Strengthened panel |
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IPC classes for russian patent Strengthened panel (RU 2469907):
        
 Rotary element to increase lift, particularly, deflecting wing tip with high aerodynamic characteristic / 2414386
Invention relates to aircraft engineering. Wing deflecting tip features profile of aircraft wing. Wing 10 comprises torsion 10a has first side with first skin 11 and second, opposite, side with second skin, and torsion box edge facing deflecting tip 20 with front skin that is a continuation of first skin 11. Wing deflecting tip 20 comprises main case 21 that allows transition from first side to second side, and first transition part 20a facing the first side, and second transition part 20b that faces second side. Said wing deflecting tip 20 can be set by retention mechanism 23 and drive that coupled said tip with torsion box 10a between first, IN, position and second, OUT, position. First transition parts 20a of the tip can slide on over wing front surface. Wing deflecting tip can be mounted between first and second positions at preset angle to wing torsion box 10a, and wing profile may be widened around imaginary rotational axis 30.
 Aircraft flexible control surface / 2408498
Flexible elastoplastic control surface (1; 11, 14) is, in fact, flat in span direction (9) ram airflow direction (6) and comprises actuators (3) acting on control surface (1; 11, 14) at various points of force application spaced apart in crosswise direction with respect to ram airflow direction (6). Actuators (3) are configured so that, when activated, they deflect aforesaid pints (2) to allow elastic deformation without jogs on control surface (1; 11, 14) in directions of span (9) and ram airflow (6).
 Hollow soft wing with air intake at tip and shaped slot on upper surface / 2389644
Invention relates to aircraft engineering. Hollow soft wing with air intake at tip comprises lower surface, upper surface and ribs. Upper surface is furnished with pockets that forms tapered duct to allow airflow to flow through upper slot on wing upper surface.
 Aircraft wing / 2380277
Invention relates to aircraft engineering, particularly to light airplanes. Proposed wing comprises spars, stringers and skin, and has back step-cutout arranged on upper surface that features slightly curved bottom and extends along wing span. Wing upper surface support panels designed to regulate the depth of aforesaid step-cutout and pivoted to vary their position with the help of springs. The latter are rigidly coupled with the recess bottom and regulation panel lower surface.
 Flat drive with foliated structure, flat drive device and lifting plane of aircraft / 2337430
Invention is related to piezoelectric instruments for control of aircraft lifting planes. Flat drive with foliated structure that is symmetrical in relation to middle plane includes flat piezoelectric layer, which has active direction and is connected with one flat passive layer of cloth with rigid haircloth and weft that are oriented in accordance with two directions that form network of cells. Both directions of every cloth layer are same. Active direction of every piezoelectric layer is oriented along single diagonal of cloth layers cells. Flat drive device consists of two drives installed top to tail. Lifting plane of aircraft, for instance, helicopter blade, includes top and bottom surfaces, and also drive device close to back edge.
 Swept-forward wing with swivel part of outer wing panels / 2296082
Each outer panel of proposed wing includes wing extension with sweep-back leading edge, root section with swept-forward leading edge and trailing edge; swivel part of wing outer panel is articulated with this section relative to vertical axis of flying vehicle; it may be turned backward in way of flow so that sweep angle of leading edge may change from initial swept-forward to swept-back magnitudes.
 Locking mechanism / 2467920
Invention relates to aerospace engineering, particularly, to locking of replaceable elements of pressure compartment porthole. Locking mechanism comprises spring-loaded holder and retainer fitted on axle in casing. Retainer is composed of two-arm lever while holder represents lug with flanges. Said flanges are composed of cylindrical rolls arranged on edges of aforesaid lug on retainer side to get in contact by outer surfaces with retainer arm that are furnished with inner shoulder to hold holder in open position by interaction with ledges arranged on the casing and to hold that holder in closed position in interaction with retainer arm. Pressure surface of the lug is provided with elastic lining. Holder and retainer are furnished with individual springs. Holder spring force is lower than that of retainer.
 Cross-butt joint of two fuselage sections / 2467919
Invention relates to aircraft engineering. Cross-butt joint (1, 24) of two fuselage sections (2, 3, 25, 26) have end areas (6, 7, 29, 30) are provided with inclined wedge skin surfaces (8, 9). Both end zones (6, 7, 29, 30) are jointed together by cross-butt strap (10, 31) with two inclined wedge surfaces (12, 13) on bottom side (11). Between wedge surface (12, 13) of said strap and skin wedge surface (8, 9) may be fitted wedge (14, 15, 33, 34) to compensate allowances between two fuselage sections (2, 3, 25, 26). Said strap (10, 31) is assembled by jointing two curvilinear segments (23, 32) to align first section (2, 25) with strap (10, 31). Then, wedges (14, 15, 33, 34) for allowance compensation are fitted in place to joint end zone (6, 29) of first fuselage section with strap (10, 31). Similarly, second fuselage section is jointed to cross-butt strap (10, 31).
 Skin element as aircraft fuselage part / 2466905
Skin element is formed to bent sheet element and made partially or completely from black-reinforced plastic. Skin element thickness varies over skin element width and/or length on skin element side facing the cabin. Skin element features lengths varying from 10 m to 60 m and is made by laminating or spraying fibers. Method of producing skin element comprises making mould to form skin element outer layer and applying layers of resin of layer of fiber, or mixing resin, reinforcing agent and fibers, and applying said mix in mould by spray gun.
 Method of compensating clearances between two fibrous composite parts / 2466059
Invention relates to aircraft engineering, particularly, to jointing aircraft two fibrous composite parts. Proposed method comprises the following stages: a) making first fibrous composite part 1. Note here that seat surface 2 may be offset from tolerances, b) making tool holder 7 to fit in surface shape 2, c) making second fibrous composite part 3 with the help of tool holder 7. Note here that shape of contact surface 9 of second part 3 matches mainly with seat surface 2. d) It includes also jointing first fibrous composite part 1 and second fibrous composite part 3 in the zone of mount surface 2 and contact surface 9.
 Method for light aeroplane covering / 2463218
Invention relates to method for covering light airplanes with polyester sheathing fabric, as well as to dispersive hot-gluing glue and its application for covering. For covering, polyester sheathing fabric with longitudinal shrinkage of 7% and transversal shrinkage of 5% at 160-180°C is used. In the process of covering light airplanes and/or their parts consisting of frame system, sheathing fabric in the area of its overhangs and frame parts is covered by dispersive hot-gluing glue and wrapped around longeron part of frame so that when connection between sheathing fabric and frame produced by glue is destructed the fabric could be held on frame carcass. Dispersive hot-gluing glue contains 80-88% of adhesive, 12-15% of hardener and 0.15-0.3% thickener, and the glue is polymerised at temperatures >40°C. The hardener contains 54-60% of solvent, 0.35-0.5% of naphthalensulfonic acid sodium salt-based stabiliser, 0.25-0.35% of propoxylated spirit-based emulsifier, 1.7-1.9% of polyetheramines-based hardener, 37-41% of polyisocyanate.
 Selfsupporting internal cabin structural block with integrated cabin equipment components / 2463207
Invention relates to aeronautical engineering, more specifically to aircraft cabin structural block and attachment method for cabin equipment components. Cabin structural block designed for attachment of cabin equipment components is made capable to be selfsupporting and to be mounted to aircraft floor (6) structure. Method of cabin equipment components attachment for aircraft which includes attachment of cabin equipment component (2) to cabin structural block (1), attachment of cabin structural block (1) to aircraft floor structure (6), where cabin structural block (1) is made capable to be selfsupporting.
 Cladding for aircraft / 2463206
Invention relates to aircraft cladding. Cladding for aircraft fuselage bay contains multiple mutually adjacent cladding panels. Between fuselage bay and cladding, interspace is provided. Cladding panel has such connection (going through interspace) with fuselage bay that in case of sudden pressure drop between fuselage bay internal space and interspace gap opening is provided to balance pressure.
 Frame element, aircraft component assembly system and method of mounting component in aircraft / 2457980
Invention relates to aircraft engineering, particularly, to frame element to be used in assembly of aircraft components, in particular, inner component or insulation pack. Frame element 10, 10' to be used in assembly system 46 may be jointed to aircraft structure 36 and comprises tie-down fitting to secure inner element 34 or insulation pack 52 to frame element 10, 10'. Said frame element 10, 10' comprises two tie-beams 12; 14; 12'; 14' arranged in parallel and connection tie-beam 16; 18; 16' extending between the latter. Connection tie-beam 16, 18 has recess 26, 28, 30, 32 to receive frame 42, 44 in jointing frame element 10 to structure 46. Assembly system 46 comprises several frame element 10, 10'. Said frame element with component 34, 53 jointed thereto is jointed to aircraft structure 36.
 Composite structure / 2455194
Invention relates to composite structure and method of its production. Composite structure comprises panel made up of two or more layers of composite material, and two or more stiffeners. Panel has surface with step whereon panel thickness varies. First of said stiffeners is attached to surface panel on step thicker side. Second stiffener is engaged with step part. Proposed method comprises making the panel with step and jointing stiffener to panel surface.
 Method, device and plant for manufacturing skin structures / 2448876
Invention relates to aircraft engineering, particularly, to methods and appliance used for production of aircraft skin structures. Equipment used in production and conveying of aircraft skin structures comprises gripping device to grip skin structures including initial skin with edge. Said gripping device comprise bearing device for conveying and handling and locking grip coupled with said bearing device. Locking grip allows gripping initial skin by gripping aforesaid edge.
 Metal structural member of skin / 2249538
Proposed structural member has built-in profiled rigidity members. Profiled rigidity members (2) have bulge (4) of definite thickness (f) which reduces towards joint area (3) or connection to definite thickness (a) at this joint area or connection; ratio of thickness (f) of base of profiled rigidity member (2) to thickness (a) at joint area or connection is more than or is equal to 2.
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FIELD: construction. SUBSTANCE: panel comprises a composite lining, multiple stringers attached to the lining, and actuation mechanisms of force application. Each mechanism is arranged between neighbouring pairs of stringers and is arranged as capable of applying a local force to the lining, causing bending of lining between stringers to form a group of several folds in the lining. The method to bend a panel includes application of a compressing load to the panel in the plane of lining and a local force to the lining between neighbouring pairs of stringers with the help of actuation mechanisms, when the load exceeds a preset threshold. EFFECT: improved operational indices and reduced weight of a structure. 15 cl, 10 dwg
The technical FIELD The invention relates to reinforced panels containing composite siding. The invention also relates to a method for bending such reinforced panels. PRIOR art Despite the high performance characteristics in certain areas, such as weight, durability and costs over the entire service life, composite materials are not widely used for load-bearing structures. The reason for this mainly is the lack of understanding of the mechanisms of their destruction and behavior at destruction. This widespread lack of knowledge and know-how that often leads to the creation of structures that are too large, which contradicts the concept of light weight, distinguishing all new design solutions. Bending is one of the most contentious issues concerning design of reinforced panels. It is well known that the load in the plane that can withstand panel reinforced composite material, without any damage, higher than the load when bending. Unfortunately, the complexity and high cost of tests simulating this behavior is associated with the destructive impact due to structural factors, making a recovery mechanism for the destruction of the debris is very complicated. For metal and yacynych structures removal and redistribution of stresses essentially provide with the help of local plastic deformations, and destruction occurs as a result of bending of plating or local/common longitudinal bending element stiffness. Local plastic deformation in composite materials are very rare, so this method of stress relief is basically unusable. The INVENTION In accordance with the first aspect of the present invention proposed reinforced panel that contains the following elements: - composite siding; many stringers attached to the sheathing, and one or more actuators application efforts, each of which is located between adjacent pairs of stringers and made with the possibility of application of local efforts to the skin, causing bending of the shell. In accordance with the second aspect of the present invention, a method for bending toughened pane, which contains the composite siding and a few stringers attached to the casing, comprising the following operations: application of load to the panel in the plane of the sheathing and application local efforts to trim between adjacent pairs of stringers using one or more actuators application of force when the load exceeds a predefined threshold. The present invention is based on the fact that the release of tension, like the way which occurs as a result of local plastic deformation of the metal structures can be obtained in composite structures by creating a pre-bend plating. The load applied in the plane of the sheathing may be compressive load, shear load, or a combination of these two loads. In accordance with one embodiments of the present invention, each of the actuators force applies a local effort by modifying their geometry between the two stable States. In this case, each of the actuators application of force can change the geometry automatically, without requiring the control system. In accordance with other variants of the invention proposes a control system for control efforts to the sheathing, and actuate the actuators effort, when controlled force exceeds a predefined threshold. In this case, the actuators of the force can be, for example, piezoelectric devices. The use of a control system for active management and control can increase the range of operating loads and to provide information about the actual configuration, behavior and holistically the tee design. A set of specially designed strain gauge sensors can provide the necessary input to the control system. However, in the preferred embodiment, the actuators provide the necessary measuring input, i.e. each of the actuators is configured to determine the effort applied to the sheathing, and generating a signal read, which is controlled by the control system. This reduces the number of required control lines, as each of the actuators can transmit signals read management system, and to receive control signals from the control system, through a double line of control. The stringers and the hull can be bonded with an adhesive, by co-curing or any other suitable connection method. Actuators application of force can be at least partially embedded in the skin. This can eliminate the need to use adhesive to attach actuators application of force to the skin. In accordance with the options described here implement the present invention, the casing is produced from a composite material that contains many unidirectional carbon fibers impregnated with epoxy resin. However, obsh the WHC can be obtained from any composite material, including, for example, a metal laminated material, reinforced with glass fibre (GLARE). A BRIEF DESCRIPTION of GRAPHIC MATERIALS The present invention will be described below only in the form of an example and with reference to the accompanying drawings. Figure 1 shows the top view of a reinforced panel in accordance with one embodiments of the present invention. Figure 2 presents the cross-section along the line a-a of figure 1. Figure 3 presents the cross-section of the panel in an enlarged scale. Figure 4 presents one of the actuators in an enlarged scale. Figure 5 shows the electronic control system. Figure 6 presents a partially built-in enforcement mechanism. Figure 7 presents a fully integrated operating mechanism. On Fig presents a cross section of a fiber carrying control line. Figure 9 presents a top view of the reinforced panels with multistable (multiple steady States) Executive mechanisms. Figure 10 shows a graph of a typical variable boundary depending on load. INFORMATION CONFIRMING the POSSIBILITY of carrying out the INVENTION Figure 1 and 2 shows a part of the reinforced panel 1. Such a panel may be formed, for example, the wing skin and the and fuselage of the aircraft. The panel contains a composite trim panel 2; the set of composite stringers 3, 4, attached to the hull by co-curing; and two rows of the piezoelectric actuators 5 effort placed between the stringers. If we are talking about an aircraft wing, the stringers are in the direction of the wing span from the wing root to the end. Figure 1 presents only a small part of the panel, which extends further in both the horizontal and vertical directions. As can be seen in figure 2, each stringer includes a rib 3A, 4A, protruding from the casing, and a pair of flanges 3b, 4b, attached to the casing 2. Each of the actuators 5 application of force attached to the casing 1 layer 5A of the adhesive, are presented in figure 3 and 4. Two electrodes 10A and 10b are attached to the upper and lower surfaces of the actuator. Each of the electrodes is connected to a corresponding line 11a, 11b management, and line management of the joint in the cable 12 leading to the control system 13, are presented in figure 5. System 13 control actuates actuators, applying an electric voltage between the electrodes 10A, 10b. This leads to the expansion or compression of the actuators at right angles to the electric field due to the piezoelectric effect. Expanded the e or compression actuators is determined by the sign of the electric voltage. Actuators 5 application efforts also act as load sensors. The deformation of the panel leads to the expansion or compression actuators application of force, which, in turn, creates an electric voltage between the electrodes 10A, 10b. This voltage provides a signal readout controlled by the control system 13. The control system 13 generates the control signal when the monitored voltage exceeds a predetermined threshold stored in the memory 14. This control signal increases or decreases the voltage between the electrodes, which in turn causes compression or expansion of the actuators. It should be noted that, as the signal read from the actuators, and the control signal actuating mechanisms may be transferred through a single control line. The deformation of the actuators apply local pressure to the skin, resulting plating is bent between the stringers forming folds 6, shown in figures 1 and 2. Currently, for most designs of aerospace vehicles is not allowed to bend below the design limit load. Figure 10 shows a graph of the typical changes of the interface (for example, shift the load or displacement out of the plane at the boundary section is La between the hull and stringer) depending on load. In the area between P0 and P1 panel is subjected to excessive pressure in the application of the pressure difference between the opposite surfaces of the panel. This pressure difference exists, for example, by the presence on one side of the cladding of fuel under pressure. In the zone between the points P1 and P2 to the panel applied compressive load in the plane. At the point P2 of the panel is bent. Thus, the point P2 represents the maximum permissible load in the absence of the actuators 5, and accordingly determines the size of the structures. When the panel, which was previously subjected to excessive pressure, reaches a critical load, there would be significant change in the shape in the perpendicular direction. In fact, in a stable phase of the deformed configuration (represented by a single fold, formed by the application of pressure) into a group of several folds, passing in the direction of the stiffeners after reaching the critical load. This is due to the phenomenon of buckling, so the abrupt change of the mechanical variables characterizes the transition from steady state to unstable. Figure 1 shows three of many of these folds 6. If the external load food is continuing to act, the internal energy of elastic deformation before and after transition should remain unchanged. Assume that if the same amount of energy of elastic deformation affects only a single fold or scattered by several folds in the last case, the maximum movement outside the plane of characterizing each fold must be decreased, so that each component of stress at the interface will be proportionately decreased. This means that the bending leads to tension at the interface between the panel and the stringer. By installing a memory 14 threshold below the critical load P2 (for example, 60% or 80% of the critical load P2) actuators 5 create an additional field, mechanical stress, which forces the casing to bend to reach the critical load P2. This pre-bending leads to the subsequent release of tension at the interface of the covering/supporting element. Then the panel is working in the post-stable mode, and a reduced voltage at the interface reaches a critical value at higher load levels. As a result the working load capacity is improved, and you can get a significant improvement in overall performance and weight savings. Figure 6 presents an alternative design in which the actuator 5 is partially embedded in the recess in the surface of the casing 2. Figure 7 presents an alternative design, in which the actuator 5 is completely embedded in the casing 2. On Fig presents the cross-section of the fiber 21, which carries the line 11b of the control. The casing 2 and each of the stringers 3, 4 obtained from a sequence of composite layers, and each layer contains many unidirectional hollow carbon fibers impregnated with epoxy resin. In a partially or fully embedded designs presented on Fig.6 and 7, one or both lines 11a, 11b of the control can pass through the hollow core of the respective carbon fiber. This can be seen in Fig, which presents a cross section of a hollow carbon fiber 21, which contains a hollow core line 11b control. The space between the conductive metal line 11b management and conductive carbon fiber 21 is filled with the resin 20, which acts as an insulator. In accordance with an alternative embodiment of the present invention, presented in Fig.9, the piezoelectric actuators 5 replace multistable actuators 30, who exert local pressure to the skin by changing geometry between two or more stable States. An example of such design is asymmetrical layered design. Various ASI is unbalanced layered structure described in the following documents: "The application of residual stress tailoring of snap-through composites for variable sweep wings", 47thAIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, 1-4 May 2006, Newport, Rhode Island; and - "Bi-stable composites with piezoelectric actuators for shape change", C.R.Bowen, A.I.T.Salo, R.Butler, E.Chang and H.A.Kim, Key Engineering Materials Vols. 334-335 (2007) pp.1109-1112. The simplest example is a square plate with stacking sequence [0°/90°]. After cooling plate, which utverjdayut in a horizontal position at a high temperature, acquires a cylindrical shape, which can easily be clipped into the second cylindrical shape by application of force. In this case, does not require any electrical control system, no memory. Instead, the actuators 30 are properties inherent in the material, forcing them to snap into place between stable States when the load in the panel exceeds the desired threshold, thus leading to the preliminary bending of the cladding as a result of local efforts actuators 30. Although the present invention is described here with reference to one or more preferred embodiments, it should be clear that various changes or modifications not beyond the scope of the invention defined by the attached claims. 1. Reinforced panel containing: 2. The panel according to claim 1, additionally containing a control system configured to control effort applied to the casing, and actuating actuators effort, when controlled force exceeds a predefined threshold. 3. The panel according to claim 2, where each of the actuators application the efforts made to determine the effort applied to the skin, and generating a signal readout controlled by the control system. 4. The panel according to claim 3, where each of the actuators is arranged to transmit signals read control system and receiving a control signal from the control system through a common dual line control. 5. The panel according to claim 1, where each of the actuators application the efforts made with the possibility of application of local efforts by changing their geometry between the two stable States. 6. The panel according to claim 1 or 2, gastinger and sheathing are jointly otverzhdennye. 7. The panel according to claim 1 or 2, where each of the actuators attached to the sheathing with a layer of adhesive. 8. The panel according to claim 1 or 2, where each of the actuators force at least partially embedded in the hull. 9. The panel according to claim 1 or 2, where each of the actuators effort is a piezoelectric device. 10. The panel according to claim 1 or 2, where the covering is formed from a layered composite material. 11. The panel according to claim 1 or 2, where the folds are formed at a distance from each other in the direction of the stiffeners. 12. The method of bending the reinforced panels containing composite siding and many stringers attached to the casing, including: 13. The method according to item 12, further comprising control effort applied to the sheathing, and actuation actuators application of force when the monitored load exceeds a predefined threshold. 14. The method according to item 12 or 13, where through one or more Executive fur the isms to the skin apply a local force, which leads to the bending of the skin between stringers with the formation of a group of several folds in the skin. 15. The method according to 14, where the folds are formed at a distance from each other in the direction of the stiffeners.
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