FR2579169A1 - Aircraft comprising at least one lift structure with at least two superimposed wings - Google Patents

Abstract

The present invention relates to an aircraft comprising, on either side of its fuselage, at least one lift structure consisting of at least two wings. The outboard ends of the wings 5, 7; 6, 8 are linked together, each structure 3, 4 forming a closed polygon together with the sides of the fuselage 2, seen from the front, the wings being fixed together, to the fuselage and, possibly, to the said main plane element in such a way as to be articulated in rotation in a plane at least substantially transversal to the aerodynamic flow direction, and the aircraft is characterised in that at least one articulation axis 13, 14, 17, 18 is situated within the area defined by the main inertial planes, parallel to the wing, of the upper 5, 7 and lower 6, 8 wings. The invention applies to aircraft with multiple main planes.

Description

The present invention relates to an aircraft comprising at least one support structure with at least two superimposed wings.

In French patent No. 2,523,072, belonging to the same applicant, an aircraft of the type is described comprising, on either side of according to the fuselage, at least one support structure consisting of at least two wings, the distal ends of which are connected, either to each other, or to at least one wing element provided between them, each structure forming, in front view, with the sides of the fuselage, a closed polygon, and comprising an upper wing and a lower wing, the zones of attachment of the upper and lower wings being optionally offset in the direction of the aerodynamic flow, the values of the moment of inertia and of the lift each taking a maximum value in the central zone of the wing considered in the longitudinal direction, the attachment of the wings to each other, to the fuselage and, optionally, to said wing element being produced in an articulated manner in rotation in a plane at least substantially transverse to the direction of flow aerodynamic. It will be noted that, by moment of inertia, is meant the moment which can oppose the bending moment due to the force exerted on the wing tending to bend the latter.

An aircraft, as defined above, exhibits significant qualities of resistance, in particular to torsion; it is not very deformable, of relatively low weight and has the characteristics necessary for achieving an aerodynamic flow under the best lift conditions.

However, after intensive research, the inventor realized that such a type of aircraft could see its qualities improved significantly by slightly modifying its basic design.

To this end, the aircraft, according to the present invention, of the type described above, is particularly remarkable in that at least one articulation axis is located inside the area defined by the planes of inertia main, parallel to the wing, upper and lower wings.

In the case where each lifting structure of the aircraft consists of two upper and lower wings, the ends of which are distant from the fuselage and are directly connected to one another in an articulated manner, so as to constitute, substantially, an articulated triangle, the axis of articulation between said ends of the upper and lower wings is situated inside the zone defined by the main planes of inertia parallel to the wing, upper and lower wings.

In the case where each lifting structure of the aircraft consists of two upper and lower wings connected, at their ends remote from the fuselage, to a wing element, in an articulated manner, so as to constitute, substantially, an articulated trapezoid, at least one axis of the articulation to said wing element is located in the zone defined by the main planes of inertia, parallel to the wing, of the upper and lower wings.

In the two aforementioned cases, at least one axis of articulation of the upper and lower wings of the fuselage may be located in the zone between the main planes of inertia, parallel to the wing, of the upper and lower wings.

According to another characteristic of the invention, at least some of the above-mentioned joints are not completely free.

According to yet another characteristic of the invention, the abovementioned lift structures comprise at least one stiffening device, which may comprise, in the vicinity of the upper and lower surfaces of each wing, in the central zone of said surfaces, at least one beam of materials, the modulus of elasticity of which is greater than that of the material constituting the surface of the wing, and which extends in the longitudinal direction of the wing.

In particular, the above-mentioned bundle, with unidirectional or twisted fibers, is housed in a tube, in particular a metal tube, integral with the wing; a connecting tube forming, preferably, part-e integral of the bundle being possibly provided between the bundle and the internal wall of the tube serving as housing for the bundle.

In a particular embodiment, the stiffening device comprises four aforementioned beams arranged, two by two, on the upper and lower surfaces respectively of the wing symmetrically with respect to the longitudinal plane of inertia of the wing, parallel to the wing, and relative to the longitudinal plane of inertia of the wing, perpendicular to the wing.

According to another characteristic of the invention, each aforementioned beam comprises at least one damping device, in particular of the piston-cylinder type, the piston of which is integral with the aforementioned beam and of which, in a first case, the cylinder comes to bear against the end of the wing and, in a second case, the cylinder is integral with the frame supporting the end hinge pin.

According to yet another characteristic of the invention, at least substantially in the median transverse plane, perpendicular to the wings, of the upper wing assembly, lower wing, at least one shoring device is provided connecting the upper and lower wings.

In particular, the shoring device comprises at least one bar, in particular a metal bar, possibly provided with a damping device such as, for example, a piston-cylinder assembly.

In another embodiment, the shoring device comprises at least one flexible bar by construction or initially arched.

According to another characteristic of the invention, each aforementioned bar is connected; at each end, to a box provided in the corresponding wing and consisting of two symmetrical spaced half-boxes, formed of spaced ribs, in particular metallic, internal and external, the space between them is filled with a material, in particular synthetic , the internal ribs protruding from the surface of the wing so as to allow connection with the bar.

The invention will be better understood and other details, characteristics and advantages thereof will appear more clearly in the light of the explanatory description which will follow of currently preferred embodiments of the invention which will be described with reference to the schematic drawings. attached, in which - Figure 1 is a schematic front view of a first embodiment of the aircraft according to the invention - Figure 2 is a top view of the aircraft of Figure I - Figure 3 is a schematic front view of a second embodiment of the aircraft according to the invention - Figure 4 is a top view of the aircraft of Figure 3 - Figure 5 is a partial front view of a lift structure of the aircraft of Figures 1 and 2 - Figure 6 is a partial front view of a lift structure of the aircraft of Figures 3 and 4 - Figures 7a to 7g schematically show alternative embodiments of the lift structure of Figure 5 5 - Figures 8a to 8p schematically show alternative embodiments of the lift structure of Figure 6 - Figure 9 is a cross-sectional view of a wing of an aircraft according to the invention - the Figure 10 is an enlarged view of the area X of Figure 9 - Figure 11 is a longitudinal sectional view of a first embodiment of a damping device mounted on a wing of the aircraft according to the invention - Figure 12 is a schematic view of a second embodiment of a damping device mounted on a wing of the aircraft according to the invention; and - Figure 13 is a schematic view in partial section of a shoring device provided between the upper and lower wings of an aircraft according to the invention.

The exemplary embodiments shown in the figures relate more particularly to aircraft without autonomous propulsion means, such as gliders, and of the plane type known as duck (that is to say without empennage but comprising additional flaps or wings in the front). Of course, the invention is in no way limited to this type of aircraft, but is likely to apply to any aircraft, whether or not provided with propulsion means, with or without tailplane.

It will also be noted that, in the various embodiments, the identical or similar parts of the aircraft have the same references.

With particular reference to FIGS. 1 to 6, the aircraft 1 comprises, on either side of its fuselage 2, at least one lift structure 3,4 made up of at least two wings 5,6; 7.8. In the embodiments shown, the aircraft 1 is provided with two lift structures 3,4 each consisting of two wings 5,6; 7.8, respectively. it is understood that one can conceive, in the same way, aircraft provided with several support structures comprising, possibly, more than two wings.

The distal ends of the wings 5,6; 7,8 are connected, either directly to each other, or to at least one wing element 9,10 provided between them (embodiment illustrated in FIGS. 3,4 and 6, in particular). Each lift structure 3,4 forms, in front view, with the sides of the fuselage 2 a closed polygon, and has an upper wing 5.7 and a lower wing 6.8, the attachment zones of the upper and lower wings being optionally offset in the direction of the aerodynamic flow as is clearly visible in Figures 2 and 4. As shown in these figures, the upper wings 5,7 have attachment points to the fuselage located forward 6.8 lower wing attachment points; but the opposite can also be considered.

The moment of inertia, as defined above, and the lift each take a maximum value in the central zone of the wing considered in the longitudinal direction, in particular in the central third of the wing (interval A represented on (Figures 5 and 6). An adequate distribution is thus obtained, both for the recovery of the stresses to which the wing is subjected, and as regards the lift characteristics.

As. illustrated, the aircraft 1 comprises, at the front of the fuselage, at least one duck plane 11 replacing the conventional tailplane.

In the case where the upper and lower wings of each lift structure are connected by a canopy element 9, 10, it is necessary to ensure the rigidity of each lift structure to provide, between the upper and lower wings, shrouds 12 (embodiment of Figures 3 and 4-). Such shrouds are not necessary in the case of a three-point lift structure, as illustrated in FIGS. 1 and 2, said three-point lift structure being self-supporting.

Attaching the wings 5.6; 7,8 to each other, to the fuselage 2 of the aircraft I and, optionally, to the wing element 9,10 is produced in an articulated manner in rotation in a plane at least substantially transverse to the direction of the aerodynamic flow. At least some of these joints may not be completely free, that is to say semi-rigid.

Referring to Figures 1 and 2, illustrating a first embodiment of the invention, the joints between the wings 5,6; 7.8, respectively, are represented by the references 17; 18, respectively, while the articulations of the wings to the fuselage 2 are represented by the references 13,15; 14.16, respectively.

In the case of the embodiment illustrated in Figures 3 and 4, the wings 5,6; 7.8, respectively, are connected to the wing elements 9, 10, respectively, by the joints 23, 25; 24.26, respectively. These same wings 5,6; 7.8, respectively, are connected to the fuselage 2 by the joints 19,21; 20.22, respectively.

According to an important characteristic of the invention, at least one axis of the aforementioned articulations is situated inside the zone defined by the main planes of inertia, parallel to the wing, of the upper and lower wings. In what follows, reference will be made more particularly to FIGS. 5 and 6. These figures, for each of the embodiments illustrated by FIGS. 1, 2 and 3, respectively, represent, schematically, a support structure (with by way of example, the structure 3) comprising an upper wing 5 and a lower wing 6 and in which an axis of articulation is situated inside the zone defined by the main planes of inertia, parallel to the wing , the upper and lower wings. In the case of FIG. 5, this axis of articulation is the axis 17 of the articulation connecting the distal ends of the upper 5 and lower wings 6. In the case of FIG. 6, l 'axis of articulation thus offset (relative to the main planes of inertia P and P') is the axis of articulation 19 connecting the upper wing 5 to the fuselage (not shown). The main planes of inertia P and P 'parallel to the wings, correspond to the median cords of the two profiles of the wings.

As illustrated by FIGS. 7a to 7g and 8a to 8p, different scenarios are possible, including the offset of a single articulation axis, either of the wings between them, and of the wings to the fuselage or, possibly, to the wing element 9,10 intermediate, or the offset of two, three or even four (in the case of the embodiment of Figures 3,4 and 6) hinge axes. In this regard, note the presence of a damping device 27, in the case of FIG. 6, between the offset articulation axis 19 and the lower wing 6.

Thus, in all cases, taking into account the compression / tension forces generated by the joints, the bending moments introduced by the offset of the axis (or axes) make it possible to induce deformations D, D 'in the same direction for the upper 5 and lower 6 wings in the case of three points of articulation (FIG. 5), and in the opposite direction in the case of four points of articulation (FIG. 6). In FIGS. 5 and 6, the deformed shapes D, D ′ are shown for positive forces. In the case of negative forces, these deformations are in the opposite direction.

The value of the axis offset (i.e. the fact, for an articulation axis, of being in the zone between the main planes of inertia parallel to the wings) can be arbitrary, but it can be optimized as a function of the reciprocal values of the flexion modules of the upper and lower wings.

Reference will be made in the following, more particularly, to FIGS. 9 to 12.

Another important characteristic of the present invention is to provide the lift structures 3, 4 with at least one stiffening device 28, making it possible to absorb, in highly stressed places, stresses clearly greater than those absorbed by the part corresponding wings.

In the figures, there is shown, by way of example, a stiffening device 28 provided on the upper wing 5 of the lifting structure 3. However, it is understood that such a stiffening device can be provided in all the wings of the aircraft, and this, for each embodiment of the aircraft.

With particular reference to FIG. 9, the stiffening device 28 comprises, in the vicinity of the upper 5a and lower 5b surfaces of the wing 5, in the central area of said surfaces, at least one bundle 29 of materials, the module of which of elasticity is greater (and preferably much greater) than that of the material constituting the surface of the wing, and which extends in the longitudinal direction of the wing.

Indeed, either a section of a wing, as illustrated in FIG. 9, the parts of the section most stressed, in the case of transmission of forces, are located on its periphery, in particular in the central part. It will then be understood that the desired stiffening can be obtained, in particular, by introducing four aforementioned beams arranged, two by two, on the upper surfaces 5a, respectively lower 5b, of the wing 5, symmetrically with respect to the longitudinal plane of inertia of the wing, parallel to the wing (the trace of which on the plane of the figure is represented by the axis X, X ') and with respect to the longitudinal plane of inertia of the wing, perpendicular to the wing ( whose trace on the plane of the figure is represented by the axis Z, Z '); i.e. symmetrically with respect to the center of inertia 0. Thanks to this symmetry, it is therefore possible to stiffen the wing in the direction, at the same time, of its inertia with respect to the X axis, X 'and with respect to the longitudinal central axis of the wing.

The particular structure of the beams 29 will now be described with reference to FIG. 10. It being understood that the beam 29 in itself has a modulus of elasticity much greater than that of the wing itself, and is in particular made of synthetic fibers unidirectional or stranded, such as for example carbon fibers, the bundle 29 can be housed in a tube 30, in particular metallic, consisting for example of an aluminum or titanium alloy. The tube 30 is intended to establish the transmission of forces along the longitudinal axis and to make the connection between the beam 29 and the wing 5, the structure of which can in particular be made from agglomerated synthetic materials. made in particular of resin filled with glass fibers, is intended to make the tube 30 integral with the wing 5 while allowing a small differential expansion.

There is also provided a connecting tube 31, disposed between the bundle 29 and the internal wall of the housing tube 30 and allowing the beam 29 to slide on the walls of the tube 30 while avoiding the cutting of the walls of the tube by the fibers of the beam. Preferably, this connecting tube 31 forms an integral part of the bundle 29.

In the case where bundles with stranded fibers are used, it suffices, in general, to modulate the level of load of the prestressing at the start in order to adjust the damping capacity of the airfoil, and thus the capacity of self amortization may be sufficient without additional device.

However, when using unidirectional fiber bundles, an additional damping device must be considered. In this case, it will always be advantageous to prestress the beams to their optimal values so that there is no (or little) introduction of parasitic resonances during the damping of the forces in cyclic regime.

With particular reference to FIGS. 11 and 12, it suffices to make integral, at one of their ends, the bundles 29 of a damping device 33, 34, in particular of the piston-cylinder type , preferably of the oleo-pneumatic type with variable flexibility. Two types of anchoring of the piston-cylinder devices 33, 34 can be envisaged. In a first case, illustrated by FIG. 11, the piston 35 is integral with the bundle 29 while the cylinder 37 comes to bear at the end of the wing. This device is especially advantageous in the case of the embodiment of the aircraft with three points of articulation.

In the case illustrated in FIG. 12, the piston 36 is secured to the bundle 29, while the cylinder 38 is secured to the frame 39 supporting the end articulation axis (13 for example). In this case, the damping is linked to a corresponding damping of the rotations of the axes.

Reference will now be made more particularly to FIG. 13 in which only the central zone of a support structure of the aircraft has been shown. At least substantially in the median transverse plane perpendicular to the wings 5, 6 of the upper wing - lower wing assembly, at least one shoring device 40 is provided connecting the upper 5 and lower wings 6. The shoring device 40 comprises at least one bar 41, in particular a metal bar, optionally provided with a damping device 42 such as, for example, a piston-cylinder assembly, preferably of the oleopneumatic type with variable flexibility.

As shown diagrammatically in FIG. 6, the shoring device 40 can also include at least one flexible bar 48 by construction or initially arched.

Each bar is connected, at each end, to a box 43 provided in the corresponding wing 5,6 and constituted by two symmetrical spaced half-boxes, formed by spaced ribs, in particular metallic, internal 45 and external 44, the space between it is filled with a material 46, in particular synthetic. The internal ribs 45 protrude from the surface of the wing so as to allow the connection, in particular by a joint 47, with the bar. It will be noted that the fixing position of the shoring device can be moved, depending on the particular construction forms of the lifting structures, along the width of the wing, in particular to obtain a clamping effect on the trailing edges.

The main action of this shoring device will be to limit the deformation of the compressed wing by flattening its central zone, and, on the contrary, arching the central zone of the stretched wing. Everything happens as if this shoring device made it possible to transfer part of the elastic energy accumulated, during deformation, from the compressed wing to the stretched wing.

Finally, it should be noted that the wings, and in particular their coating, can be made from an extruded aluminum alloy profile, with a constant profile, as is, for example, described in French patent application No. 83 08725; the interior parts of the hollow profile thus formed possibly being stiffened by composite materials with a honeycomb structure.

Claims (14)

1 - Aircraft, of the type comprising, on either side of its fuselage, at least one lift structure consisting of at least two wings, the distal ends of which are connected, either to each other, or to at least one wing element provided between them, each structure forming, in front view, with the sides of the fuselage a closed polygon, and comprising an upper wing and a lower wing, the attachment zones of the upper and lower wings being optionally offset in the direction of the aerodynamic flow, the values of the moment of inertia and of the lift each taking a maximum value in the central zone of the wing considered in the longitudinal direction, the attachment of the wings to each other, to the fuselage and, optionally, to said element of airfoil being produced in an articulated manner in rotation in a plane at least substantially transverse to the direction of the aerodynamic flow, characterized in that at least one articulation axis (13-18 19-26) is located inside the area defined by the main planes of inertia, parallel to the wing, upper (5.7) and lower (6.8) wings. 2 - aircraft according to claim 1, of the type in which each lifting structure consists of two upper and lower wings whose ends remote from the fuselage are, directly, interconnected in an articulated manner, so as to constitute, substantially, a triangle articulated, characterized in that the axis (17,18) of the articulation between said ends of the upper (5,7) and lower (6,8) wings is situated inside the zone defined by the planes d 'main inertia, parallel to the wing, upper and lower wings. 3 - aircraft according to claim 1, of the type in which each lift structure consists of two upper and lower wings connected, at their ends remote from the fuselage, to a wing element in an articulated manner, so as to constitute, substantially, a articulated trapezium, characterized in that at least one axis (23,25); (24,26) of the articulation to said airfoil element (9,10) is located in the zone defined by the main planes of inertia, parallel to the wing, upper wings (5,7) and lower wings (6 , 8). 4 - Aircraft according to any one of claims 1 to 3, characterized in that at least one axis (13-16; 19-22) of articulation of the upper (5.7) and lower (6.8) wings fuselage (2) is located in the area between the main planes of inertia, parallel to the wing, upper and lower wings. 5 - An aircraft according to any one of claims 1 to 4, characterized in that at least some of the joints (13-18); (19-26) above are not completely free. 6 - An aircraft according to any one of claims 1 to 5, characterized in that the aforementioned lift structures (3,4) comprise at least one stiffening device (28). 7 - aircraft according to claim 6, characterized in that the stiffening device (28) comprises, in the vicinity of the upper (5a) and lower (5b) surfaces of each wing (5), in the central area of said surfaces, at least a bundle (29) of materials, the modulus of elasticity of which is greater than that of the material constituting the surface of the wing, and which extends in the longitudinal direction of the wing. 8 - An aircraft according to claim 7, characterized in that the aforementioned bundle (29), with unidirectional fibers or stranded, is housed in a tube (30), in particular metallic, integral with the wing; a connecting tube (31) preferably forming an integral part of the bundle (29) possibly being provided between the bundle (29) and the internal wall of the tube (30) serving to accommodate the bundle. 9 - aircraft according to claim 7 or claim 8, characterized in that the stiffening device (28) comprises four aforementioned beams (29) arranged, two by two, on the upper surfaces (5a), respectively lower (5b), of the wing (5) symmetrically with respect to the longitudinal plane of inertia of the wing, parallel to the wing, and with respect to the longitudinal plane of inertia of the wing, perpendicular to the wing. 10 - aircraft according to any one of claims 7 to 9, characterized in that each beam (29) above comprises at least one damping device (33,34), in particular of the piston-cylinder type, including the piston (35 , 36) is integral with the bundle (29) and of which, in a first case, the cylinder (37) comes to bear at the end of the wing; and, in a second case, the cylinder (38) is integral with the frame (39) supporting the end hinge pin. 11 - Aircraft according to any one of claims 1 to 10, characterized in that at least substantially in the median transverse plane, perpendicular to the wings (5,6), of the upper wing - lower wing assembly, provision is made for less a shoring device (40) connecting the upper and lower wings. 12 - Aircraft according to claim 11, characterized in that the shoring device (40) comprises at least one bar (41), in particular metallic, possibly provided with a damping device (42) such as, for example, a piston-cylinder assembly. 13 - Aircraft according to claim 11, characterized in that the shoring device (40) comprises at least one bar (48) flexible by construction or initially arched. 14 - Aircraft according to claim 12 or claim 13, characterized in that each aforementioned bar is connected, at each end, to a box (43) provided in the corresponding wing and consisting of two half-boxes spaced symmetrical, and formed spaced ribs, in particular metallic, internal (45) and external (44) whose space (46) between them is filled with a material, in particular synthetic, the internal ribs (45) projecting from the surface of the wing so as to allow connection with the bar.