suction side half-airfoil, thereby modifying the airfoil of each cross-section of the blade and modifying the camber of each airfoil. Consequently, the absolute value of the negative stall angle of attack of the blade is increased, thus making it possible to increase the aerodynamic performance of the blade under a negative angle of attack compared with a blade that is not modified, and without significantly degrading its aerodynamic performance under a positive angle of attack. 1. A method of improving a blade , the blade extending in a longitudinal direction (X) spanwise from a first end to a second end , and in a transverse direction (Y) from a leading edge to a trailing edge , the blade having successive cross-sections , each cross-section being defined by an airfoil , each airfoil being defined by two half-airfoils including a suction side half-airfoil and a pressure side half-airfoil , each of the two half-airfoils comprising a leading edge segment , an intermediate segment , and a terminal segment , wherein the method includes:a first step of determining an initial leading edge circle for at least one of the two half-airfoils of at least one airfoil, each initial leading edge circle being attached to a respective half-airfoil;a second step of modifying at least one half-airfoil of at least one airfoil;a third step of moving the leading edge for each airfoil of the blade; anda fourth step of fabricating the blade with the modified airfoils; defining a straight line segment connecting the leading edge to the trailing edge of the airfoil;', 'creating a construction circle passing through the ...

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Номер записи: 73
29-01-2015 дата публикации

Aluminum Material Having Improved Precipitation Hardening

Номер: US20150027595A1
Автор: PALM Frank
Принадлежит:

An aluminum material for producing light-weight components includes aluminum (Al), scandium (Sc), zirconium (Zr) and ytterbium (Yb), where a weight ratio of scandium (Sc) to zirconium (Zr) to ytterbium (Yb) [Sc/Zr/Yb] is in a range from 10/5/2.5 to 10/2.5/1.25. 1. An aluminum material for producing light-weight components , the aluminum material comprising:aluminum (Al), scandium (Sc), zirconium (Zr) and ytterbium (Yb), wherein a weight ratio of scandium (Sc) to zirconium (Zr) to ytterbium (Yb) [Sc/Zr/Yb] is in a range from 10/5/2.5 to 10/2.5/1.25.2. The aluminum material of claim 1 , wherein the material further comprises:scandium (Sc) in an amount from 0.3 to 1.5% by weight, based on the total weight of the aluminum material.3. The aluminum material of claim 1 , wherein the material further comprises:zirconium (Zr) in an amount from 0.075 to 0.75% by weight, based on the total weight of the aluminum material, orytterbium (Yb) in an amount from 0.0375 to 0.375% by weight, based on the total weight of the aluminum material.4. The aluminum material of claim 1 , whereina) at room temperature, the material has a tensile strength or a yield strength in a range from 350 to 800 MPa, orb) following further heat treatment of the material in a temperature range from 300 to 400° C., at room temperature the tensile strength or the yield strength are higher than the tensile strength or yield strength of the same material produced without further heat treatment.5. A method for producing an aluminum material claim 1 , the method comprising the steps:a) providing an aluminum (Al)-scandium (Sc)-zirconium (Zr) base alloy;b) adding ytterbium (Yb) to the AlScZr base alloy from step a) to produce a molten aluminum (Al)-scandium(Sc)-zirconium(Zr)-ytterbium (Yb) alloy;{'sub': liquidus', '350° C., 'c) cooling the molten AlScZrYb alloy obtained in step b) in a temperature interval Tto Tat a cooling rate of ≧100 K/sec to produce a solidified AlScZrYb alloy; and'}d) heat treating the ...

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Номер записи: 74
28-01-2021 дата публикации

HYBRID METALLIC/COMPOSITE TUBE DESIGN TO TRANSFER BENDING, AXIAL, AND FLEXURAL SHEAR

Номер: US20210025526A1
Автор: BAIRD BRADLEY WILLIAM
Принадлежит: GOODRICH CORPORATION

A tube arrangement includes a composite tube defining a centerline axis, wherein the composite tube comprises a proximal surface and a distal surface, and an end fitting comprising a first end disposed within the composite tube and a second end extending from the composite tube, wherein an outer surface of the end fitting defines a flared portion defining a terminus of the first end, a lobe portion disposed axially from the flared portion, and a terminating portion disposed axially from the lobe portion, the proximal surface conforms to a geometry of the outer surface of the end fitting, the lobe portion and the flared portion mechanically lock the end fitting to the composite tube to mitigate movement of the end fitting relative to the composite tube. 1. A tube arrangement , comprising:a composite tube defining a centerline axis, wherein the composite tube comprises a proximal surface and a distal surface; andan end fitting comprising a first end disposed within the composite tube and a second end extending from the composite tube;wherein an outer surface of the end fitting defines a flared portion defining a terminus of the first end, a lobe portion disposed axially from the flared portion, and a terminating portion disposed axially from the lobe portion,the proximal surface conforms to a geometry of the outer surface of the end fitting,the lobe portion and the flared portion mechanically lock the end fitting to the composite tube to mitigate movement of the end fitting relative to the composite tube.2. The tube arrangement of claim 1 , wherein the lobe portion defines an annular ridge disposed around the end fitting.3. The tube arrangement of claim 2 , wherein the annular ridge defines a first convex fitting surface.4. The tube arrangement of claim 1 , wherein an annular groove is formed into the proximal surface of the composite tube claim 1 , the annular groove receives the lobe portion.5. The tube arrangement of claim 4 , wherein the annular groove defines a ...

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Номер записи: 75
29-01-2015 дата публикации

Integral Composite Bushing System and Method

Номер: US20150030389A1
Автор: Pollett Brandon
Принадлежит:

A composite bearing comprising a densified portion, wherein a hole location is positioned at the approximate center of said densified portion; and a plurality of filament tendrils, wherein the plurality of filament tendrils are configured to wrap around the hole location to create a “U” shape. 1. A composite aircraft structure , said composite aircraft structure comprising:a body portion having a first density; wherein the disk-shaped densified portion is configured to receive a hole positioned at an approximate center of said disk-shaped densified portion,', 'wherein, the densified portion has a density that is greater than said first density and is configured to resist a load imparted via said hole; and, 'a disk-shaped densified portion comprising ceramic, metal, high-strength plastic, or a combination thereof,'} wherein, the plurality of filament tendrils are configured to wrap around at least a portion of said hole to form a form a “U” shape,', 'wherein at least one of said plurality of filament tendrils comprises a carbon fiber material, a para-aramid synthetic fiber material, or a fiberglass material,', 'wherein, said plurality of filament tendrils direct stress away from said hole and into the body portion., 'a plurality of filament tendrils,'}2. The composite aircraft structure of claim 3 , wherein said densified portion's diameter is approximately 1.5 to 3 times said hole's diameter.3. A composite bearing for use in a composite structure claim 3 , the composite bearing comprising:a densified portion, wherein the densified portion is configured to receive a hole positioned at an approximate center of said densified portion; and wherein the densified portion is configured to resist a load imparted upon the composite bearing,', 'wherein plurality of filament tendrils direct stress away from the composite bearing and into the composite structure., 'a plurality of filament tendrils, wherein the plurality of filament tendrils are configured to encircle at least a ...

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Номер записи: 76
02-02-2017 дата публикации

Systems and methods for making composite structures

Номер: US20170028606A1
Автор: Daniel J. Perron, Jamie J. Langabeer, Jeffrey Olberg, Jonathan D. Robbins, Stephen G. Moore, Weidong Song
Принадлежит: Boeing Co

A system for depositing a composite filler material into a channel of a composite structure includes an end-effector configured to extrude a bead of the filler material into the channel. The filler material can comprise a first group of relatively long fibers, a second group of relatively short fibers and a resin. A drive system is configured to move the end-effector relative to the channel, and a position sensor is configured to detect the position of the bead relative to the channel. A controller is configured to operate the drive system in response to the detected position and to operate the end-effector to heat and compress the filler material so as to orient the longer fibers in a substantially longitudinal direction relative to the channel and the shorter fibers in substantially random directions relative to the channel when the bead is extruded into the channel.

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Номер записи: 77
02-02-2017 дата публикации

COMPOSITE STRUCTURE

Номер: US20170028673A1
Автор: Tilbrook David
Принадлежит: Hexcel Composites Limited

A composite structure () comprising one or more electrically conductive pathways () and one or more isolators for isolating the pathways () from the bulk of the structure (). 1. A composite structure comprising one or more electrically conductive pathways and one or more isolators for isolating the pathways from the bulk of the composite structure.2. A composite structure according to wherein the composite structure comprises fibre reinforcement and a reinforcement resin matrix claim 1 , said electrically conductive pathways being formed from said fibre reinforcement and said reinforcement resin matrix.3. A composite structure according to wherein the electrically conductive pathways are formed from the same fibre reinforcement and the same resin matrix as the bulk of the composite structure.4. A composite structure according to wherein the electrically conductive pathways are discrete.5. A composite structure according to wherein the isolators are formed by an isolator resin matrix.6. A composite structure according to wherein the isolator resin matrix comprises the reinforcement resin matrix.7. A composite structure according to wherein the composite structure comprises multiple ply layers of fibre reinforcement claim 2 , the isolator extending across at least two ply layers.8. A composite structure according to wherein the length of the isolator is n times the critical fibre length wherein n=1 to 10.9. A composite structure according to wherein the electrically conductive pathways are formed by unidirectional carbon fibre.10. A composite structure according to claim 9 , wherein the carbon fibre is coated with a metal.11. A method of controlling current paths in a composite structure comprising providing one or more electrically conductive pathways in the structure and isolating the pathways from the bulk of the structure.12. A method according to claim 11 , wherein the electrically conductive pathway is isolated from the composite structure by means of isolators. ...

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Номер записи: 78
02-02-2017 дата публикации

METHOD FOR JOINING THERMOSET COMPONENTS

Номер: US20170028698A1
Автор: Beier Uwe, Weiland Frank
Принадлежит:

A method for joining one or more first-type thermoset components to a second-type thermoset component, each first-type thermoset component being manufactured by providing an uncured starting thermoset component on which a thermoplastic material layer is placed, such that an interpenetrating network forms between the thermoset polymer of the starting thermoset component and the corresponding thermoplastic material layer when each first-type thermoset component is cured; and each first-type thermoset component being then placed on an uncured second-type thermoset component so that when the latter is cured, a further interpenetrating network forms between each thermoplastic material layer and the thermoset polymer of the second-type thermoset component; this results in a strong joint between the first-type and the second-type thermoset components. 1. A method of joining thermoset components comprising the steps of:providing at least one first-type thermoset component, each at least one first-type thermoset component being manufactured by:providing a starting thermoset component, the starting thermoset component being uncured,placing a thermoplastic material layer on a surface of the starting thermoset component, the thermoplastic material layer having a thermoplastic glass transition temperature,curing the starting thermoset component at a first-type curing temperature, thus giving rise to the first-type thermoset component having, once cured, a filmed surface coated with the thermoplastic material layer, the thermoplastic material layer being joined to the first-type thermoset component by means of a first-type interpenetrating network;providing a second-type thermoset component, the second-type thermoset component being uncured;placing the filmed surface of each at least one first-type thermoset component on the second-type thermoset component; andcuring the second-type thermoset component at a second-type curing temperature, so that a second-type interpenetrating ...

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Номер записи: 79
02-02-2017 дата публикации

CONTROL SURFACE FOR AN AIRCRAFT

Номер: US20170029090A1
Автор: CHU JAMES, Northam Robert
Принадлежит:

The present application relates to a control surface for an aircraft. The control surface has a leading edge, a trailing edge, and a chord-line defined between the leading edge and the trailing edge. A first aerodynamic surface is between the leading and trailing edges and a second surface is between the leading and trailing edges. The leading edge is formed by a nose, the nose having a hinge axis about which the control surface is deflectable. A maximum thickness of the control surface perpendicular to the chord-line between the first aerodynamic surface and the second surface is located aft of the hinge axis. The present application also relates to a control surface for an aircraft having a maximum curvature of the first aerodynamic surface of the control surface located aft of the hinge axis. The present application also relates to an aircraft or part of an aircraft comprising a fixed section and a control surface. 1. A control surface for an aircraft comprisinga leading edge, a trailing edge, and a chord-line defined between the leading edge and the trailing edge,a first aerodynamic surface between the leading and trailing edges,a second surface between the leading and trailing edges,the leading edge being formed by a nose, the nose having a hinge axis about which the control surface is deflectable,the first aerodynamic surface having an exposed airflow surface and the nose comprising an arced nose profile section extending from the exposed airflow surface wherein the arced nose profile section is configured to be at least partially exposable to airflow flowing over the control surface, wherein a maximum thickness of the control surface, perpendicular to the chord-line between the first aerodynamic surface and the second surface, is located aft of the hinge axis, anda radius of the arced nose profile section is greater than a perpendicular distance between the chord-line and the point of maximum thickness of the control surface on the first aerodynamic surface.2 ...

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Номер записи: 80
02-02-2017 дата публикации

METHODS FOR MANUFACTURING AN I-STRINGER OF AN AIRCRAFT AND DEVICES FOR USE IN SUCH METHODS

Номер: US20170029137A1
Автор: Aitken Charles, Autry Brenden A.
Принадлежит:

Methods for manufacturing a reinforced composite structure for an aircraft and devices used in such methods are provided. A device includes a base, a first support member fixedly attached to the base, and a second support member fixedly attached to the base and aligned longitudinally with the first support member. The first support member and the second support member are spaced a first distance apart. Two pinching wheels are spaced a second distance apart. The two pinching wheels are positioned proximate to ends of the first support member and the second support member. The second distance is less than the first distance. The two pinching wheels are configured to receive a composite material layout between them and to cause two lengths of the composite material layout to contact each other. 1. A device for forming a cap section of an I-stringer of an aircraft , the device comprising:a base;a first support member fixedly attached to the base;a second support member fixedly attached to the base and aligned longitudinally with the first support member, wherein the first support member and the second support member are spaced a first distance apart; andtwo pinching wheels spaced a second distance apart, wherein the two pinching wheels are positioned proximate to ends of the first support member and the second support member, wherein the second distance is less than the first distance, and wherein the two pinching wheels are configured to receive a composite material layout between them and to cause two lengths of the composite material layout to contact each other.2. The device of claim 1 , further comprising a first diagonal member supported by the first support member and having a first end claim 1 , anda second diagonal member supported by the second support member and having a first end, wherein the first diagonal member forms an angle with the second diagonal member and the first end of the first diagonal member is spaced a third distance from the first end of the ...

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Номер записи: 81
01-02-2018 дата публикации

COMPOSITE MATERIALS

Номер: US20180029686A1
Автор: Frulloni Emiliano, Lenzi Fiorenzo, Restuccia Carmelo Luca
Принадлежит: CYTEC TECHNOLOGY CORP.

A composite material that includes a layer of reinforcing fibres impregnated with a curable resin matrix and a plurality of electrically conductive composite particles positioned adjacent or in proximity to the reinforcing fibres. Each of the electrically conductive composite particles is composed of a conductive component and a polymeric component, wherein the polymeric component includes one or more polymers that are initially in a solid phase and are substantially insoluble in the curable resin, but is able to undergo at least partial phase transition to a fluid phase during a curing cycle of the composite material. 1. A curable composite material comprising:i) at least one structural layer of reinforcing fibres impregnated with a curable resin matrix; and{'sub': 'g', 'ii) at least one electrically conductive composite particle adjacent or in proximity to said reinforcing fibres, said conductive composite particle comprising a conductive component and a polymeric component, wherein the polymeric component of the conductive composite particle comprises one or more thermoplastic polymers that are initially in a solid phase and substantially insoluble in the curable resin matrix prior to curing of the composite material, but is able to undergo at least partial phase transition to a fluid phase by dissolving in the resin matrix during a curing cycle of the composite material, and wherein the one or more thermoplastic polymers has/have a glass transition temperature (T) of greater than 200° C.'}2. The composite material of claim 1 , wherein said curable resin matrix is a thermoset composition in which at least 50% of the polymeric component of the conductive composite particle is soluble in the resin matrix during curing of the composite material claim 1 , and wherein the phase transition to the fluid phase occurs by dissolution of the polymeric component in the resin matrix.3. The composite material of claim 1 , wherein the conductive component of each electrically ...

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Номер записи: 82
01-02-2018 дата публикации

SYSTEMS AND METHODS FOR ASSEMBLING A STRUCTURALLY REINFORCED COMPOSITE STRUCTURE

Номер: US20180029687A1
Автор: Koncz Tibor Albert
Принадлежит:

Systems and methods for assembling a structurally reinforced composite structure are disclosed herein. The methods include deforming a composite tubular skin to a deformed conformation to generate clearance to permit a frame assembly to be conveyed into an internal volume that is defined by the composite tubular skin. The methods further include conveying the frame assembly into the internal volume, permitting the composite tubular skin to deform from the deformed conformation to the target conformation, and operatively attaching the frame assembly to the composite tubular skin to form the structurally reinforced composite structure. The systems include a frame support that is configured to support the frame assembly, a frame deformation assembly, a skin support that is configured to support the composite tubular skin, and a skin deformation assembly. 1. A system for assembling a structurally reinforced composite structure , the system comprising:a frame support that is configured to support a frame assembly of the structurally reinforced composite structure during insertion of the frame assembly into an inner volume that is at least partially bounded by a composite tubular skin of the structurally reinforced composite structure;a frame deformation assembly that is configured to selectively deform the frame assembly to a deformed frame conformation to decrease a first dimension of the frame assembly in a first direction;a skin support that is configured to support the composite tubular skin during insertion of the frame assembly into the composite tubular skin; anda skin deformation assembly that is configured to selectively deform the composite tubular skin to a deformed skin conformation to decrease a first dimension of the composite tubular skin in the first direction and to concurrently increase a second dimension of the composite tubular skin in a second direction.2. The system of claim 1 , wherein the system further includes a conveyance structure that is ...

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Номер записи: 83
01-02-2018 дата публикации

ACTUATOR HARDOVER MONITOR

Номер: US20180029690A1
Автор: Eddy Stephen M., Hagerott Steven G., Peters Russell
Принадлежит:

An actuator hardover monitor for a control surface includes an actuator sensor for detecting an actuator position, a command model of an expected position of the actuator based on an input command, and a monitor to determine whether a difference between the actuator position and the expected position exceeds a threshold for a predetermined duration. A method of preventing a hardover event for a control surface includes commanding an actuator valve to a commanded position, determining continuously when the commanded position, or an actuator valve position, or a control-surface position, or a modeled actuator valve position exceeds a predetermined limit to provide an exceedance. The method may further include filtering a signal of the exceedance based on a time constant to provide a filtered exceedance, and switching to a backup control-surface actuator when the filtered exceedance exceeds the predetermined limit for a predetermined duration. 1. An actuator hardover monitor for a control surface , comprising:an actuator sensor for detecting an actuator position of a first control-surface actuator configured to control the control surface; a command model of an expected position of the control-surface actuator based on an input command;', 'a monitor to determine whether a difference between the actuator position and the expected position exceeds a threshold for a predetermined duration; and, 'a computer having a processor for executing software instructions stored in non-transitory memory, the software instructions comprisinga switch for switching to a second control-surface actuator configured to control the control surface when the difference exceeds the threshold for at least the predetermined duration.2. The actuator hardover monitor of claim 1 , further comprising a controller to adjust the actuator position of the first control-surface actuator to reduce the difference between the actuator position and the expected position.3. The actuator hardover monitor of ...

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Номер записи: 84
17-02-2022 дата публикации

CFRP FUSELAGE FRAME WITH SECUREMENT TO VERTICAL TAIL FIN

Номер: US20220048610A1
Автор: Diep Paul, Pham Phiyen T.
Принадлежит:

A securement assembly for securing a vertical tail fin assembly to an aircraft includes a first lug member secured to the vertical tail fin assembly. The securement assembly further includes a first clevis member. The first clevis member includes a first end portion of the first clevis member is engaged to the first lug member. A second end portion of the first clevis member is secured to a first fuselage frame constructed of a composite material with a first fastener which extends through the second end portion and the first fuselage frame in a first direction transverse to the first fuselage frame. 1. A securement assembly for securing a vertical tail fin assembly to an aircraft , comprising:a first lug member secured to the vertical tail fin assembly; and a first end portion of the first clevis member is engaged to the first lug member; and', 'a second end portion of the first clevis member is secured to a first fuselage frame constructed of a composite material, with a first fastener which extends through the second end portion and the first fuselage frame in a first direction transverse to the first fuselage frame., 'a first clevis member, wherein2. The securement assembly of claim 1 , wherein:the first clevis member has a first prong which defines a first opening at the first end portion of the first clevis member and has a second prong which defines a second opening at the first end portion of the first clevis member; andthe first lug member is positioned between the first prong and the second prong of the first clevis member.3. The securement assembly of claim 2 , further includes a first pin which extends through the first opening of the first prong claim 2 , the second opening of the second prong and through a first lug opening defined by and through the first lug member engaging the first clevis member to the first lug member such that the first pin extends in the first direction transverse to the first fuselage frame.4. The securement assembly of claim 2 ...

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Номер записи: 85
17-02-2022 дата публикации

AERIAL VEHICLE

Номер: US20220048617A1
Автор: FUJIMOTO Keiichiro, SHIMA Eiji, TSUTSUMI Seiji
Принадлежит:

[Objective] To provide, as to an aerial vehicle equipped with a multicopter mechanism, an aerial vehicle having both a vertical take-off and landing function and a horizontal cruise function and having an excellent cruising performance. 1. An aerial vehicle , comprising:a propulsion unit that includes a rotary shaft extending in a first direction and thrust producing mechanisms provided at both ends of the rotary shaft and produces a propulsion force for flying in air; anda fuselage unit that is suspended from the propulsion unit below the rotary shaft, has a center of gravity at a position below the rotary shaft, is configured to be freely rotatable around the rotary shaft, and is capable of storing an article.2. The aerial vehicle according to claim 1 , whereinin the propulsion unit, the thrust producing mechanisms turn into a vertical attitude in which thrust axes of the thrust producing mechanisms become parallel to a direction of gravity or in a horizontal attitude in which the thrust axes become perpendicular to the direction of gravity, the fuselage unit takes off and lands in the vertical attitude with the propulsion force, and the fuselage unit horizontally cruises in the horizontal attitude with the propulsion force.3. The aerial vehicle according to claim 1 , whereinthe fuselage unit has a shape having a short direction in the first direction and having a longitudinal direction in a second direction orthogonal to the first direction and maintains a horizontal state in a state in which the propulsion unit is suspended.4. The aerial vehicle according to claim 1 , whereinthe fuselage unit is provided with a tail unit at rear of the propulsion unit.5. The aerial vehicle according to claim 4 , whereinthe tail unit is provided with a first aileron.6. The aerial vehicle according to claim 1 , whereinthe fuselage unit is provided with a fan that provides a rotational force that rotates the fuselage unit around the rotary shaft at a rear of the propulsion unit.7. ...

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Номер записи: 86
31-01-2019 дата публикации

VTOL Aircraft for External Load Operations

Номер: US20190031339A1
Автор: McCullough John Richard, Oldroyd Paul K.
Принадлежит: Bell Helicopter Textron Inc.

An aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation. The aircraft includes an airframe having first and second wings with first and second pylons extending therebetween. The first and second wings each having first and second outboard nacelle stations. A two-dimensional distributed thrust array is attached to the airframe. The thrust array including a plurality of outboard propulsion assemblies coupled to the first and second outboard nacelle stations of the first and second wings. A flight control system is coupled to the airframe and is operable to independently control each of the propulsion assemblies. A cargo hook module is coupled to the airframe. The cargo hook module is operable for external load operations. 1. An aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation , the aircraft comprising:an airframe having first and second wings with first and second pylons extending therebetween, the first and second wings each having first and second outboard nacelle stations, the airframe having a longitudinal axis and a lateral axis in the VTOL orientation;a two-dimensional distributed thrust array attached to the airframe, the thrust array including a plurality of outboard propulsion assemblies coupled to the first and second outboard nacelle stations of the first and second wings;a flight control system coupled to the airframe and operable to independently control each of the propulsion assemblies; anda cargo hook module coupled to the airframe, the cargo hook module operable for external load operations.2. The aircraft as recited in wherein the cargo hook module is coupled between the first and second pylons.3. The aircraft as recited in wherein the cargo hook module further comprises a fixed cargo hook.4. The aircraft as recited in wherein the cargo hook module further comprises a winch mounted cargo hook.5. The ...

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Номер записи: 87
30-01-2020 дата публикации

Burnthrough Resistant Floor Panels and Method of Manufacturing and Installation

Номер: US20200031450A1
Автор: Slaton Daniel B.
Принадлежит:

Aspects of the present disclosure present floor panels and floor panel assemblies that include incorporated burnthrough resistant materials and/or include burnthrough resistant layers applied to floor panels used in floor panel assemblies, along with methods for inhibiting fire penetration into compartments and methods of making and installing burnthrough resistant floor panels and floor assemblies including such apparatuses and methods as they relate to vehicles, including aircraft. 1. A burnthrough resistant floor panel comprising:an internal support structure having an internal support structure first surface and an internal support structure second surface;at least one skin configured to substantially cover the internal support structure first surface and the internal support structure second surface to form a floor panel, said at least one skin having an exterior skin surface; anda burnthrough resistant layer proximate to the at least one exterior skin surface.2. The floor panel of claim 1 , wherein the internal support structure comprises a honeycomb panel.3. The floor panel of claim 1 , wherein the floor panel comprises:{'sup': '2', 'the burnthrough resistant layer proximate to at least one of the internal support structure first surface and the internal support structure second surface, said burnthrough resistant layer resisting a backside heat flux of less than about 2.0 BTU/ft/sec. for at least from about 4 mins. to about 5 mins.'}4. The floor panel of claim 1 , wherein the burnthrough resistant layer comprises a coating claim 1 , said coating configured to be applied to at least one of the internal support structure first surface and the internal support structure second surface.5. The floor panel of claim 1 , wherein the burnthrough resistant layer comprises at least one of: a spray coating claim 1 , a film claim 1 , and a tape.6. The floor panel of claim 1 , wherein the burnthrough resistant layer comprises a ceramic paper.7. The floor panel of claim 1 ...

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Номер записи: 88
04-02-2021 дата публикации

Frame component and method for producing a frame component, frame and fuselage structure for an aircraft

Номер: US20210031897A1
Автор: Bernhard Hörger, Carsten Paul, Cord Haack
Принадлежит: PREMIUM AEROTEC GMBH

A frame component for a frame of a fuselage structure of an aircraft includes a central web extending along a longitudinal direction and having an inner edge region with respect to a radial direction running transversely with respect to the longitudinal direction and an outer edge region with respect to the radial direction. An inner web is bent from the inner edge region of the central web towards a first side. An outer web is bent from the outer edge region of the central web towards the first side. The central web, the outer web and the inner web are produced integrally from a metal sheet and together define a C-shaped cross section of the frame component. At least one stringer recess is formed in the outer web and in the outer edge region of the central web. The central web has, in the region of the stringer recess, a first reinforcing formation which forms a protrusion on the first side of the central web.

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Номер записи: 89
04-02-2021 дата публикации

Hybrid body fuselage

Номер: US20210031899A1
Автор: Andrew W. Sklar
Принадлежит: Individual

A supersonic aircraft fuselage includes a fuselage body having a first end, a second end, a length extending between the first end and second end, a surface, a first flat plane extending from the first end to a center of the fuselage body along the length on the surface, and a second flat plane extending from the second end to the center of the fuselage body along the length on the surface. The surface includes a curved portion conforming to a Sears-Haack body shape and abutting the first flat plane and second flat plane and extending between the first end and second end. A supersonic aircraft includes a first fuselage, a second fuselage, and a space between the first fuselage and second fuselage. The first fuselage and second fuselage form a Busemann biplane geometry within the space.

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Номер записи: 90
04-02-2021 дата публикации

NOSE STRUCTURE FOR AN AIRCRAFT AND METHOD OF MAKING AN AIRCRAFT

Номер: US20210031904A1
Автор: Murray Robert W., Skilton Richard A.
Принадлежит: The Boeing Company

An aircraft includes an airframe, forming a nose structure of the aircraft, and at least one high-level system. The aircraft also includes a wheel well assembly, coupled to the airframe and forming a portion of a nose landing gear bay. The wheel well assembly includes a pressure deck that extends from a right side of the airframe to a left side of the airframe and that forms a portion of a pressure boundary delimiting a pressurized space and a non-pressurized space. The aircraft further includes a floor-panel support, supported by the pressure deck. The aircraft also includes a plurality of transport elements, located between the floor-panel support and the pressure deck. 1. An aircraft comprising:an airframe, forming a nose structure of the aircraft;at least one high-level system;a wheel well assembly, coupled to the airframe and forming a portion of a nose landing gear bay, the wheel well assembly comprising a pressure deck that extends from a right side of the airframe to a left side of the airframe and that forms a portion of a pressure boundary delimiting a pressurized space and a non-pressurized space;a floor-panel support, supported by the pressure deck; anda plurality of transport elements, located between the floor-panel support and the pressure deck; and wherein:the pressure deck and the floor-panel support form a portion of a floor of the aircraft that delimits a flight deck, arranged over the floor in the pressurized space, and the nose landing gear bay, arranged under the floor in the non-pressurized space;the plurality of transport elements are in communication with the at least one high-level system; andthe plurality of transport elements is accessible from within the flight deck.2. The aircraft of claim 1 , wherein:the plurality of transport elements is coupled to the floor-panel support to form a subfloor assembly; andthe subfloor assembly is coupled to the pressure deck within the airframe.3. The aircraft of claim 1 , further comprising a plurality ...

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Номер записи: 91
04-02-2021 дата публикации

LIFT ROTOR SYSTEM

Номер: US20210031909A1
Автор: PACHIDIS Vasileios, SAIAS Chana A.
Принадлежит: ROLLS-ROYCE PLC

A lift rotor arrangement () for a VTOL aircraft (). The lift rotor arrangement () comprises: a fairing () mounted on a wing segment (); and first and second rotor blades () mounted on a first shaft () extending vertically from the fairing (). The first shaft () is movable between an extended position in which the first and second rotor blades () are vertically spaced above the wing segment () and are rotatable to provide vertical lift, and a retracted position in which the first and second rotor blades () are rotationally-fixed with the first rotor blade () stowed within the wing segment (). The blades () may be rotatable around an axis substantially perpendicular to the axis of the respective first shaft () so as to act as ailerons/elevons in the retracted position. 1. A lift rotor arrangement for a VTOL aircraft , the lift rotor arrangement comprising: a fairing mounted on a wing segment; and first and second rotor blades mounted on a first shaft extending vertically from the fairing , wherein the first shaft is movable between an extended position in which the first and second rotor blades are vertically spaced above the wing segment and are rotatable to provide vertical lift , and a retracted position in which the first and second rotor blades are rotationally-fixed with the first rotor blade stowed within the wing segment.2. The lift rotor arrangement according to wherein claim 1 , in the retracted position claim 1 , the first rotor blade is stowed within a recess or a void in the wing segment.3. The lift rotor arrangement of wherein the first and second rotor blades are aerofoils with a respective leading edge and trailing edge and wherein the trailing edge of the first rotor blade is aligned with a rearward edge of the wing segment in the retracted position.4. The lift rotor arrangement according to further comprising third and fourth rotor blades mounted on a second shaft extending vertically from the fairing in an opposing vertical direction to the first ...

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Номер записи: 92
19-02-2015 дата публикации

TWO-DIMENSIONAL MORPHING STRUCTURE FOR WING

Номер: US20150047337A1
Автор: Gandhi Umesh N.
Принадлежит: Toyota Motor Engineering & Manufacturing North America, Inc

An apparatus and methods for changing the shape of a wing using a plurality of morphing structures. One example method includes coupling a plurality of morphing structures to each other. Each morphing structure includes an anchor, a plurality of hinges, a plurality of shape-memory alloy members wherein each shape-memory alloy member extends from the anchor to a different hinge, a plurality of springs wherein each spring extends from the anchor to a different hinge, and a plurality of rigid members wherein each rigid member extends between two hinges. The method further includes actuating the plurality of shape-memory alloy members in at least some of the morphing structures wherein the shape-memory alloy members contract when actuated to pull against the hinges and anchors within the actuated morphing structures to rotate the rigid members and change the shape of the actuated morphing structure. 1. A morphing structure , comprising:an anchor;a plurality of hinges;a plurality of shape-memory alloy members wherein each shape-memory alloy member extends from the anchor to a different hinge;a plurality of springs wherein each spring extends from the anchor to a different hinge; anda plurality of rigid members wherein each rigid member extends between two hinges.2. The morphing structure of wherein the anchor includes a control structure for actuating the plurality of shape-memory alloy members.3. The morphing structure of wherein the plurality of springs counteract the forces applied to the anchor and the plurality of hinges when one or more of the plurality of shape-memory alloy members are actuated.4. The morphing structure of wherein the plurality of hinges surround the anchor.5. The morphing structure of wherein each hinge is coupled to one of a shape-memory alloy member and a spring.6. The morphing structure of wherein each hinge is coupled to two rigid members and one of a shape-memory alloy member and a spring.7. The morphing structure of wherein four rigid ...

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Номер записи: 93
16-02-2017 дата публикации

DECOMPRESSION PANEL FOR USE IN AN AIRCRAFT

Номер: US20170043856A1
Автор: Arm Hannah Rebecca, Glaze Tatum Jeremy Ryan, Loken Allan A., Olliges Michael J., Perkins Thomas
Принадлежит:

A method of installing and replacing a plurality of decompression panels into a sidewall assembly of an aircraft includes coupling a first decompression panel to at least one of a sidewall and a floor panel of the sidewall assembly. The method also includes inserting a tab of the first decompression panel into a slot of a second decompression panel. The slot is defined by a flange extending along a surface of the second decompression panel. The method further includes coupling the second decompression panel to at least one of a sidewall and a floor panel of the sidewall assembly. 120-. (canceled)21. A method of installing and replacing a plurality of decompression panels into a sidewall assembly of an aircraft , said method comprising:coupling a first decompression panel to at least one of a sidewall and a floor panel of the sidewall assembly;inserting a tab of the first decompression panel into a slot of a second decompression panel, wherein the slot is defined by a flange extending along a surface of the second decompression panel; andcoupling the second decompression panel to at least one of a sidewall and a floor panel of the sidewall assembly.22. The method in accordance with claim 21 , wherein inserting the tab comprises deforming the flange to enable the slot to receive the tab.23. The method in accordance with claim 22 , wherein deforming the flange comprises deforming the flange along a channel formed through the flange claim 22 , wherein the channel facilitates deformation of the flange.24. The method in accordance with claim 21 , further comprising releasing the flange to secure the tab within the slot.25. The method in accordance with claim 21 , further comprising coupling a clamping device to the first decompression panel to secure the first decompression panel to the second decompression panel.26. The method in accordance with claim 25 , wherein coupling the clamping device comprises coupling the clamping device to the first decompression panel and at ...

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Номер записи: 94
15-02-2018 дата публикации

ROTARY WING AIRCRAFT WITH A FUSELAGE THAT COMPRISES AT LEAST ONE STRUCTURAL STIFFENED PANEL

Номер: US20180043982A1
Автор: FIETZ Helmut, FINK Axel, RACK Andreas
Принадлежит: AIRBUS HELICOPTERS DEUTSCHLAND GMBH

A rotary wing aircraft with a fuselage that comprises at least one structural stiffened panel, the structural stiffened panel comprising a stressed skin and a stiffening framework that is rigidly attached to the stressed skin, wherein the stressed skin comprises an inner skin, an outer skin and a core element assembly that is arranged between the inner skin and the outer skin, the core element assembly comprising at least one viscoelastic core element and at least one intermediate core element that are tessellated, wherein the at least one viscoelastic core element is provided for noise and vibration damping. 1. A rotary wing aircraft with a fuselage that comprises at least one structural stiffened panel , the structural stiffened panel comprising a stressed skin and a stiffening framework that is rigidly attached to the stressed skin , wherein the stressed skin comprises an inner skin , an outer skin and a core element assembly that is arranged between the inner skin and the outer skin , the core element assembly comprising at least one viscoelastic core element and at least one intermediate core element that are tessellated , wherein the at least one viscoelastic core element is provided for noise and vibration damping.2. The rotary wing aircraft of claim 1 ,wherein the at least one intermediate core element is arranged in an attachment area between the inner skin and the outer skin, the inner skin being rigidly attached via the at least one intermediate core element to the outer skin in the attachment area.3. The rotary wing aircraft of claim 2 ,wherein the at least one intermediate core element comprises a grid of longitudinal components and transversal components that define at least one intermediate free space, the at least one viscoelastic core element being arranged in the at least one intermediate free space.4. The rotary wing aircraft of claim 3 ,wherein the stiffening framework comprises a grid of stringers and frames that are rigidly attached to the ...

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Номер записи: 95
19-02-2015 дата публикации

EFFICIENT CONTROL AND STALL PREVENTION IN ADVANCED CONFIGURATION AIRCRAFT

Номер: US20150048215A1
Автор: McGinnis John William
Принадлежит:

An apparatus forming an aircraft which is designed for flight by movement through the air, the aircraft has a front and rear portions and a center of mass, with left and right sides when divided by a central plane of reference. The aircraft has inboard portions closer to said central plane of reference and outboard portions farther from said central plane of reference. Further, the aircraft contains at least one positive lifting aerodynamic surface configured to affect the flow of air near said at least one positive lifting aerodynamic surface when said aircraft is appropriately moving forward, and at least one elevon structure configured to create negative aerodynamic force when said aircraft is appropriately moving forward. The elevon structure is constructed so as to have outboard portions thereof positioned outward of said central plane of reference to a distance at least three-fourths of the distance from said central plane of reference to a tip end of said at least one wing. 1. An apparatus forming an aircraft which is designed for flight by movement through the air , said aircraft having front and rear portions and a center of mass , said aircraft having left and right sides when divided by a central plane of reference , said aircraft having thereby inboard portions closer to said central plane of reference and outboard portions farther from said central plane of reference , comprising:at least one aerodynamic lifting surface configured to affect the flow of air near said at least one aerodynamic lifting surface when said aircraft is appropriately moving forward, said at least one aerodynamic lifting surface thereby configured to create positive lift when said aircraft is appropriately moving forward, said at least one aerodynamic lifting surface thereby forming at least one wing, said at least one wing having a center of lift which is rearward of said center of mass of said aircraft in flight;at least one airfoil structure configured to create aerodynamic ...

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Номер записи: 96
14-02-2019 дата публикации

Stringer stiffened composite panels having improved pull-off strength

Номер: US20190047676A1
Автор: Forouzan Behzadpour, Jiangtian Cheng
Принадлежит: Boeing Co

Stringer stiffened composite panels having improved pull-off strength are disclosed. An example stringer includes a first surface, a second surface, an edge, and a chamfer. The second surface is located opposite the first surface and is to be coupled to a composite structure of an aircraft. The edge extends from the second surface toward the first surface. The chamfer extends from the first surface to the edge at an angle between twelve and eighteen degrees relative to the first surface.

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Номер записи: 97
14-02-2019 дата публикации

PANEL ASSEMBLY

Номер: US20190047679A1
Автор: JAISWAL Vimal, WOOLCOCK John
Принадлежит:

A panel assembly with a panel and a stringer is disclosed. The stringer has a stringer foot and an upstanding stringer web. The stringer foot has a flange which extends in a widthwise direction between the stringer web and a lateral edge and in a lengthwise direction alongside the stringer web, and a foot run-out which extends between the flange and a tip of the stringer foot. The foot run-out is bonded to the panel at a foot run-out interface. Reinforcement elements, such as tufts, pass through the foot run-out interface into the panel. The reinforcement elements are distributed across the foot run-out interface in a series of rows including an end row nearest to the tip of the stringer foot and further rows spaced progressively further from the tip of the stringer foot. At least the end row has three or more of the reinforcement elements which are distributed along a polygonal curve. 1. A panel assembly comprising:a panel;a stringer comprising a stringer foot and an upstanding stringer web, wherein the stringer foot comprises a flange which extends in a widthwise direction between the stringer web and a lateral edge and in a lengthwise direction alongside the stringer web, and a foot run-out which extends between the flange and a tip of the stringer foot, wherein the foot run-out is bonded to the panel at a foot run-out interface; andreinforcement elements which pass through the foot run-out interface,wherein the reinforcement elements are distributed across the foot run-out interface in a series of rows including an end row nearest to the tip of the stringer foot and further rows spaced progressively further from the tip of the stringer foot, and at least the end row comprises three or more of the reinforcement elements which are distributed along a polygonal curve.2. The panel assembly of wherein the reinforcement elements are bonded to the foot run-out and/or the panel.3. The panel assembly of wherein at least the end row comprises four claim 1 , five claim 1 , ...

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Номер записи: 98
13-02-2020 дата публикации

METHODS AND APPARATUS FOR A DISTRIBUTED AIRCRAFT ACTUATION SYSTEM

Номер: US20200047873A1
Автор: Huynh Neal Van, McCormick Patrick Joseph, Moser Matthew Alexander
Принадлежит:

Methods, apparatus, and articles of manufacture for a distributed aircraft actuation system are disclosed. An example apparatus includes a non-responsive component detector to determine a position difference between a first position of a first control surface of an aircraft and a second position of a second control surface of the aircraft, the first position lagging the second position, and a command generator, in response to determining that the position difference satisfies a threshold, the command generator is to cease movement of the second control surface at the second position, attempt to move the first control surface to the second position, and cease movement of the first control surface when moved to the second position. 1. An apparatus comprising:a non-responsive component detector to determine a position difference between a first position of a first control surface of an aircraft and a second position of a second control surface of the aircraft, the first position lagging the second position; and cease movement of the second control surface at the second position;', 'attempt to move the first control surface to the second position; and', 'cease movement of the first control surface when moved to the second position., 'a command generator, in response to determining that the position difference satisfies a threshold, the command generator is to2. The apparatus of claim 1 , wherein the command generator is to claim 1 , in response to the first control surface being unable to move to the second position claim 1 , deactivate a first set of control surfaces while a second set of the control surfaces remains active claim 1 , the first set of the control surfaces including the first control surface on a first side of the aircraft and the second control surface on a second side of the aircraft claim 1 , the second set different from the first set claim 1 , the deactivating including:ceasing movement of the first control surface at the first position;moving the ...

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Номер записи: 99
13-02-2020 дата публикации

SYSTEM AND METHOD OF PROTECTION OF AIRCRAFT FROM FOREIGN OBJECT STRIKES

Номер: US20200047915A1
Автор: Marimuthu Anand Kumar, Smith Michael, Tho Cheng-Ho
Принадлежит:

An aircraft structure protected from collisions with foreign objects includes a first support proximate to an aircraft structure root; first sheets attached to the first support to form a first leading edge portion of the aircraft structure; a second support proximate to an outer end of the aircraft structure; second sheets attached to the second support to form a second leading edge portion; and a door including one or more ribs disposed between the first and second supports and configured to give shape to a third leading edge portion of the aircraft structure; and third sheets, attached to the ribs to form the third portion of the leading edge; wherein the first support, the first sheets, the second support, the second sheets, the one or more ribs, and the third sheets are disposed to mitigate damage from a foreign object collision by absorbing kinetic energy from a collision. 1. An aircraft structure protected from collisions with foreign objects comprising:a first support proximate to a root of the aircraft structure;one or more first sheets comprising a first material and attached to the first support to form a first portion of a leading edge of the aircraft structure;a second support proximate to an outer end of the aircraft structure;one or more second sheets comprising a second material and attached to the second support to form a second portion of the leading edge; and one or more ribs disposed between the first support and the second support and configured to give shape to a third portion of a leading edge of the aircraft structure; and', 'one or more third sheets comprising a third material, attached to the ribs to form the third portion of the leading edge;, 'a door comprisingwherein the first support, the one or more first sheets, the second support, the one or more second sheets, the one or more ribs, and the one or more third sheets are disposed to mitigate damage from a collision with a foreign object by absorbing kinetic energy from a collision.2. ...

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Номер записи: 100