| Class / Patent application number | Description | Number of patent applications / Date published |
|---|
| 244133000 | Materials of construction | 26 |
| 20080258009 | Leading edge for aircraft made of reinforced composite material - The present invention relates to a leading edge ( | 10-23-2008 |
| 20080277531 | Hybrid Composite Panel Systems and Methods - Hybrid composite panel systems and methods are disclosed. In one embodiment, an assembly includes a primary section, a matrix member engaged with the primary section, and a secondary section engaged with the matrix member opposite the primary section. The primary section includes a plurality of first composite layers reinforced with a first reinforcing material, and the secondary section includes a plurality of second composite layers reinforced with a second reinforcing material. The primary and secondary sections are configured to bear an operating load at least partially transversely to the first and second composite layers, and are asymetrically configured such that the primary section bears a majority of the applied operating load. | 11-13-2008 |
| 20080283667 | HYBRID COMPOSITE-METAL AIRCRAFT LANDING GEAR AND ENGINE SUPPORT BEAMS - A hybrid composite-metal component is provided. The component includes an elongate inner metal piece, an outer metal piece disposed about at least a portion of the inner metal piece, and composite material disposed between the inner metal piece and the outer metal piece. The component may further include at least one of a seal and at least one fastener joining the inner metal piece and outer metal piece. Both the inner metal piece and the outer metal piece may include at least one tapered end. The tapered ends of both the inner metal piece and the outer metal piece each may include a double taper. | 11-20-2008 |
| 20080283668 | Composite material structure for aircraft fuselage and process for manufacturing it - The present invention provides a closed composite material structure for aircraft fuselage shaped on a male jig from which it can be separated in a certain direction, said structure comprising a single outer panel and a plurality of inner longitudinal stiffeners integrated in said panel, such that the expansion coefficient of the male jig is greater than the expansion coefficient of the composite material of the structure, thus being able to remove the already manufactured structure, formed by the panel and the integrated inner stiffeners, in a single operation. The present invention further provides a process for manufacturing such a closed structure. | 11-20-2008 |
| 20080308678 | APPLICATION OF INSULATING COATING - Systems and methods for discontinuously applying an insulating primer to a carbon fiber reinforced plastic (CFRP) component are disclosed. In one embodiment, a method for mitigating electrical surface discharges from a CFRP component includes first applying an insulating primer to a metallic component. Next, an insulating primer is applied discontinuously to the CFRP component adjacent the metallic component. The discontinuous application of the insulating primer forms a primed portions and unprimed portions. The unprimed portions are configured to enable electrostatic dissipation. | 12-18-2008 |
| 20090001219 | Fiber polymer matrix composites having silicon-containing inorganic-organic matrices and methods of making the same - Components made from solid fiber inorganic-organic-polymer matrix composites and methods of making the components and composites are provided. In one embodiment, and by way of example only, the method includes the steps of impregnating a fiber preform with a liquid resin comprising at least one compound selected from the group consisting of silsesquioxane and polysilazane, and heating the resin-impregnated preform at a temperature below the resin pyrolysis temperature to form the solid fiber inorganic-organic-polymer matrix composite. In other embodiments, a component, such as a mechanical or an aircraft component is provided that includes a solid fiber inorganic-organic-polymer matrix composite. The solid fiber inorganic-organic-polymer matrix composite may include a fiber preform impregnated with a resin comprising at least one silsesquioxane or a fiber preform impregnated with a resin comprising at least one polysilazane. | 01-01-2009 |
| 20090127393 | FOAM STIFFENED HOLLOW COMPOSITE STRINGER - A method and apparatus for a composite stringer. A method is used for manufacturing a hollow composite stringer. Foam is formed with a mandrel installed into the foam. A composite material and the foam is laid up onto a tool in a form of a stringer. The composite material and the foam in the form of the stringer is cured to form a cured stringer. The mandrel from the foam is removed to form the hollow composite stringer. | 05-21-2009 |
| 20090152400 | Acoustically attenuated fuselage for aircraft - The invention relates to a covering ( | 06-18-2009 |
| 20090184204 | DISTRIBUTION OF POINT LOADS IN HONEYCOMB PANELS - Composite panels used in aircraft are formed by a sandwich including a lightweight honeycomb core reinforced by a solid plank in medial regions of the panel. The plank is bonded to the honeycomb and forms an integral part of the core where a strut is attached to the panel. The plank distributes point loads imposed on the panel by the strut. | 07-23-2009 |
| 20090230247 | METHODS AND APPARATUS FOR INCORPORATING LUMINOPHORES INTO DECORATIVE LAMINATES - A process for incorporating light emitting materials into aircraft interiors is described. The method includes selecting at least one light emitting material, incorporating the selected at least one light emitting material into one or more of a laminate coating and an extruded component, and installing the one or more of a laminate coating and extruded component into an aircraft cabin configuration. | 09-17-2009 |
| 20100025535 | Protection of elements in composite material - Protection of elements ( | 02-04-2010 |
| 20100059629 | System and Method for Fabricating Composite Parts - The present application is directed to systems for controlling the direction of thermal expansion of a composite part during thermal processing. The systems may comprise at least one guiding mechanism capable of being attached to a surface of a tool for fabricating composites and at least one moveable composite insert capable of being attached to a composite part. The moveable composite insert has an opening for receiving the at least one guiding mechanism. The opening is capable of controlling movement of the at least one moveable composite insert in a desired direction during thermal expansion of the composite part, when the guiding mechanism is received by the opening. Methods for controlling the direction of thermal expansion of a composite part during thermal processing are also disclosed. | 03-11-2010 |
| 20100155537 | ENERGY-ABSORBING STRUCTURAL COMPOSITE ELEMENT - A structural, energy absorbing composite element for aircraft structures includes a metallic or nonmetallic skin shell with a porous filler material insert positioned within the skin shell. The porous filler material is a material having ligaments that collapse resulting in a densification of the porous filler material in response to impact loading or a compression force sufficient to cause failure of the combined assembly. | 06-24-2010 |
| 20100170996 | WELDABLE HIGH-STRENGTH ALUMINUM ALLOYS - An aluminum alloy comprises aluminum, magnesium, scandium, and an enhancing system. The magnesium is from about 0.5 percent to about 10.0 percent by weight based on the aluminum alloy. The scandium is from about 0.05 percent to about 10.0 percent by weight based on the aluminum alloy. The enhancing system is from about 0.05 percent to about 1.5 percent by weight based on the aluminum alloy. | 07-08-2010 |
| 20110062287 | Method of Joining a Thermoplastic Material to a Fibre Composite Material - The invention relates to a method of joining a thermoplastic material ( | 03-17-2011 |
| 20110315824 | COMPOSITE STRUCTURES HAVING INTEGRATED STIFFENERS AND METHOD OF MAKING THE SAME - A composite structure is made by placing a stiffener preform and a composite structure reinforcement on a tool. The preform and the reinforcement are vacuum bagged processed and co-infused with a thermoset resin to form a stiffened, unitized structure. | 12-29-2011 |
| 20140151507 | Split Resistant Composite Laminate - A composite laminate, method of forming same, and use for same are disclosed. One example of a composite laminate has multiple layers or plies ( | 06-05-2014 |
| 20150136910 | COMPOSITE MATERIAL SUITABLE FOR A MORPHING SKIN - A morphing skin for an air vehicle structure, the skin being formed by a composite material which includes: a multiplicity of fibres, a matrix incorporating the fibres, and a thermo-sensitive material, such as a thermo-plastic. The thermo-sensitive material can be reversibly transitionable in response to a change in temperature, between (i) an ambient temperature mode in which the thermo-sensitive material is capable of transferring loads in the composite material, and (ii) a heated mode in which the ability of the thermo-sensitive material to transfer the loads is reduced. This can allow the stiffness of the skin to be temporarily reduced to enable the skin to be morphed into a different shape. | 05-21-2015 |
| 20150298423 | Structural Inserts for Honeycomb Structures - A method and apparatus for manufacturing a sandwich structure. A cavity is formed in a core of the sandwich structure. A structural insert is placed into the cavity. The structural insert comprises a first number of parts configured to receive a load and a second number of parts assembled with the first number of parts to form the structural insert. | 10-22-2015 |
| 20150344125 | ENERGY ABSORPTION DEVICE FOR AIRCRAFT STRUCTURAL ELEMENT - A device for absorbing kinetic energy for an aircraft structural element undergoing a dynamic impact. The device includes an outer enclosure made from a braided composite material configured to preserve its integrity after an impact, a foam core, contained in the outer enclosure and to at least partially fill the outer enclosure. The foam core is configured to at least partially absorb the kinetic energy generated by the impact. Reinforcing elements are integrated at least partially into the foam core to dissipate, combined with the form core, the kinetic energy generated by the impact. The reinforcing elements includes discontinuous threads inserted into the foam core by stitching, and each discontinuous thread includes an L- or T-shaped head, folded down outside the outer enclosure. | 12-03-2015 |
| 20150375843 | Elongated Structures and Related Assemblies - An elongated structure comprising a web extending along a length of the elongated structure and a flange comprising a first flange portion, the first flange portion extending away from an area of the web. The first flange portion comprises a variable width along at least a portion of the length of the elongated structure. The first flange portion of the elongated structure may comprise the variable width the length of the elongate structure. The first flange portion may comprise a constant width along at least a portion of the length of the elongated structure. The elongated structure may further comprise a second flange portion. The second flange portion may comprise a variable width along at least a portion of the length of the elongate structure. The first flange portion may comprise a first top flange portion. | 12-31-2015 |
| 20160083824 | Methods and Systems for Shape Memory Alloy Structures - There is provided a method of training a shape memory alloy (SMA) workpiece. The method includes applying a force couple to a shape memory alloy (SMA) workpiece to impart a generally planar transformational behavior to the SMA workpiece to obtain a trained shape memory alloy (SMA) workpiece. | 03-24-2016 |
| 20160101576 | METHOD FOR MANUFACTURING A FIBRE COMPOSITE COMPONENT, FIBRE COMPOSITE COMPONENT, AND STRUCTURAL COMPONENT FOR AN AIRCRAFT OR SPACECRAFT - A method for manufacturing a fibre composite component is disclosed. A single-piece foam body or the parts of a multi-part foam body having at least one recess which is open towards a first side of the foam body are prepared and positioned. A rigidifying element for the backing structure or a preform or first semi-finished product for forming the rigidifying element is arranged in the recess at least in portions. A skin portion or a second semi-finished product for forming the skin portion is provided, and brought into contact, in regions, with the foam body on the first side thereof and with the rigidifying element. The skin portion or second semi-finished product and the rigidifying element or preform or first semi-finished product in contact therewith are processed further in such a way that a fibre composite component is obtained in which the skin portion and the rigidifying element are interconnected. | 04-14-2016 |
| 20160167763 | SANDWICH PANEL AND METHOD TO FORM THE PANEL | 06-16-2016 |
| 20160185442 | METHOD AND APPARATUS FOR REDUCING STRUCTURAL VIBRATION AND NOISE - A conjugate damper for a structural panel includes a constraining sheet extending between a first edge and a second edge. Each of the first edge and the second edge is at least partially coupled to a first surface of the structural panel. The conjugate damper also includes a damping layer coupled between the constraining sheet and the first surface such that, when the structural panel is in a compressively deformed state, a thickness of the damping layer in a direction generally normal to the first surface is decreased relative to a baseline state. The damping layer includes a viscoelastic material. | 06-30-2016 |
| 20160185451 | LANDING GEAR WELL ROOF - An aircraft landing gear well roof including a first reinforced main structure and a second reinforced main structure spaced from each other in a transverse direction of the roof and equipped with respective mounting for articulating a structural element of the first landing gear and mounting for articulating a structural element of the second landing bear. A membrane connecting the first reinforced main structure and the second reinforced main structure is disposed between them in the transverse direction. | 06-30-2016 |
