Abstract: A multi-layer ceramic matrix composite structure (40) having a plurality of fiber-reinforced cooling passages (42) formed therein. The cooling passages are formed by the removal of a fugitive material (24). The fugitive material is part of a wrapped fugitive material structure (28) containing a layer of reinforcing ceramic fibers (26) that is used to lay-up the multi-layer structure. An intermediate layer of ceramic fabric 56 may be placed alternately over and under the wrapped fugitive material structure to separate the cooling passages into alternating upper (54) and lower (52) cooling passages. The transversely oriented fibers surrounding the cooling passages serve to increase the interlaminar strength of the structure when compared to prior art designs. An airfoil member (112) incorporating such reinforced integral cooling passages (120) is provided.

Abstract: A reinforced structural member having a tubular member and a support structure. The tubular member has a sidewall that defines a hollow interior portion. The support structure is formed from a composite material and bonded to the sidewall in at least two predetermined locations. The support structure is positioned within the tubular member and configured to transmit force between the predetermined locations so as to prevent the tubular member from being crushed in response to the application of a force of a predetermined magnitude to the tubular member. A method for forming a reinforced structural member is also provided.

Abstract: An aircraft composite assembly 10 is provided, including a first composite member 12, at least one first flange joint 26 comprising a first plurality of plies 24 cured to the first composite member 12, a second composite member 30, and at least one second flange joint 38 comprising a second plurality of plies 36 cured to the second composite member 30. The at least one second flange joint 38 is bonded to the at least one first flange joint 26 such that the first composite member 12 is affixed to the second composite member 30.

Abstract: In order to manufacture a sandwich panel using the RTM technique, a stack comprising a core (10) with open cells, a film (12) of intumescent material covering each of the faces of the core (10), a dry barrier fabric (14) covering each of the films (12) and an overlay of dry fibers (16) covering each of the barrier fabrics is placed in a mold. During polymerization of the films (12), pressurization of the mold and the presence of the barrier fabrics (14) prevent penetration of the foam into the overlays (16). The resin is then injected into the mold and then polymerized, without danger of penetration into the cells of the core (10).

Abstract: A ruddervator for an aerospacecraft including a monolithic, one-piece, oxide/oxide-based ceramic matrix composite (Oxide-CMC) shell having a hollowed interior area. A graphite composite structural member is inserted into the hollowed interior area and bonded thereto. The Oxide-CMC shell is comprised of one or more plies of Oxide-CMC fabric which are fused over a thick substrate of rigid ceramic foam insulation to form the monolithic shell. An outer mold line ply of the Oxide-CMC shell is further infused with a high-emissivity coating such as reaction-cured glass (RCG) to provide plasma heating re-radiation outward to reduce internal temperatures in the ruddervator. A torque box transition structure is secured to the graphite composite structural member to interface the ruddervator to a fuselage actuator spindle of the aerospacecraft. The ruddervator is reusable and formed from a reduced number of independent component parts, and weighs less than previously developed ruddervators.

Abstract: A movable structural component for a thermomechanically stressed assembly at least partially from fiber reinforced ceramic is disclosed, wherein the movable structural component comprises at least one structural element made by an polymer infiltration and pyrolysis process and at least one structural element made by a chemical vapor process.

Abstract: Integrally stiffened and formed, load carrying structures comprising a plurality of elongated thin-walled tubes placed co-extensively in a complementary side-by-side fashion which together form a hollow structure having a desired external contour. Integral skins forming the external and internal surfaces of the structure cooperatively therewith. The structure can be formed with an underlying internal support member spanning the interior of the load carrying structure, thereby connecting opposite sides of the structure together. Also, each of the tubes are wound with fibers in controlled orientations generally paralleling the direction of the loads applied to the tubes to optimize the strength to weight ratio of the tubes.

Abstract: A multi-layer tile material produced from layers of alumina enhanced thermal barrier material having different densities. The insulation layers are bound together by a high strength, high temperature alumina or silica binder having a coefficient of thermal expansion similar to that of the insulation layers. Use of the multi-layered tile allows the problems of tile slumping and of insufficient heat management associated with low density alumina enhanced thermal barrier tile to be overcome.

Abstract: The present invention provides an apparatus and method for varying a wall skin to alter airflow over the skin. The apparatus has a first and second end-plates, a plurality of flexible rods are arranged substantially parallel to substantially define a plane, each flexible rod has a mid-point and first and second ends secured to the first and second end-plates respectively. Each mounting structure is slidingly attached to the flexible rods between the mid-point and each end-plate, proximate to the end. An elastomer envelops the rods to form a flexible skin. A plurality of shape memory alloy rods are arranged parallel to and define a plane and having each end secured to an end-plate, the plurality of shape memory alloy rods being contractible when heated such that upon contraction the plurality of shape memory allow rods will buckle the plurality of flexible rods and the flexible skin to alter airflow over the skin.

Abstract: A subdermally-reinforced elastomeric transition is provided in which an elastomeric skin is attached to a plurality of subdermal supporting members that engage subdermal reinforcing members. The reinforcing members may be rods or support rails which the supporting members engage. The supporting members may be attached to the elastomeric skin in an orientation that is perpendicular to a direction of strain of the skin and may have a plurality of holes or slots for receiving the reinforcing members. Alternatively, the supporting members may be oriented to be parallel to the direction of strain of the skin, the reinforcing members being located substantially within the supporting members.

Abstract: The development of polymer composite liquid oxygen (LOX) tanks is a critical step in creating the next generation of launch vehicles. A composite LOX tank will weigh significantly less than conventional metal tanks. This benefit of reduced weight is critical to enable future launch vehicles to meet required mass fractions. The high strength and low weight of a composite tank allows for a lighter weight craft and/or higher payload capacity which results in lower costs per pound to place an object in orbit. The unique, nontraditional idea described here is to use resin-based composite materials to make such a composite tank or other oxygen carrying component such as a feedline. Polymer composites have traditionally been regarded as being more flammable than metals and therefore deemed not oxygen compatible. However, PEAR (polyether amide resin) composites, due to their low flammability and high toughness, could be made into liquid oxygen tanks.

Abstract: The missile comprises a nose, fixed fins or movable fins, gas rudders, propelling nozzles and blast pipe inserts, combustion chamber liners, tail cone, grid fins, fluid elements and radome or subcomponents of these made of carbon fiber-reinforced silicon carbide (C/SiC) and/or carbon fiber-reinforced carbon (C/C) and/or silicon carbide fiber-reinforced silicon carbide (SiC/SiC) which are integrated in a missile structure. The missile components may be made of C/SiC and/or C/C and/or SiC/SiC with continuous fiber reinforcement and/or chopped fiber reinforcement or combinations thereof. The missile components may be prepared by mechanical machining of C/SiC and/or C/C and/or SiC/SiC blanks either in a single piece or by co-infiltration with carbon or silicon or silicon carbide and/or co-siliconizing of separate C/SiC and/or C/C and/or SiC/SiC segments to result in a monolithic structure.

Abstract: A metallic structural component for an aircraft includes stiffening profile members integrally formed or connected onto a skin sheet. Each stiffening profile member includes a first part integrally connected to the skin sheet, and a second part non-integrally connected to the first part so as to form an internal boundary surface within the multi-part stiffening profile member. If a crack forms in the skin sheet and propagates into the first part of the stiffening profile member, the internal boundary surface resists the further propagation of the crack past this boundary surface. The crack propagation resistance is improved, the remaining strength of the structural component after formation of a crack is increased, and the overall structural component is light in weight. The outer contour of the stiffening profile member is not affected by the provision of the separate first and second parts or the internal boundary surface therebetween.

Abstract: Provided is a fabrication that is simple to make and cost effective in optimizing performance for moderately extended high velocity missions of high performance airborne vehicles. To counter rapid thermal loading while providing a relatively low cost structure, a laminate, in a preferred embodiment an intermediate modulus carbon fiber laminated as fabric layers in a resin matrix, is stitched through with glass fiber material. This laminate is then cured and formed as a section of skin of a supersonic missile, enabling extended supersonic operation, nominally some minutes after the laminate's Tg has been reached or exceeded. Additionally, an airframe and airframe members constructed in accordance with a preferred embodiment of the present invention are disclosed. Compared to existing materials, it reduces cost and weight, increases thermal stability, and maximizes internal payload.

Abstract: A metal airplane component frame is fastened to a skin of composite material by a countersunk, beveled-head rivet extended through aligned openings in the metal layer and in a high load bearing strength material inlay embedded in the composite material skin. The beveled head of the rivet lies within the skin.

Abstract: A reinforced elastomer panel (100) has a first rigid member (104) with an edge (116) attached to a first end of an elastomer panel (114). The first rigid member (104) has a removable plate (106) that provides access to a cavity (108). Attached to the first rigid member (104) is a reinforcing member (102) that extends through the elastomer panel (114). A second ridge member (118) has an edge attached to a second end of the elastomer panel (114).

Abstract: The present invention is a nanocomposite layer applied to the leading edge of an airfoil. The nanocomposite is an elastomeric matrix reinforced with nanosized reinforcing particles ranging in size from about 0.5 to 1000 nanometers and preferably ranging in size from about 5 to 100 nanometers. These nanosized reinforcing particles improve the mechanical properties of a pure elastomer. Improving such properties assists the nanocomposite layer in absorbing and dissipating the impact energy caused by the repeated impact of particulate matter, such as sand and rain. Adding the nanosized reinforcing particles to an elastomer allows the nanocomposite to absorb the impact energy over a larger volume in comparison to a pure elastomeric material. The nanocomposite layer is, therefore, capable of withstanding the repeated impact caused by the particulate matter for a greater length of time compared to a pure elastomer.

Abstract: A thermal blanket (28) for use in connection with a spacecraft (10) or spacecraft component for providing a thermal control coating. The thermal blanket (28) includes a plastic substrate (22) on which is deposited a silicon film (24) by a vacuum deposition process. The silicon film (24) provides a relatively high infrared light transmission and moderate absorption of high energy bandwidths in the solar spectrum that allows for the reflectance of high energy visible light and the emittance of infrared radiation. Additionally, the silicon film (24) is a hard, robust coating that provides resistance to humidity and other corrosive environments.

Abstract: A composite member having increased resistance to delamination includes a reinforcing wrap of fiber-matrix composite material which is encircled about the member so that the edge surfaces of the member are covered by the wrap. The wrap includes fibers which are oriented from about −60° to about +60° relative to the direction along which the reinforcing wrap is wrapped about the member. The fibers of the reinforcing wrap impart increased third-axis or Z-direction strength at the edges of the member where interlaminar tensile stresses are greatest.

Abstract: A system is provided for filling a gap between adjoining coplanar first and second sheet members mounted on an outer surface of a common substrate and having juxtaposed first and second edge surfaces. A permanent gap filler member of predetermined shape and size intended for eventual permanent reception in the gap is employed for partially filling the gap. A temporary gap filler tool which includes a substitute gap filler element having substantially the same predetermined size and shape of the permanent gap filler member, has an abutting surface for contiguous engagement with the first edge surface of the first sheet member when advanced to an attached position before the permanent gap filler member is mounted in the gap, a minor gap surface facing opposite said abutting surface, and means for releasably attaching the temporary gap filler tool to the first sheet member. A minor gap is thereby defined between the temporary gap filler element and the second edge surface of the second sheet member.

Abstract: A missile has a body with a substantially circular nose opening therein, and a ceramic radome sized to cover the nose opening. A compliant metallic circular transition element is disposed structurally between the radome and the body. The transition element includes an elongated compliant arm region and an upper crossbar region positioned adjacent to the radome such that the lower margin surface of the radome is adjacent to an upper side of the crossbar region. A brazed butt joint is formed between the lower margin surface of the radome and the upper side of the crossbar region of the transition element. A second brazed butt joint is formed between the vehicle body and the lower side of a lower crossbar region joined to the arm region. The brazed joints are formed with a single active braze alloy which permits the entire joining operation to be accomplished in a single furnace cycle.

Abstract: A box-structure airfoil is constructed of a composite material upper skin, a composite material lower skin 12 and a spar. Ribs and an elongate projection are formed integrally on the inner surface of each of the two skins. The upper and lower skins and the spar are simultaneously bonded by an adhesive to form a single structure. Since the ribs and the elongate projection are one-piece formed with each skin, it is possible to reduce the number of principal constituent components. Furthermore, since assembly operation is made by using the adhesive, there is no need for fasteners or the like for assembling. Moreover, the spar is bonded not only to Moreover, because not only a web but also flanges on both sides of the spar are bonded to the respective skins, it is possible to obtain a large strength. According to the invention, it is possible to reduce the number of principal constituent components and that of assembling components of an airfoil, thereby achieving its cost reduction.

Abstract: The present invention is directed to polishing the surface of an airplane part to an improved specularity by adding the additional step of sanding prior to the polishing step.

Abstract: A coated aerospace aluminum alloy part, such as a lipskin of a nacelle or a leading edge of a wing or tail, is disclosed which part is protected from corrosion due to severe environmental stresses, including high and low temperatures, and rain drop erosion, by a multilayer coating comprising a phosphate bonded metal containing basecoat and a resin based topcoat. A method of coating the part is disclosed.

Abstract: A computer system automatically generates CNC code for a stitching machine. The computer determines the locations of a present stitching point and a next stitching point. If a constraint is not found between the present stitching point and the next stitching point, the computer generates code for making a stitch at the next stitching point. If a constraint is found, the computer generates code for changing a condition (e.g., direction) of the stitching machine's stitching head.

Abstract: The invention provides a hybrid laminate and skin panels of hybrid laminate structure that are suitable for a supersonic civilian aircraft. The hybrid laminates include layups of layers of titanium alloy foil and composite plies, that are optimally oriented to counteract forces encountered in use, that are bonded to a central core structure, such as titanium alloy honeycomb. The reinforcing fibers of the composite plies are selected from carbon and boron, and the fibers are continuous and parallel oriented within each ply. However, some plies may be oriented at angles to other plies. Nevertheless, in a preferred embodiment of the invention, a substantial majority of, or all of, the fibers of the hybrid laminates are oriented in a common direction. The outer surfaces of the laminates include a layer of titanium foil to protect the underlying composite-containing structure from the environment, and attack by solvents, and the like.

Abstract: A tunable, mechanical and acoustical damping system is integrated into a lightweight composite aerospace vehicle structure. In one embodiment, the damping system (22) is incorporated into the wall of a spacecraft payload fairing (12) and includes an outer skin (24), a grid structure (26), an inner skin (28) and a tuning conduit (30). The grid structure (26) forms a number of cells (36) some of which may be interconnected. The cell volumes and the tuning conduit (30) are effective as a Helmholz resonator to reduce potentially damaging noise in the payload compartment (20) of the fairing (12). By appropriate selection of the conduit's dimensions in conjunction with the cell dimensions, the resulting resonator can be tuned to damp a selected frequency or frequencies. For example, a resonant frequency of the fairing (12) with its payload (18). A viscoelastic material (34) is provided between the grid structure (26) and inner skin (28) for mechanical damping.

Abstract: A reinforced elastomer panel (100) has a first rigid member (104) with an edge (116) attached to a first end of an elastomer panel (114). The first rigid member (104) has a removable plate (106) that provides access to a cavity (108). Attached to the first rigid member (104) is a reinforcing member (102) that extends through the elastomer panel (114). A second rigid member (118) has an edge attached to a second end of the elastomer panel (114).

Abstract: A coated ceramic has alumina as an amorphous residue from an alumina sol containing alumina powder or and, optionally, a high emittance powder. The alumina powder alters the rheology and wetting characteristics of the sol to aid in forming the residue and increases the alumina concentration (% solids) in the residue. The high emittance powder radiates heat out of the residue and away from the ceramic when the coated ceramic is heated, such as during orbital reentry.

Abstract: Currently, fiber/metal laminate sheets are limited in width due to restrictions in the width of thin metal sheets available. This invention solves that problem by providing an integral splice concept and a method for manufacturing a laminate employing the inventive concept. Specifically, a fiber/metal laminate sheet is provided having at least two metal plies and at least one fiber layer. Each of the metal plies comprises at least two metal sheets which are arranged side-by-side and have sheet metal breaks therebetween. All of the sheet metal breaks lie within the integral splice. An important feature of the invention is that the sheet metal breaks in each metal ply are staggered widthwise across the laminate with respect to the sheet metal breaks in the remaining metal plies, in accordance with a predetermined metal break staggering pattern, thereby maximizing the bond engagement of the discontinuous metal plies with the adjacent fiber layers.

Abstract: A hinge line system (52) for an aircraft has a structural block (22) attached to a first edge of a hinge line. The structural block (22) has a flange (28) attached to a first end of an elastic sheet (32). A second structural block (24) is attached to a second edge of the hinge line. The second structural block (24) has a second flange (30) attached to a second end of the elastic sheet (32).

Abstract: An aircraft with supporting wings including wing shells with good shear strength made of fibrous composite materials, particularly fiber-reinforced plastics, having members taking up tensile and compressive forces on the inside of the wing shells. The members have unidirectional fibers extending longitudinally of the wing, and fixing elements which can be joined detachably to fuselage attachments are provided on the members at the root end of the wing. Stringers are constructed on the inside of the wing shells and spaced longitudinally of the shells, their fiber component being formed by a fiber layer joined to the fiber layer of the wing shell. Fiber bundles are arranged between spaced stringers as unidirectional stiffening elements. The fiber bundles are embedded in the synthetic resin matrix of the plane load-bearing structure and extend longitudinally of the supporting wing.

Abstract: Method and apparatus for a reinforced skin structure such as an aircraft fuselage which has reduced interior cabin noise. The reinforced skin structure according to the present invention, includes a frame assembly which has a plurality of spaced apart frame members and an outer skin mounted on the frame assembly. A plurality of inner panels, which generally defines an interior compartment, are incorporated within the skin structure. Each of the inner panels is typically supported in a spaced apart fashion from the frame members. A septum is disposed between the outer skin and at least some of the plurality of inner panels. The septum acts to provide an air barrier and increases the air resonant frequency of the cavity between the outer skin and the inner panels. A layer of insulation is placed on the inner side of the septum removing the need to place insulation against the outer skin.

Abstract: A constrained damping layer assembly includes a first constraining layer adapted for being secured to a first portion of an aircraft frame, and a second constraining layer adapted for being secured to a second portion of the aircraft frame. A damping material is sandwiched between the first constraining layer and the second constraining layer. The damping material includes a first side, which is adapted for slidably contacting the first constraining layer, and further includes a second side, which is adapted for slidably contacting the second constraining layer. The first constraining layer is adapted for sliding in a first direction upon introduction of a sheer load onto the interior panel, and the second constraining layer is adapted for sliding in a second direction upon introduction of a sheer load onto the interior panel.

Abstract: A sheath for covering and protecting a component leading edge of an airfoiled component is disclosed. The sheath includes a sheath leading edge, and a first protective side and a second protective side, wherein the first and second protective sides are merged at the sheath leading edge. A cavity is formed between the first and second protective sides, wherein the cavity is adapted to have the airfoiled component positioned therein and engage an inside surface of each of the first and second protective sides. The sheath is formed from a material including cobalt, and preferably a nickel-cobalt composition. Preferred embodiments include cobalt in the nickel-cobalt composition present between 8-32 wt. % and 30-54 wt. %, based on the weight of the sheath. An airfoiled component construction is also disclosed.

Abstract: An aircraft surface for use on composite structure. The detail surface includes a conductive foil, a conductive double pulled expanded screen, and an adhesive for joining the conductive foil and screen to a composite material. The conductive screen is layered between the conductive foil and composite structure, providing electrical contact between the foil and the composite structure.

Abstract: The invention is a sealing assembly for reducing the gap between an aircraft's movable flight control member and adjacent structure thereto, the flight control member and structure having external surfaces and side surfaces facing each other. In detail, the assembly includes first and second flexible bulb seals mounted on the periphery of the side surface of both the flight control member and the adjacent structure forming an extension of the external surface of both. Each of the bulb seals includes a compressible foam bulk absorber core; the core having a bottom surface, a top surface generally parallel to the bottom surface, an outer side surface, and an inward facing surface at an acute angle to the top surface. A first sheet of dielectric material is bonded to the bottom surface of the core. A second sheet of flexible resistive material is bonded to the top, and sides of the core. A third sheet of flexible magnetic radar absorbing material is bonded to the second sheet.

Abstract: Hygroscopic ceramic materials which are difficult to waterproof with a silane, substituted silane or silazane waterproofing agent, such as an alumina containing, fibrous, flexible and porous, fibrous ceramic insulation used on a reentry space vehicle, are rendered easy to waterproof if the interior, porous surface of the ceramic is first coated with a thin coating of silica. The silica coating is achieved by coating the interior surface of the ceramic with a silica precursor, converting the precursor to silica either in-situ or by oxidative pyrolysis and then applying the waterproofing agent to the silica coated ceramic. The silica precursor comprises almost any suitable silicon containing material such as a silane, silicone, siloxane, silazane and the like applied by solution, vapor deposition and the like. If the waterproofing is removed by, e.g., burning, the silica remains and the ceramic is easily rewaterproofed.

Abstract: A continuous moldline technology system (100) has a pair of end blocks (102, 104). One (102) of the pair of end blocks (102, 104) has an edge (112) attached to a first edge of an elastomer panel (110). The other block (104) has an edge attached to a second edge (114) of the elastomer panel (110). A composite slat (116) having a rectangular cross section is partially disposed inside the elastomer panel (110).

Abstract: A structural component, especially a substantially hollow aircraft structural component such as a wing or tail section having an upper chord or shell and a lower chord or shell interconnected to form a substantially hollow bending beam in use, is so constructed that the chord or shell which is primarily taking up tension stress is made of fiber reinforced composite material, while the other chord or shell which primarily takes up compression stress is made of metal. In an aircraft wing the upper wing shell will be made of metal and the lower shell will be made of fiber reinforced composite material. However, the tail plane and elevator assembly will have an upper shell made of fiber reinforced composite material while the lower shell will be made of metal.

Abstract: A refractory composite structure comprising a ductile refractory metallic layer with a roughened surface which is tightly bonded to a refractory composite structural shell. The roughened surface is dendritic in form and is produced by chemical vapor deposition techniques.

Abstract: A vehicle having a ceramic radome includes a vehicle body having an opening therein, a ceramic radome sized to cover the opening of the vehicle body, and an attachment structure joining the radome to the vehicle body to cover the opening. The attachment structure has a substantially cylindrically symmetric compliant metallic transition element disposed structurally between the radome and the body. The radome is joined to a first end of the transition element by a first brazed butt-joint. A primary seal is disposed between the transition element and the vehicle body. A spanner nut overlies the transition element and has a first engagement to the transition element. A threaded engagement between the spanner nut and the vehicle body allows the radome/transition element assembly to be attached and tightened to the vehicle body and to be later removed.

Abstract: In a wing with wing shells with good shear strength made of fibrous composite materials, particularly fiber-reinforced plastics, for aircraft, members taking up tensile and compressive forces are attached to the inside of the wing shells. These members have unidirectional fibers extending longitudinally of the wing. Stringers spaced longitudinally of the wing are provided on the inside of the wing shells, their fiber component being formed by a fiber ply joined to the fiber ply of the wing shell. Unidirectional fibre bundles are arranged between spaced stringers, the bundles being embedded with good shear strength in the synthetic resin matrix of the wing shells, extending longitudinally of the wing and having a substantially rectangular cross-section.

Abstract: A gap filler device applicable to a gap formed between side walls of adjacent panels comprises an elongated horizontal portion extending across the gap and a vertical portion extending into the gap. The vertical portion includes a pair of opposed, spaced apart, planar bearing surface members biased into frictional engagement with the side walls of the gap to resist removal of the gap filler device. The vertical portion has descending legs extending from opposed sides of the horizontal portion to a bight. Ascending legs extend from the bight to terminal ends. Slots are formed in the vertical portion and extend upwardly from the bight at spaced locations along the length of the legs. In one embodiment, a plurality of outwardly projecting spike members are formed on the bearing surface members for engagement with the side walls of the gap.

Abstract: An insect abatement system prevents adhesion of insect debris to surfaces which must be kept substantially free of insect debris. An article is coated with an insect abatement coating comprising polyorganosiloxane with a Shore A hardness of less than 50 and a tensile strength of less than 4 MPa. A method for preventing the adhesion of insect debris to surfaces includes the step of applying an insect abatement coating to a surface which must be kept substantially free of insect debris.

Abstract: A radome for protecting electromagnetic equipment includes a polyimide foam that preferably is a closed cell foam. The polyimide foam imparts improved impact and moisture resistance to the radome without adversely affecting electromagnetic transmission thereof.

Abstract: A pliant, controllable contour control surface comprising a first flexible facesheet formed to a first initial contour of the control surface, and a second flexible facesheet formed to a second initial contour of the control surface. The first and second facesheets each have a set of prestrained shape memory alloy tendons embedded therein, extending from a leading edge to a trailing edge of the control surface. Each set of the shape memory alloy tendons is separately connected to a controlled source of electrical current such that tendons of the first and second flexible facesheets can be selectively heated in an antagonistic, slack-free relationship, to bring about a desired modification of the configuration of the control surface. A computer based control system is utilized for maintaining a constant temperature of the antagonists to establish conditions conducive to the stress induced transformation from austenite to martensite, accomplished by causing constant current to flow through the antagonists.

Abstract: A flush mounted panel fastener is provided for use in aircraft or the like of the so-called stealth type designed to avoid detection by radar. The panel fastener comprises a fastener element such as a threaded bolt having a head and shank for fastening a panel onto a substrate or frame. The head is sized for recessed seating within a counterbore formed in the panel, and a fastener cap is mounted and magnetically retained on the head of the fastener element. The fastener cap has a surface coating or finish thereon disposed substantially coplanar with an outer surface of the panel, wherein the surface coating is formed from a suitable nonreflective radar material and cooperates with the panel to present a substantially uninterrupted surface of low radar signature.

Abstract: A process for making a layered product having a honeycomb core having cells filled with a foam material, the process comprising the steps of: (a) placing a first layer of an uncured, heat-expandable, foamable material on the top side of a central honeycomb core having empty cells, and placing a second layer of an uncured, heat-expandable, foamable material on the bottom side of the central honeycomb core; (b) placing a first layer of an uncured preform material above the first layer of uncured, heat-expandable, foamable material, and placing a second layer of an uncured preform material above the second layer of uncured, heat-expandable, foamable material; (c) placing the charge made by steps (a) and (b) inside a mold and closing the mold; (d) heating the mold to the cure temperature of the heat-expandable, foamable material, and holding the mold at this temperature for sufficient time to expand and cure the heat-expandable, foamable material; (e) reducing the temperature of the mold to the injection temperat

Abstract: A process for making a layered product having a honeycomb core having empty cells, the process comprising the steps of: (a) placing a first layer of an uncured adhesive film on the top side of a central honeycomb core having empty cells, and placing a second layer of an uncured adhesive film on the bottom side of the central honeycomb core; (b) placing a first layer of an uncured prepreg material above the first layer of uncured adhesive film, and placing a second layer of an uncured prepreg material below the second layer of uncured adhesive film; (c) placing a first layer of a dry fiber preform above the first layer of uncured prepreg material, and placing a second layer of a dry fiber preform below the second layer of uncured prepreg material; (d) placing the charge made by steps (a)-(c) inside a mold and closing the mold; (e) heating the mold to the cure temperature of the adhesive film and the prepreg material, and holding the mold at this temperature for sufficient time to cure the adhesive film and the