Airbus Operations Gmbh

Bae Systems, Inc.

Airbus Defence and Space Gmbh

Stephen Williams

Timothy J Sanderson

Andrew Paul Godbehere

Kevin Jones

Keith Travis

Ray James

Robert D Spear

Chetan Korya

Jonathan Price

Peter Baker

Paul S Stacey

An industrial control system includes an encryption device, a decryption device and a mission-critical communication link connecting the encryption device to the decryption device. The encryption device is configured to send encrypted messages to the decryption device over the mission-critical communication link. The system includes a failure monitor configured to monitor for a failure of the encryption device or the decryption device, and, in the event of failure, to continue the sending of the messages to the decryption device but without encrypting the messages.

Mission-critical communication links for industrial control systems

A communication system includes an encryption device, a decryption device and a communication link connecting the encryption device to the decryption device. The encryption device is configured to send encrypted messages to the decryption device over the communication link. The system includes a failure monitor configured to monitor for a failure of the encryption device or the decryption device, and, in the event of failure, to continue the sending of the messages to the decryption device but without encrypting the messages.

Communication links

A method of manufacturing an aircraft component, the component comprising a pair of covers joined by a pair of spars. The covers and spars form a closed box structure with four corners, each corner providing a transition between one of the covers and one of the spars. The method comprises: laying up a laminated sheet of fiber plies; and after the sheet has been laid up, folding the sheet to form the four corners and then joining the sheet to itself to provide the closed box structure.

Aircraft component with closed box structure

A structure comprising a space frame including a plurality of strut sections which cross at one or more locations within the space frame to define a node at each respective crossing location, wherein the strut sections are laminates comprising a plurality of ply layers, and wherein the ply layers of the crossing strut sections are interleaved at the node(s). Also a method of forming the structure. Also An aerofoil structure, comprising a space frame defining a cage adapted to receive upper and lower covers so as to form upper and lower aerodynamic surfaces of the aerofoil, and at least one spar extending generally spanwise across the structure, wherein the spar includes a spar web and upper and lower spar caps, and wherein the spar web is disposed on one of an interior or exterior side of the cage, and the spar caps are attached to the other of the interior or exterior side of the cage.

Space frame structure

A method for forming a laminate component, the method comprising the steps of: a) providing a tool having a molding surface; b) arranging a stack of ply layers on the molding surface of the tool; and c) molding the stack of ply layers over the molding surface in a molding process, wherein the tool has a wrinkle control feature in its molding surface for managing the development of wrinkle in the ply layers during the molding process, and wherein the wrinkle control feature cooperates with the stack of ply layers during the molding process to produce a local surface undulation in a nominal surface of the laminate component nearest the tool to either prevent wrinkling in the ply layers or initiate wrinkling in one or more of the ply layers at a predetermined location. Also, a tool having a wrinkle control feature, and a laminate component so formed.

Wrinkle control method and tool therefor

A composite structure has a first section, a second section, and a curved corner joining the first and second sections. The structure has a stack of fiber-reinforced plies including first and second external plies which run from the first section into the second section on an inside and an outside of the corner, respectively. A discontinuous first internal ply is sandwiched between the first and second external plies so that more of the first internal ply is located within the first section than within the second section. A second internal ply is sandwiched between the first and second external plies. The fibers of the first internal ply run at an angle &#x3b8; to the fibers of the second internal ply. In some embodiments the angle &#x3b8; is greater than 3&#xb0; but less than 20&#xb0;. In one embodiment the angle &#x3b8; is between 35&#xb0; and 42&#xb0;.

Composite structure

A composite structure has first and second sections, and a curved corner joining the first and second sections. The structure includes a stack of fiber-reinforced plies. First and second external plies run from the first section into the second section round an inside and an outside of the corner, respectively. First and second internal plies are sandwiched between the first and second external plies, and run from the first section into the second section round the corner. The first internal fly is dropped off inside the second section. The second internal ply is dropped off inside the first section. The structure has an increased thickness where the first and second internal plies overlap.

A method of stiffening a rib () during the assembly of an aircraft wing is described. A stiffener panel () is attached to a face of the rib (). The rib is aligned with one or more wing components with the stiffener panel () attached. An assembly operation such as drilling () is performed with the rib aligned and the stiffener panel attached. The stiffener panel () is removed from the rib after the assembly operation. The stiffener panel has a resilient external sealing ring which is mounted on a face of the panel and forms an external perimeter of a vacuum cavity; and at least one resilient internal sealing ring which is mounted on the panel () within the perimeter defined by the external sealing ring and forms an internal perimeter of the vacuum cavity. The stiffener can be attached to the rib by a vacuum clamping force, the vacuum being sealed by the external and internal sealing rings. The internal sealing ring(s) seal any holes which are present in the rib. The stiffener () can support the rib () during alignment to give it the required flatness.

Method of stiffening a rib

An apparatus and method for controlling the curing process during the manufacture of composite laminates and components. The apparatus including a temperature controlled vessel such as an autoclave in which the material is placed during curing and an infra-red temperature measuring device such as a camera located remotely from the component to measure the temperature of at least part of the material during curing.

Apparatus for curing a composite laminate

A method of reinforcing a laminated wing skin () for an aircraft. The method comprises laying-up a plurality of fibers to form the laminated wing skin (). A composite reinforcement member () is positioned on a surface () of the laminated skin () and one or more further layers are laid up so as to enclose the composite reinforcing member () at a position adjacent one surface of the laminated skin ().

Method of reinforcing a laminated member such as a skin for an aircraft

An aircraft wiring installation connecting a set of fuel side wires in an aircraft fuel tank to a set of airside wires outside the fuel tank, the installation comprising: a connector box having a fuel side wall () and an airside wall (); one or more connectors () carried by the fuel side wall () which connect the fuel side wires to the airside wires; a fairing () partially covering the airside wall (); and one or more holes () in the airside wall through which the airside wires pass from the connector box into the fairing ().

Aircraft wiring installation

The method of constructing a skin for a member of a body such as an aircraft comprises providing from a multi-axial fabric first and second pieces A, A shaped to correspond respectively to the shape of the member to be covered by the skin and to the shape of a similar member on the opposite side of the body, and inverting one of the pieces and placing it on the other. The method may include providing third and fourth pieces B, B from multi-axial fabric shaped to correspond respectively to the shape of the member to be covered by the skin and to the shape of the similar member on the opposite side of the body and inverting one of the third and fourth pieces B, B and placing it on the other to form a skin for the similar member. In order to minimise wastage the method preferably includes forming the first piece A and third piece B from areas of the fabric which lie alongside each other and/or forming the second piece A and the fourth B piece from areas of the fabric which lie alongside each other.

Method of constructing skin for a member of a body such as an aircraft

A method of constructing a composite material which comprises a laminate of plies comprising a warp dominated ply and a weft dominated ply, includes arranging first and second pieces of material alongside each other with a recess being defined at least partly by a warp dominated ply so that parts thereof form a lap joint which is received at least partly in a recess formed in or by a warp dominated or other ply, and arranging at least part of one of said first and second pieces of material in said recess so as to form lap joint with part of the other of the first and second pieces of material . The parts of said first and second pieces of material include weft dominated plies

Composite material construction

A method of forming a substantially planar, multi-layer, composite article that comprises a plastics materials reinforced with fibers that extend substantially continuously in one direction in the plane of each layer is provided. The majority of the layers are aligned in a main direction B and the article ( ) is disposed over a contoured surface of a forming tool ( ). The article ( ) is constrained against movement in a direction perpendicular to the plane of the main direction layers and forces applied to the article ( ) in the plain thereof in a direction substantially perpendicular to the main direction of the fibers. The article ( ) is thereby urged into conformity with the contoured surface of the tool ( ) while substantially maintaining the continuity of the fibers.

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Bristol, United Kingdom

Stephen Williams