Abstract: Systems and methods are provided for installing splice fittings. One embodiment is a method for forming a splice between stringers. The method includes aligning a web of a splice fitting with a web of a first stringer and a web of a second stringer, and coupling the first stringer and the second stringer by installing fasteners that secure the web of the first stringer to the web of the splice fitting, and by installing fasteners that secure the web of the second stringer to the web of the splice fitting.

Abstract: A duct stringer is disclosed including a structural member with a hat-shaped cross-section. The structural member has a crown, a pair of webs and a pair of feet. A channel member with a U-shaped cross-section has a base and a pair of flanges. The flanges of the channel member are co-cured to opposed inner faces of the webs of the structural member. The structural member and the channel member together provide a duct with a closed cross-section which is adapted to transport fluid, for instance in an aircraft wing to provide a vent function in an aircraft fuel system.

Abstract: Adhesion of a metal surfaces (e.g. titanium and titanium alloys) to an epoxy resin is improved by creating a sol-gel layer at the metal/epoxy resin surface, with the sol-gel comprising a mixture of organometallic compounds to react with or bond to both the metal surfaces and an interfacing epoxy resin.

Abstract: A fastener system for ignition prevention triggered by a lightning strike. The fastener has a fastener body and a fastener head, and a fastening member configured to be mounted on a portion of the fastener body, wherein the fastener and the fastening member are configured to fasten at least a first part and a second part between the fastener head and the fastening member, and the fastener comprises an electrically insulating material. A structure comprises the fastener system, the first part that faces an outside from where the lightning strike may hit the first part and the second part.

Abstract: The present disclosure is directed to a composite component defining a component aperture extending between a first surface and a second surface. The composite component includes an insert having an insert annular wall positioned in the component aperture. The insert annular wall defines an insert aperture therethrough. An insert flange extends radially outwardly from the insert annular wall and contacts the first surface of the composite component. The insert flange includes a diameter about 1.5 times to about 5 times greater than a smallest diameter of the component aperture defined by the composite component.

Abstract: The disclosure herein provides for disbond resistant stringer runouts. Various aspects may be used independently or in combination to mitigate disbonding associated with the stiffener runout under operational loads. Aspects include a stiffener runout having rounded base flange corners. A recessed notch may be incorporated within the base flange to provide a flange termination point that is forward of the web termination point, allowing the web to terminate prior to the base flange. The web may be trimmed from a full height to a reduced height at the web termination point. The base flange may be co-bonded to the underlying composite structure via a scarf joint. Perimeter clamp radius fillers may be used to concentrate a clamping force around the perimeter of the base flange.

Abstract: Disclosed is a multi-spar torsion box structure comprising a plurality of spars of composite material arranged to form a multi-cell structure with two or more cells extending longitudinally along the torsion box, and upper and lower skin covers formed of composite material and joined to upper and lower surfaces of the multi-cell structure, respectively. The structure further comprises at least one belt-like reinforcing element extending around the outer perimeter of the torsion box, and fixed to the upper and lower surfaces of the multi-cell structure, and transversely arranged with respect to the longitudinal direction of the torsion box. The belt-like reinforcing element is made of a metallic or a composite material. The disclosed torsion box can advantageously be used in the manufacture of aircraft lifting surfaces, such as horizontal tailplanes (HTP) or wings.

Abstract: A device and method 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.

Abstract: In the present invention, a through-hole (2) and a compression periphery part (3a) at which the core of the resin foam (1c) is compressed are provided in a fastening part of a resin-foam core composite plate (1). A cylindrical part (11a) of a metallic fastening member (11) is inserted into the through-hole (2) from the proximal end side of a wall part (4). The fastening member (11) is crimped to the distal end of the wall part (4) via an eyelet member (12) placed on the outside of the wall part (4) so as to forcibly spread the cylindrical part (11a). The fastening member (11) and a counterpart member (A) are fastened together by tightening a nut (14) onto a bolt (13) inserted into the cylindrical part (11a).

Abstract: A system for connecting spaced components of a gas turbine engine includes connecting one end of a T-bar to each of the spaced components. One component is attached to the crossbar portion of the T-bar by trapping the crossbar between the component and a bracket. The second component is attached to the shaft portion using a hanger nut connected to the shaft portion of the T-bar and bolted to the second component. The system may be installed by attaching the crossbar portion, then advancing the first component towards the second component until the shaft portion passes through an aperture in the second component. The hanger nut is attached to the shaft portion and the second component.

Abstract: A wind blade is provided. The wind blade includes a self-supporting structural framework, having multiple chord-wise members and one or more span-wise members. Each of the multiple chord-wise members and the one or more span-wise members have an aerodynamic contour. The wind blade also comprises a fabric skin located over the self-supporting structural framework. Further, the wind blade includes a tensioning mechanism configured for providing pretensioning in the fabric skin. The tensioning mechanism includes multiple mechanical force elements coupled with the fabric skin for providing a predetermined force for maintaining an aerodynamic surface of the fabric skin during operation of the wind blade.

Abstract: The disclosure disclosed a connection device for connecting first means of the airplane to second means of the airplane. The first means and the second means are provided with connection regions, which lengthwise extend in a first direction and are connected with each other, respectively. A plurality of finger-like grooves, which are parallel with each other and extend respectively in a second direction inclined relative to the first direction in which the connection region extends, are provided respectively on a connection face of the connection region of the first means and a connection face of the connection region of the second means.

Abstract: An aircraft wing includes a rib attached by a joint to a tubular spar that varies in cross sectional shape during flight. The joint includes a ring attached to the tubular spar, and pivot pins connecting the ring with the rib. The pivot pins allow relative movement between the rib and the tubular spar as the tubular spar varies in cross sectional shape.

Abstract: A symmetrical rib structure and methods are presented. A structural web comprises a web center plane and a plurality of web stiffeners, and is symmetrical about the web center plane. A flange is configured on a periphery of the symmetrical rib structure, and a fastener hole is configured in the flange and comprises a fastener hole centerline aligned with the web center plane. A cutout is configured in the structural web around the fastener hole, and two support columns are coupled to the flange on opposite sides of the cutout.

Abstract: An assembly comprising first and second structural components and a packer disposed between the first and second components, wherein the first component and the packer have mating part-spherical surfaces, and the first and second components are fixed relative to one another. Also, a method of aligning first and second structural components relative to one another to form an assembly, the method comprising bringing together first and second structural components with a packer disposed between the components, wherein the first component and the packer have mating part-spherical surfaces, sliding adjustment of the first component relative to the packer until the first and second components are aligned in a desired relationship, and then fixing the first and second components relative to one another.

Abstract: The invention relates to a method of connecting two fuselage sections (1, 2, 21, 22) formed by a fiber-reinforced plastics material, in particular wound fuselage sections (1, 2, 21, 22) for aircraft fuselage cells, having a multiplicity of stringers (10, 11, 35, 36) arranged with a uniform distribution at a distance from and parallel to one another on the inner side of an outer skin (3, 4, 25, 26), with the formation of a transverse joint (5, 30).

Abstract: A flow body includes a flow body base body with a first shell element, a leading flow body edge with a second shell element and a clamping body, and a receiving space partially delimited by the clamping body. The flow body base body includes on a front end a projection, having a shape corresponding to that of the receiving space for engaging into the receiving space. The clamping body is internally arranged on the leading flow body edge spaced apart from a rear end of the second shell element. The rear end of the second shell element is flushly positioned on the first shell element when the projection engages into the receiving space. A smooth and harmonious transition that does not influence a laminar flow around the flow body can be achieved between the two components.

Abstract: A wing assembly for an aircraft is disclosed having a spar, a wing cover, and a rib. The spar has a web portion and a flange portion joined by a radius corner, wherein a gap is formed at an interface between the spar, the wing cover and the rib. The assembly further includes a seal member which substantially covers the gap, and sealant provided between the seal member and the rib and/or the wing cover and/or the spar so as to seal the gap.

Abstract: 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.

Abstract: A system and method that allow inspection of hollow structures made of composite material, such as an integrally stiffened wing box of an aircraft. A wing box comprises top and bottom skins connected by a plurality of spaced spars. The system employs a plurality of scanners for inspecting different portions of each spar. The system uses dynamically controlled magnetic coupling to connect an external drive tractor to computer-controlled scanners that carry respective sensors, e.g., linear ultrasonic transducer arrays. A system operator can control the various components by means of a graphical user interface comprising multiple interaction regions that represent the individual scanner motion paths and are associated with respective motion script files.

Abstract: A fabrication method for providing a relatively smooth surface associated with lightning strike protection for composite materials joined with fasteners recessed below the surface is described. The method includes coating at least one of a head of the fastener and a preformed dielectric insert with an adhesive, inserting the preformed dielectric insert into the recess such that the adhesive engages both contours of the recess and the preformed dielectric insert, and removing any excess adhesive from the surface area of the composite material and a top surface of the insert.

Abstract: A fairing device for creating an aerodynamic cover for an area of an aircraft structure, including a first fairing part with an outer side forming a first exterior flow contour, and having a fairing section for contacting a first section of the aircraft structure and a connecting section. The first fairing part includes a mounting device for receiving a connecting element. The fairing device includes a second fairing part including a fairing section having an outer side forming a second exterior flow contour and is designed for contacting a second section of the aircraft structure and a connecting section. The shape of the outer side of the second fairing part is adapted to the inner side of the connecting section of the first fairing part in an overlapping area, so the connecting section of the first fairing part forms a joint piece protruding over the second fairing part in the overlapping area.

Abstract: Technologies for reducing the risk of disbonding between two bonded members at areas of differing strain in the members are provided. Where the difference in strain between the two bonded members becomes excessing, movement of the members relative to each other may cause a traditional rigid adhesive to fail, resulting in a disbond or delamination in one or both composite members at the point of differing strain. Disbonding may be minimized or prevented by placing a compliant material, such as a thin sheet of rubber, between the structural members at the point of differing strain. The compliance of the material may allow the compliant material to deform enough to remain bonded to both structural members when subjected to a stress that might otherwise disbond or delaminate the bonded members.

Abstract: Structural panel assemblies include first and second composite panels positioned in edgewise alignment, with at least one tab extending from a mating edge of the first panel and including an exterior surface in facing registration with a corresponding portion of an interior surface of the second composite panel. The corresponding portion includes a recessed area to interfit with the tab or is otherwise configured to receive the tab. The exterior surface of the tab is spaced from the corresponding portion, to receive a bonding material therebetween for bonding the composite panels.

Abstract: A method for joining, by adhesive bonding, at least two large joining partners in vehicle construction, in particular in aircraft construction, comprises the following steps: detecting the geometric data of the joining partners in an automated manner, detecting the joint gap dimensions of the joining partners from the geometric data, applying adhesive to one or both joint faces of the two joining partners to be joined as a function of the joint gap dimensions, joining the joining partners in the joining position, and sequentially applying joining pressure to the joint faces along the joint gap to bring the joining partners into the final joining position.

Abstract: The disclosed shell structure splice and method may include a first panel having a first edge, a second panel having a second edge, the second edge being positioned in edgewise alignment with the first edge to form a splice joint, a strap bridging the splice joint and attached to the first panel and the second panel, the strap having a first tapered region and a second tapered region, a first fitting having a tapered section and a flat section, the tapered section being attached to the first tapered region of the strap, and a second fitting having a tapered section and a flat section, the tapered section being attached to the second tapered region of the strap.

Abstract: The subject matter disclosed herein discloses a system with attachment clips designed to fix an attachment element of an aircraft fuselage frame onto a fuselage skin and/or to stiffeners fitted onto this fuselage skin, the system also comprising tooling elements arranged alternately with the clips that they connect to each other, in order to reduce the fabrication time and cost of an aircraft fuselage.

Abstract: 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 ? to the fibers of the second internal ply. In some embodiments the angle ? is greater than 3° but less than 20°. In one embodiment the angle ? is between 35° and 42°.

Abstract: Structures and methods for joining composite fuselage sections and other panel assemblies together are disclosed herein. In one embodiment, a shell structure configured in accordance with the present disclosure includes a first panel portion positioned adjacent to a second panel portion. The first panel portion can include a first stiffener attached to a first composite skin, and the second panel portion can include a second stiffener attached to a second composite skin. The shell structure can include a fitting extending across a first edge region of the first panel portion and a second edge region of the second panel portion. A first end portion of the fitting can be attached to the first stiffener and the first composite skin, and a second end portion of the fitting can be attached to a second stiffener and a second composite skin, to join the first panel portion to the second panel portion.

Abstract: The present invention relates to the use of connecting elements for fastening CFK components and metal components in an aircraft or spacecraft, wherein a CFK component and a metal component can be connected to each other via at least one connecting element. The connecting element in this case is produced from an electrically non-conducting material, or has a shank and is produced from an electrically conducting or non-conducting material, wherein the shank is provided with a sleeve consisting of an electrically non-conducting material.

Abstract: 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.

Abstract: The invention relates to a radome (1) for an aircraft, and more specifically to a device for connecting said radome to the structure of said aircraft. The radome that is the subject of the invention comprises a plurality of locking units able to bring together the opposing surfaces of the aircraft's fuselage (10) and the radome (1), each unit comprising: locking means (20) able to exercise a traction force on the surface of the fuselage, substantially normal to the surface, via attachment means; centering means able to withstand the shear forces substantially tangential to the opposing surfaces of the fuselage and radome; the locking means (20) are placed such that they are located inside the cone of resulting forces on the centering means so as to minimize the bending stresses generated by the assembly in the radome (1).

Abstract: Methods for attaching a window to an airframe, methods of maintaining structural integrity of an airframe while attaching a window to the airframe, and portions of an aircraft are set forth herein. One method for attaching a window to an airframe comprises the steps of: applying to a bonding surface of a nutplate an adhesive; inserting through an airframe hole a fixture attached to a nut of the nutplate, and applying tension to the fixture to hold the adhesive-coated bonding surface of the nutplate adjacent an interior surface of the airframe; allowing the adhesive to cure; removing the fixture from the nut of the nutplate; and installing a bolt through a hole of a retainer, through the hole in the airframe, and into the nutplate; the retainer being in contact with and securing the window to the airframe. The adhesive is selected such that very specific conditions are met.

Abstract: A connector for securing floating double skin pipework to a structure includes means for securing to the structure, and an external body forming an outer channel and an internal tube forming an inner channel around the longitudinal axis of the connector. The means for securing to the structure includes a fixed connection backed up by a sliding connection along the longitudinal axis of the connector. The connector includes at least one bore for producing the fixed connection and at least one holding rail for producing the sliding connection.

Abstract: The invention seeks to improve the integrity in respect of tolerance to damage following the spread of cracks in failure-mitigating structures of the failsafe type. To do this, the invention proposes to form a composite hybrid structure within the failure-mitigating structure. A failure-mitigating mainframe, particularly of an aircraft fuselage, equipped with a structural element made of composite according to the invention, has profiled longitudinal structural components assembled back to back by fasteners. First sides are longitudinally aligned and bonded to the fuselage skin, and the other sides are longitudinally aligned some distance away from the first sides. Fasteners couple the composite structural element to the external face of the sides of the metal components that are longitudinally aligned some distance away from the first sides.

Abstract: Embodiments of reinforced composite structures for aircrafts and methods for making such reinforced composite structures are provided herein. A reinforced composite structure for an aircraft comprises a fiber reinforced composite stringer. The fiber reinforced composite stringer comprises a beam body portion and a runout portion that extends from the beam body portion in a first direction. The beam body and runout portions are defined by projection of a variable I-shaped cross-section in the first direction along at least a portion of an entire combined length of the beam body and runout portions. The variable I-shaped cross-section has a cap section, a foot section, and a web section that extends between the cap and foot sections. The variable I-shaped cross-section is configured such that a height of the web section tapers in the first direction along at least a portion of an entire length of the runout portion.

Abstract: The invention relates to a strut system for the stabilization of the shell (2) of an aerodynamic aircraft component in a commercial aircraft, comprising a pair of connection supports (4a, 4b) connected on the inside to the shell (2) and opposing each other, between which connection supports (4a, 4b) several struts (5) extend to form a latticework structure (3), which struts (5) in nodal points (6a, 6b) arranged on the ends are non-detachably connected to the connection supports (4a, 4b), wherein the area extension of the latticework structure (3) is aligned in such a manner relative to a middle plane of symmetry (11) that the struts (5), formed with an open symmetrical cross section, as well as the cross section of the connection supports (4a, 4b), are arranged so as to be symmetrical to the aforesaid.

Abstract: A fastening assembly for the fastening of systems in an aircraft comprising frames, a skin, and clips between the skin and the frames. The fastening assembly includes: ramps, that extend between two consecutive frames of the aircraft; brackets for fastening the ramps to two of the clips situated respectively on two consecutive frames of the aircraft; system supports, fastened to the ramps, each system support being able to receive at least one system.

Abstract: A lightning-protection fastener is provided that is capable of reliably preventing peeling off of an insulation layer during the operation of an aircraft and improving the anti-lightning-strike capability and reliability. There is a lightning-protection fastener that fastens a skin of an aircraft and a structural member positioned inside the skin, in which an insulation layer is melt adhered so as to cover one end surface of a head portion and also to mechanically engage with a fastener-side engagement portion (engagement portion) formed on the end surface.

Abstract: A composite frame includes a plurality of truss elements covered by a cap. The truss elements are flexibly coupled with each other and with the cap. The frame flexes to conform to a contoured surface to which it is attached. Once attached, the installed frame provides the required rigidity to the structure.

Abstract: A circumferential stiffener configured for a fuselage of an aircraft includes a web disposed at an angle to a normal axis of the circumferential stiffener, at least one inner flange, at least one outer flange and an integral angle profile configured for connection to a skin section of the fuselage.

Abstract: It is intended for installing and fitting, in conjunction with other iron fittings, a fairing on an aircraft in correspondence with a housing located at the junction of a stabilizer and a fuselage of an aircraft structure. Wherein each iron fitting comprises: a base plate (5) joined integrally to a torsion box (6) of the stabilizer (3); a mobile support (7) joined integrally to the fairing (1) and which is guided, over a linear movement, over the base plate (5); an adjustment screw (8) that is coupled into a threaded opening (15) of the base plate (5), which screw joins this base plate (5) to the mobile support (7), with the rotation of said adjustment screw (8) in combination with a retaining plate (23) dragging the respective mobile support (7) that is integral to the fairing (1).

Abstract: Aircraft and airframes include a skin operatively coupled to a structural reinforcement member. The structural reinforcement member and the skin collectively define a run-out region that encompasses a terminal edge of the structural reinforcement member. Within the run-out region, the skin is more flexible than in adjacent portions of the skin. Methods include defining a region of increased flexibility of a skin relative to adjacent portions of the skin, and operatively coupling the skin to a structural reinforcement member.

Abstract: A fuselage structure for an aircraft fuselage in composite material and an aircraft including such a fuselage structure. The fuselage structure provides savings in mass while allowing grounding of electric systems, and with which substantial mechanical stiffness characteristics may be guaranteed. In the fuselage structure attachment modules provide an assembling function as well as an electric function.

Abstract: A connecting arrangement for connecting two reinforcing elements of an aircraft or spacecraft, wherein the reinforcing elements have different cross-sectional profiles with at least one foot portion and at least one comb portion, at least one foot portion connecting element which on one side can be adapted to the geometric shape of the foot portion of the first reinforcing element and on the opposite side can be adapted to the geometric shape of the foot portion of the second reinforcing element and can be connected rigidly thereto; and at least one comb portion connecting element which on one side can be adapted to the geometric shape of the comb portion of the first reinforcing element and on the opposite side can be adapted to the geometric shape of the comb portion of the second reinforcing element and can be connected rigidly thereto.

Abstract: A method of manufacturing a structure, the method comprising: forming a panel assembly by bonding a stack of thickness control plies of composite material to a laminar composite panel, the stack of thickness control plies having a first edge proximate an edge of the laminar composite panel, a second edge opposite the first edge, and a ramp where the thickness of the stack of thickness control plies decreases towards the first or second edge; measuring the thickness of the laminate composite panel or the panel assembly; controlling the number of thickness control plies in accordance with the measured thickness; attaching a first component to the panel assembly, the first component having a surface which engages the laminar composite panel and a ramp which engages the ramp in the stack of thickness control plies; and attaching the laminar composite panel at or near its edge to a further component.

Abstract: A joint construction for an aeronautical structure made from composite material having an internal structural element and an adjoining skin element which define the aeronautical structure, which includes a coupling element disposed between and contiguous with the internal structural element and the adjoining skin element, wherein the coupling element has opposing adjoining diagonally configured end portions which form a wedge shaped diagonal joint therebetween which terminates at the internal structural element at one end and the skin element at the other end, and a plurality of rivets extending from the internal structured element to the adjoining skin, across the coupling element and across the diagonal joint.

Abstract: A line system arrangement in an aircraft or spacecraft having a fuselage includes at least one receiving region for receiving at least one component of a line system, wherein the at least one receiving region is accessible, at least in some sections, from outside the fuselage of the aircraft and spacecraft for installation and/or maintenance of the line system; and an aircraft or spacecraft with a corresponding arrangement.

Abstract: A coupling device for connecting fuselage sections, in each case forming a transverse abutting region for creating an aircraft fuselage comprising fuselage sections with a fuselage skin, a multitude of annular frame elements and stringers, wherein the coupling device has a rib-like design and includes: a strip-shaped transverse splicing plate for coupling the fuselage skins, a rib, which extends in the middle on the transverse splicing plate, for coupling an annular frame element, and a multitude of rib-like stringer coupling parts that extend on both sides of the transverse splicing plate, for coupling the stringers, as well as a method for producing the coupling device from a solid, rectangular, material trip of suitable thickness and length.

Abstract: Technologies for reducing the risk of disbonding between two bonded members at areas of differing strain in the members are provided. Where the difference in strain between the two bonded members becomes excessing, movement of the members relative to each other may cause a traditional rigid adhesive to fail, resulting in a disbond or delamination in one or both composite members at the point of differing strain. Disbonding may be minimized or prevented by placing a compliant material, such as a thin sheet of rubber, between the structural members at the point of differing strain. The compliance of the material may allow the compliant material to deform enough to remain bonded to both structural members when subjected to a stress that might otherwise disbond or delaminate the bonded members.