Abstract
A wing segment for an aircraft includes a first skin element with a material having anisotropic characteristics and a rib element arranged between leading and trailing edges of the wing segment. The first skin element is attached to the rib element such that deformation of the skin element results in a twist of the wing segment with respect to a main extension direction of the wing segment. The wing segment further includes a control unit to control deformation of the skin element in order to adjust the twist of the wing segment to achieve a predetermined twist of the wing segment.
Claims
1. A wing segment for an aircraft, comprising: a first skin element comprising a material having anisotropic material characteristics; a rib element arranged between a leading edge and a trailing edge of the wing segment, wherein the first skin element is attached to the rib element such that a deformation of the first skin element results in a twist of the wing segment with respect to a main extension direction of the wing segment; and a control unit to control a deformation of the first skin element to adjust the twist of the wing segment such that a predetermined twist of the wing segment is achievable.
2. The wing segment according to claim 1, the first skin element comprising a material with a main fiber direction, wherein stiffness of the first skin element in the main fiber direction is higher than stiffness of the skin element in a direction different to the main fiber direction.
3. The wing segment according to claim 2, wherein: the main fiber direction of the first skin element is angled with respect to the main extension direction of the wing segment such that the angle between the main fiber direction of the first skin element and the main extension direction of the wing segment is between 0 and 90; and/or the main fiber direction of the first skin element is angled with respect to the main extension direction of the wing segment such that the angle between the main fiber direction of the first skin element and the main extension direction of the wing segment is between 0 and 90.
4. The wing segment according to claim 1, wherein the first skin element comprises a plurality of skin layers, each of the skin layers having a corresponding fiber direction.
5. The wing segment according to claim 4, wherein the plurality of skin layers together forms the first skin element such that the main fiber direction of the first skin element is defined by the corresponding fiber direction which is present in the majority of skin layers.
6. The wing segment according to claim 4, wherein: a first number of skin layers of the plurality of skin layers has a first fiber direction and a second number of skin layers of the plurality of skin layers has a second fiber direction; an angle +0 between the first fiber direction and the main extension direction of the wing segment is between 0 and 90; an angle between the second fiber direction and the main extension direction of the wing segment is between 0 and 90; and the second number of skin layers comprises more skin layers than the first number of skin layers.
7. The wing segment according to claim 1, the control unit controls the deformation of the first skin element by actively or passively inducing a force into the first skin element.
8. The wing segment according to claim 7, wherein: the control unit controls the deformation of the first skin element by a stiffening element within the wing segment; and/or the control unit controls the deformation of the first skin element by actively controlling an actuation unit within the wing segment.
9. The wing segment according to claim 1, wherein: the control unit comprises at least one stiffening element arranged at the rib element such that a predetermined deformation of the rib element is generated to adjust the twist of the wing segment; and/or the control unit comprises at least one actuation unit to adjust a predetermined deformation of the rib element to adjust the twist of the wing segment.
10. The wing segment according to claim 1, further comprising a second skin element comprising a material having anisotropic material characteristics, wherein the second skin element is attached to the rib element opposite to the first skin element.
11. An aircraft comprising: a wing segment that comprises: a first skin element comprising a material having anisotropic material characteristics; a rib element arranged between a leading edge and a trailing edge of the wing segment, wherein the first skin element is attached to the rib element such that a deformation of the first skin element results in a twist of the wing segment with respect to a main extension direction of the wing segment; and a control unit to control a deformation of the first skin element to adjust the twist of the wing segment such that a predetermined twist of the wing segment is achievable.
Description
BRIEF DESCRIPTION OF THE DRAWINGS
[0070] A more complete understanding of the subject matter may be derived by referring to the detailed description and claims when considered in conjunction with the following figures, wherein like reference numbers refer to similar elements throughout the figures.
[0071] FIG. 1 schematically shows an aircraft having a wing segment according to an exemplary embodiment of the invention.
[0072] FIG. 2 schematically shows a cross-section through a wing segment according to an exemplary embodiment of the invention.
[0073] FIG. 3A schematically shows a first skin element, e.g. a part of an upper cover of a wing segment and a second skin element, e.g. a part of a lower cover of a wing segment according to an exemplary embodiment of the invention.
[0074] FIG. 3B schematically shows a first or second skin element according to an exemplary embodiment of the invention.
[0075] FIG. 4 schematically shows cross-sections through a wing segment comparing a jig-shape twist and a cruise shape twist according to an exemplary embodiment of the invention.
[0076] FIG. 5 schematically shows a relative movement of an upper cover with respect to a lower cover of the wing segment according to an exemplary embodiment of the invention.
[0077] FIG. 6 shows a numerical model of a rib element according to an exemplary embodiment of the invention.
[0078] FIG. 7A shows an arrangement of rib elements in a loaded aircraft wing with balanced skin laminates according to an exemplary embodiment of the invention.
[0079] FIG. 7B shows an arrangement of rib elements in a loaded aircraft wing with unbalanced skin laminates according to an exemplary embodiment of the invention.
[0080] FIG. 8A schematically shows a control unit for passively controlling a deformation of a rib element according to an exemplary embodiment of the invention.
[0081] FIG. 8B schematically shows a control unit for passively controlling a deformation of a rib element according to another exemplary embodiment of the invention.
[0082] FIG. 9A schematically shows a control unit for actively controlling a deformation of a rib element according to an exemplary embodiment of the invention.
[0083] FIG. 9B schematically shows a control unit for actively controlling a deformation of a rib element according to another exemplary embodiment of the invention.
[0084] FIG. 10A schematically shows a top view and a bottom view of a wing segment according to an exemplary embodiment of the invention.
[0085] FIG. 10B schematically shows a top view and a bottom view of a wing segment according to another exemplary embodiment of the invention.
[0086] FIG. 10C schematically shows a top view and a bottom view of a wing segment according to another exemplary embodiment of the invention.
[0087] FIG. 10D schematically shows a top view and a bottom view of a wing segment according to another exemplary embodiment of the invention.
[0088] FIG. 11A shows a wing segment which is loaded in up bending according to an exemplary embodiment of the invention.
[0089] FIG. 11B shows a wing segment which is loaded in up bending according to another exemplary embodiment of the invention.
DETAILED DESCRIPTION
[0090] The following detailed description is merely illustrative in nature and is not intended to limit the embodiments of the subject matter or the application and uses of such embodiments. As used herein, the word exemplary means serving as an example, instance, or illustration. Any implementation described herein as exemplary is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, there is no intention to be bound by any expressed or implied theory presented in the preceding technical field, background, brief summary or the following detailed description.
[0091] FIG. 1 shows an aircraft 100 having a wing 1 with a wing segment 10. The wing 1 may comprise several wing segments 10. The wing segment 10 comprises stringer elements 11 extending along a main extension direction 12 of the wing segment 10, a rib element attached to at least one stringer element 11 and extending between a leading edge 14 and a trailing edge 15 of the wing segment 10. For reasons of clarity, the rib elements are not shown in FIG. 1. The wing segment 10 further comprises a first skin element 20 comprising a material having anisotropic material characteristics. The first skin element 20 is attached to the rib element in such a way that a deformation of the first skin element 20 results in a twist of the wing segment 10 with respect to the main extension direction 12 of the wing segment 10. It is noted that the stringer elements 11 are depicted in FIG. 1 for a better understanding although the skin element 20 is usually not transparent and the stringer elements should therefore be hidden by the skin element 20. The wing segment 10 comprises a control unit 30 for controlling a deformation of the first skin element 20 in order to adjust the twist 40 of the wing segment 10 such that a predetermined twist 40 of the wing segment 10 is achievable.
[0092] In the example described in FIG. 1, the control unit 30 is arranged at or within a fuselage 110 of the aircraft 100. The aircraft fuselage 110 extends along a longitudinal axis 111 of the aircraft 100. The main extension direction 12 of the wing segment 10 and the longitudinal axis 111 of the aircraft fuselage 110 enclose an angle which is defined as the sweep angle of the aircraft wing 1 or wing segment 10.
[0093] FIG. 1 further shows the different fiber directions 12, 16, 21, 22 of different layers of the first skin element 20. The first fiber direction 21 is angled 45 with respect to the main extension direction 12 of the wing segment 10. The first fiber direction 21 may correspond to a first number of layers of the skin element 20. The second fiber direction 22 is angled 45 with respect to the main extension direction 12 of the wing segment 10. The second fiber direction 22 may correspond to a second number of layers of the skin element 20. The wing segment 10 may comprise a third number of layers which has a third fiber direction 12, e.g. a 0-fiber direction, which is parallel to the main extension direction 12 of the wing segment 10. The wing segment 10 may comprise a fourth number of layers which has a fourth fiber direction 16, e.g. a 90-fiber direction, which is perpendicular to the main extension direction 12 of the wing segment 10.
[0094] FIG. 2 shows a cross-section A-A through the wing segment 10 of FIG. 1. This is only a schematic view of the wing segment 10 since the wing segment 10 of an aircraft wing usually has a curved surface which is provided by the upper cover 20a, e.g. the first skin element 20a, and the lower cover 20b, e.g. the second skin element 20b as shown in FIG. 3A.
[0095] In FIG. 2, it can be recognized that the wing segment 10 comprises stringer elements 11 at each side of the wing segment 10, i.e. at the upper cover 20, 20a which is at least partially formed by the first skin element 20a and at the lower cover 20, 20b which is at least partially formed by the second skin element 20b. Although the examples described with respect to the Figures show wing segment 10 with stringer elements 11, it should be understood that the adjustment of the wing twist by means of the control unit in combination with unbalanced skin laminates also applies to the case in which the wing segment 10 has no stringer elements 10, for example in a sandwich construction of the wing segment 10.
[0096] FIG. 3A shows the curved surface of the upper cover 20a and lower cover 20b. Furthermore, FIG. 3A shows the first fiber direction 21 which is angled with respect to the main extension direction 12 of the wing segment 10 in any angle between 0 and 90, preferably +45. The first fiber direction 21 may correspond to a first number of layers of the skin element 20. The second fiber direction 22 is angled with respect to the main extension direction 12 of the wing segment 10 in any angle between 0 and 90, preferably 45. The second fiber direction 22 may correspond to a second number of layers of the skin element 20.
[0097] The wing segment 10 comprises a third number of layers which has a third fiber direction 12 which is parallel to the main extension direction 12 of the wing segment 10 and therefore parallel to the stringer elements 11. The wing segment 10 comprises a fourth number of layers which has a fourth fiber direction 16 which is perpendicular to the main extension direction 12 of the wing segment 10.
[0098] FIG. 3B shows a first skin element 20, 20a or a second skin element 20, 20b. The skin element 20 comprises a plurality of skin layers 23. Therefore, the skin element 20 can be imagined as a stack of layers 23 in which the different layers 23 are arranged one on top of the other. In FIG. 3B, an example for possible fiber directions of fibers within the different skin layers 23 are shown by the two-dimensional coordinate system depicted on the top layer 23 of the stack of layers 23. An angle between the first fiber direction 21 and the main extension direction 12 of the wing segment 10 may be between 0 and +90, for example about 45. An angle between the second fiber direction 22 and the main extension direction 12 of the wing segment 10 may be between 0 and 90, for example about 45. The second number 23b of layers 23 defined by the second fiber direction 22 is greater than the first number 23a of layers 23 defined by the first fiber direction 21. In this manner, an unbalanced skin laminate 20 can be provided.
[0099] The third fiber direction 12 may be parallel to the main extension direction 12 of the wing segment 10 such that the third fiber direction 12 may also be defined as 0-fiber direction. The fourth fiber direction 16 may be perpendicular to the main extension direction 12 of the wing segment 10 such that the fourth fiber direction 16 may also be defined as 90-fiber direction.
[0100] FIG. 4 shows a cross-section through a wing segment 10 for comparing a jig-shape twist and a cruise shape twist. The jig-shape twist is defined as the shape in which the wing segment 10 or wing 1 is built and represents a configuration of the wing segment 10 in an unloaded state. The jig-shape twist of the wing segment 10 is shown in the lower picture of FIG. 4. The cruise-shape twist which is depicted in the upper picture of FIG. 4 is defined as the wing segment's 10 deflected shape during a cruise flight condition in which the wing segment is loaded, i.e. bended. For example, the wing segment 10 is subjected to gravity loads, engine loads and aerodynamic loads. The gravity load is indicated by the lower arrow in FIG. 4 and lift forces are indicated by the upper arrows in FIG. 4.
[0101] As can be gathered from FIG. 4, the twist of the wing segment is higher in the loaded state, e.g. during cruise flight, than in the unloaded state, e.g. during ground operation. In particular, during cruise flight, the leading edge of the wing segment 10 twists downward compared to the jig-shape twist. In other words, the wing segment 10 twists nose down when loaded in bending.
[0102] FIG. 5 schematically shows a relative movement of the upper cover 20a with respect to the lower cover 20b of the wing segment 10. The unbalanced skin laminates which are represented by the first skin element 20a and the second skin element 20b are due to shear restrictions enforced by the rib elements 13, which produces a resulting torsion moment that twists the wing segment 10, for example the wing 1, a winglet or a horizontal or vertical tail plane when loaded, e.g. bended. The generated torsion moment is a result of an elastic equilibrium established between the unbalanced skin elements 20 forcing to shear forward and/or aft and the rib elements 13, e.g. the wing box ribs, prevent their motion. If the rib elements 13 have only low shear stiffness, the covers would be allowed to shear freely and would not generate any shear forces at the rib element 13 interfaces and hence no torsion. In contrast, a maximum torsion would be generated for fully rigid rib elements 13.
[0103] When using unbalanced skin laminates 20, a direct coupling exists between the rib element sizing and the global deflections of the wing segment 10. Compared to metallic wing designs and composite wing designs with balanced or symmetric skin laminates, a design with unbalanced skin laminates presents additional design opportunities. Changing the rib element stiffness has a significant effect on the wings bending and torsion coupling characteristics. Introducing additional stiffness into the rib elements 13 may, for example, be used to correct or adjust a non-optimum cruise wing angle of attack and thereby improve the aircraft performance. Angle of attack may be understood as the angle between the chord line of the wing's airfoil and the direction of the oncoming air.
[0104] FIG. 5 shows that the upper cover 20a, e.g. the first skin element 20, is shifted forward if the upper cover 20a is loaded, i.e. bended. The upper cover 20a is shown in the unloaded state 26 and in the loaded state 27, wherein a loading or bending of the upper cover 20a causes a shift forward along the longitudinal axis 111 of the upper cover 20a from the unloaded state 26 to the loaded state 27.
[0105] FIG. 5 also shows the that the lower cover 20b, e.g. the second skin element 20, is shifted backwards if the upper cover 20b is loaded, i.e. bended. The lower cover 20b is shown in the unloaded state 28 and in the loaded state 29, wherein a loading or bending of the lower cover 20b causes a shift backwards along the longitudinal axis 111 of the lower cover 20b from the unloaded state 28 to the loaded state 29.
[0106] This shifting of the covers 20a, 20b, e.g. skin elements 20, is caused by the specific arrangement of fiber directions 12, 16, 21, 22 in the different layers 23 of the skin elements 20 such that anisotropic characteristics of the skin elements are achieved. The skin elements 20 having such anisotropic material characteristics can be defined as unbalanced skin laminates. Modifying the bending and/or torsion characteristics for a composite wing, winglet and/or a horizontal or vertical tail plane may be achieved by providing such composite skins using unbalanced laminates.
[0107] In contrast, balanced/symmetric laminates may comprise laminates with only 0, 45 and 90-fiber directions having the 0-fibers extending along the main extension direction 12 of the stringer elements 11, wherein the number of skin layers with +45-fiber direction is equal to the number of skin layers with 45-fiber direction. With this orientation the 0-fibers in the skin elements 20 are aligned to the principle loading direction of the skin elements 20 and with an equal amount of +45/45-fibers in the skin elements 20. However, using such skin elements 20, wing level bending and torsion coupling effects, in addition to the already existing or system inherent twist effect due to geometrical constraints and the sweep angle, are usually not present.
[0108] By allowing an unbalancing of the +45 and 45-fibers in the skin elements 20, it is possible to introduce extension and shear coupling effects into the top and bottom covers 20a, 20b. Using unbalanced laminates, the top and bottom skin elements 20a, 20b of a composite wing segment 10, wing 1, winglet or a horizontal or vertical tail plane may, as an example, be designed such that the top skin element 20a shears forward during compression loads generated by an up-bend case, whilst the bottom skin 20b simultaneously shears backwards under the tension loads generated by the same up-bend case as shown in FIG. 5. As the two wing elements 20a, 20b are tied together by the wing rib elements 13, the shearing of the top skin element 20a and the bottom skin element 20b of the wing segment 10 is constrained. This means that the skin elements 20a, 20b are pulled back onto the rib elements 13 generating additional shear loads in the interface between the rib elements 13 and the skin elements 20a, 20b. As a net effect, the skin elements 20, 20a, 20b, e.g. the upper cover 20a and the lower cover 20b generate a twisting moment, which twists the wing segment 10, e.g. the wing box, nose down for an up-bend case. The opposite case occurs for a down-bending of the wing 1.
[0109] The general principle is illustrated in FIG. 5, which shows the shearing of the covers 20a, 20b in both plan view and side view.
[0110] Furthermore, FIG. 5 shows a load introduction via the rib element 13 into the skin elements 20a, 20b which is the result of the relative movement of the first skin element 20a, e.g. the upper cover 20a, and the second skin element 20b, e.g. the lower cover 20b, combined with the fixed attachment between the skin elements 20a, 20b and the rib element 13 preventing a relative movement between these components.
[0111] FIG. 6 shows a numerical model of a deformation of a rib element 13 as a result of the shear forces generated by the relative shifting of the upper or top cover 20a and lower or bottom cover 20b not shown in FIG. 6.
[0112] FIG. 7A shows the arrangement of rib elements 13 in a wing 1 or wing segment 10 during an up-bending case, which usually occurs during cruise flight. Bending forces 41 apply at the wing 1 or wing segment 10. The shown up-bending case refers to balanced skin laminates which are not shown in FIG. 7A.
[0113] FIG. 7B shows the arrangement of rib elements 13 in a wing 1 or wing segment 10 during an up-bending case, which usually occurs during cruise flight. The same bending forces 41 as in FIG. 7A apply at the wing 1 or wing segment 10. The shown up-bending case refers to unbalanced skin laminates which are not shown in FIG. 7A.
[0114] It is recognizable from FIGS. 7A and 7B that, using unbalanced skin laminates attached to the rib elements 13, cause an additional twist 40 on the wing 1 or wing segment 10. The unbalanced skin laminates, e.g. the skin elements having a different number of layers, for example, with +45 fiber direction 21 and layers with 45-fiber direction 22, cause this additional twist 40 of the wing 1 or wing segment 10.
[0115] FIG. 8A shows a passive control of the wing twist 40. In particular, FIG. 8A shows a rib element 13 having a control unit 30 which is configured for controlling the deformation of the rib element 13 by passively adjusting the introduction of a predetermined force into the rib element 13 such that in turn a predetermined compression or tension load can be introduced into the skin elements 20. The control unit 30 comprises at least one stiffening element 31, preferably two stiffening elements 31 which are arranged at the rib element 13 such that the deformation of the rib element 13 can be controlled in order to adjust the compression or tension load in the skin elements 20 and therefore the twist 40 of the wing segment 10. The stiffening elements 31 and the rib element 13 may be separate parts which are attached to each other. The stiffening elements 31 may be parallel to each other and extend along the lateral surfaces of the rib element 13. In particular, the stiffening elements 31 have a straight design like a beam.
[0116] By adding a stiffening element 31, the amount of chord-wise or transverse loading that is generated in the skin elements 20 during bending can be changed. This directly affects the shear deformation of the skin elements 20 and the twist of the wing segment 10 when it is bending. The passive control can also comprise a fixed pre-spanning of the stiffening element 31, which changes the twist on the wing segment 10 during ground operation. In this manner, it is possible to correct an already designed wing 1 or wing segment 10 that has an unfavorable twist. However, it is also possible to tune a wing 1 or wing segment 10 for different mission requirements.
[0117] FIG. 8B shows another example for a passive control of the wing twist 40. The stiffening elements 31 are arranged diagonally with respect to the rib element 13. This means that the stiffening elements 31 cross each other in a central portion of the rib element 13 which is substantially formed as a rectangle with outwardly curved edges as shown in FIG. 8B. The outwardly curved edges may align to the curved contours of the skin elements 20a, 20b.
[0118] FIG. 9A shows an example for an active control of the wing twist 40. In particular, FIG. 9A shows a rib element 13 having a control unit 30. The control unit 30 comprises an actuation unit 32 and stiffening elements 31 which are connected to the actuation unit 32 by electrical or signal lines. The control unit 30 is configured for controlling the deformation of the rib element 13 by actively adjusting the introduction of a force into the rib element 13 via the stiffening elements 31 such that in turn a predetermined compression or tension load can be introduced into the skin elements 20. The control unit 30 comprises at least one stiffening element 31, preferably two stiffening elements 31 which are arranged at the rib element 13 such that a predetermined deformation of the rib element 13 is generated in order to adjust the compression or tension load in the skin elements 20 and therefore to adjust the twist 40 of the wing segment 10. The stiffening elements 31 and the rib element 13 may be separate parts which are attached to each other. The stiffening elements 31 may be parallel to each other and extend along the lateral surfaces of the rib element 13. In particular, the stiffening elements 31 have a straight design like a beam. The stiffening elements 31 may, for example, be mechanically, hydraulically or magnetically actuated.
[0119] FIG. 9B shows another example for an active control of the wing twist 40. In particular, FIG. 9B shows a rib element 13 having a control unit 30. The control unit 30 comprises an actuation unit 32 and stiffening elements 31 which are connected to the actuation unit 32 by electrical or signal lines. The stiffening elements 31 and the rib element 13 may be separate parts which are attached to each other. In particular, the stiffening elements 31 have a straight design like a beam. The stiffening elements 31 may, for example, be mechanically, hydraulically or magnetically actuated. The stiffening elements 31 are arranged diagonally with respect to the rib element 13. This means that the stiffening elements 31 cross each other in a central portion of the rib element 13 which is substantially formed as a rectangle with outwardly curved edges as shown in FIG. 8B. The outwardly curved edges may align to the curved contours of the skin elements 20a, 20b.
[0120] FIGS. 10A to 10D illustrate top views and bottom views of the wing 1 or wing segment 10. In particular, these Figures show a possible combination of a control unit 30, e.g. an actuation unit 32 and a specified laminate definition, for example a specified main fiber direction or main stiffness direction, in the first skin element 20a, e.g. the top cover, and within the second skin element 20b, e.g. within the lower cover. The main fiber direction which may also be referred to as the main stiffness direction of the skin elements 20a, 20b is shown as three parallel lines in FIGS. 10A to 10D.
[0121] FIG. 10A shows a combination in which an actuation unit 32 introduces a transverse compression C into the top cover 20a whilst at the same time another actuation unit 32 introduces a transverse tension T into the bottom cover 20b. Actuation units 32 are aligned perpendicular to the main extension direction 12 of the wing segment 10 and therefore parallel to the rib elements 13. The introduction of tension and compression into the top and bottom covers 20a, 20b may generally be carried out by the same or separated actuation units 32. In case of separate actuation units 32, each actuation unit 32 may introduce the tension or compression independently of the other actuation unit 32.
[0122] In the shown configuration, to generate a twist, the top cover 20a and the bottom cover 20b comprise a similar direction of extension and shear coupling or of the main stiffness direction of the covers 20a, 20b as indicated in the Figures. The extension and shear coupling may be dependent on the fiber directions within the different skin layers 23 i.e. or it may depend on the main fiber direction of the skin elements. The opposite forces acting on the top cover 20a and bottom cover 20b create a differential shear of the two skin elements 20a, 20b and will twist the wing segment 20. As the extension and shear coupling is a directional property, it exists for a loading direction orthogonal to the wing main extension direction 12.
[0123] FIG. 10B shows a combination where actuation units 32 act in a similar direction, hence introducing either tension T or compression C into the top cover 20a and bottom cover 20b. Actuation units 32 are again aligned perpendicular to the main extension direction 12 of the wing segment 10 and therefore parallel to rib elements 13. In this configuration, to generate a twist, the top cover 20a and the bottom cover 20b comprise an opposite extension and shear coupling or an opposite main stiffness direction of the covers 20a, 20b as indicated by three parallel lines in the Figures. The similar direction of the actuation units 32 in combination with the opposite effect of unbalancing the skin elements 20a, 20b creates a differential shear of the two skin laminates 20a, 20b and will twist the wing segment 10.
[0124] FIG. 10C shows a combination where an actuation unit 32 introduces a transverse compression C into the top cover 20a whilst at the same time another actuation unit 32 introduces a transverse tension T into the bottom cover 20b. The actuators of the actuation units 32 are aligned diagonally across the rib bay, i.e. diagonally in the space between the rib elements 13. For example, the actuation unit extends from a front end of a rib element 13 to a rear end of a neighboring rib element 13. This means that the actuation units 32 extend parallel to the longitudinal direction of the fuselage of the aircraft. In this configuration, to generate a twist, skin laminates showing no extension shear along and perpendicular to the main extension direction of the wing segment 10 are used. In this case, the laminate is loaded at an angle to the main extension direction 12. This creates a differential shear of the two skins and twist the wing. It should be understood that the embodiments described herein may also be combined with metallic skin elements, e.g. metallic top and bottom covers.
[0125] FIG. 10D shows a combination where an actuation unit 32 introduces a transverse compression C into the top cover 20a whilst at the same time another actuation unit 32 introduces a transverse tension T into the bottom cover 20b. The actuators of the actuation units 32 are aligned diagonally across the rib bay, i.e. in the space between the rib elements 13. In particular, the actuation units 32 extend parallel to the longitudinal axis of the aircraft fuselage or in the flight direction. In this configuration, to generate a twist, the top cover 20a and the bottom cover 20b comprise a similar direction of extension and shear coupling as indicated by three parallel lines in the Figures. The opposite forces acting on the top cover 20a and bottom cover 20b create a differential shear of the two skin elements 20a, 20b and will twist the wing segment 20.
[0126] FIG. 11A shows a wing segment 10 or a wing 1 which is loaded in up bending, i.e. by a force F. The material on the top cover 20a is loaded in compression C and expands in a transverse direction. Since material cannot expand freely, a bi-axial compressive state is created. To add an independent control of the wing twist, three different actuation and control mechanisms are possible.
[0127] Therefore, FIG. 11A shows three different control units 30. The first example shows a control unit 30 in the form of a stiffener 31 which is attached to the rib element 13. The additional stiffener 31 increases the bi-axial compression C in the top skin element 20a, wherein a larger transverse compression occurs. An opposite effect occurs in the bottom skin 20b. Combining this with a laminate definition as shown in FIG. 10A, a predetermined twist of the wing 1 during flight can be adjusted.
[0128] The second example shows the control unit 30 in the form of a stiffener 31, which in contrast to the first example is pre-spanned. In addition to the stiffness effect described above, this allows to further add a static pre-tension or pre-compression into the top cover 20a and/or bottom cover 20b. This may change the twist on the wing segment 10 during ground operation and therefore may also have an influence of the twist and shape of the wing segment during flight. In other words, an unfavorable twist of an already designed wing 1 or wing segment 10 can be corrected or the twist of a wing 1 or wing segment 10 can be adapted for different mission requirements.
[0129] The third example shows the control unit 30 having an actuation unit 32 or actuator mechanism attached to the wing rib element 13 to introduce an opposite tension and compression load into the top and bottom. In particular, this makes it possible to introduce a compression C into the upper cover 20a and a tension T into the lower cover 20b or vice versa.
[0130] FIG. 11B shows a wing segment 10 or a wing 1 which is loaded in up bending, i.e. by a force F. The material on the top cover 20a is loaded in compression T and expands in a transverse direction. Since the material cannot expand freely, a bi-axial compressive state is created. To add an independent control of the wing twist, an actuator 32 is arranged between rib elements 13 of the wing segment 10. The actuator 32 is connected via two pin/lug connections at the top and at the bottom corners of a front and rear spar within the wing segment 10. This illustrates a case in which the skin forces are not introduced via the rib elements 13 as shown in FIG. 11A.
[0131] While embodiments of the invention have been illustrated and described in detail in the drawings and the foregoing description, such illustration and description are to be considered illustrative and exemplary and not restrictive; the subject matter is not limited to the disclosed embodiments. Other variations to the disclosed embodiments can be understood and effected by those skilled in the art and practicing the claimed invention, from a study of the drawings, the disclosure, and the appended claims. In the claims, the term comprising does not exclude other elements, and the indefinite article a or an does not exclude a plurality. The mere fact that certain measures are recited in mutually different dependent claims does not indicate that a combination of these measures cannot be used to advantage. Any reference signs in the claims should not be construed as limiting the scope of protection.
