• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Disclosed is an aircraft structure (11) for flow control, comprising a perforated panel (13) having an inner surface (15) directed to a structure interior (17), an outer surface (19) in contact with an ambient flow (21), and a plurality of micro pores (23) connecting the inner and outer surfaces (15, 19). The object to reduce the weight of the aircraft structure while maintaining the required fatigue strength, is achieved in that one or more elongate crack stopper elements (25) are attached to the inner surface (15) of the perforated panel (13), and the crack stopper elements (25) are configured to inhibit crack propagation within the perforated panel (13).
    公开号:EP3702263A1
    申请号:EP20153868.3
    申请日:2020-01-27
    公开日:2020-09-02
    发明作者:Dort DAANDELS；Bernhard Schlipf；Christian Heck
    申请人:Airbus Operations GmbH；
    IPC主号:B64C1-00
    专利说明:
    [0001] The present invention relates to an aircraft structure configured for flow control, preferably configured for hybrid laminar flow control (HLFC). A further aspect of the invention relates to an aircraft comprising such an aircraft structure.
    [0002] The aircraft structure comprises a perforated panel having an inner surface directed to a structure interior, an outer surface in contact with an ambient flow, and a plurality of micro pores distributed over the perforated panel and connecting the inner and outer surfaces. The perforated panel is preferably formed of a metal material, such as titanium.
    [0003] Such aircraft structures are known in connection with HLFC systems where air is sucked in or blown out through the micro pores in the perforated panel to advantageously influence the ambient flow along the aircraft structure to increase lift and reduce drag, and as a consequence reduce fuel consumption. The micro pores necessary for sucking in or blowing out air, however, decreases the fatigue strength of the perforated panel as crack initiation and propagation through the perforated panel is simplified by the micro pores. Thus, a higher panel thickness is required to obtain the required fatigue strength, which however adds to the weight of the aircraft structure.
    [0004] Accordingly, the object of the present invention is to reduce the weight of the aircraft structure while maintaining the required fatigue strength.
    [0005] This object is achieved in that one or more elongate crack stopper elements are attached to the inner surface of the perforated panel. Preferably, a plurality of crack stopper elements are provided that extend in parallel to one another. The crack stopper elements are preferably bonded to the inner surface of the perforated panel, but might also be bolted or riveted thereto. The crack stopper elements are configured to inhibit crack propagation within the perforated panel past the crack stopper element, i.e. from one side of a crack stopper element to the other side. Inhibiting crack propagation means either to stop or at least to decelerate crack propagation. It might also mean that crack propagation is deflected in an uncritical direction in parallel to the crack stopper elements.
    [0006] By such crack stopper elements crack propagation, specifically with respect to a direction transverse the crack stopper elements, can be stopped or at least decelerated, so that the fatigue strength of the perforated panel and thus of the aircraft structure is increased. In other words, the thickness of the perforated panel can be reduced without reducing the fatigue strength of the perforated panel at the same time, which in turn leads to a weight reduction of the aircraft structure, and consequently to a reduced fuel consumption of the related aircraft.
    [0007] According to a preferred embodiment, the crack stopper elements extend, longitudinally, in a direction transverse to a direction of flight of an associated aircraft. Preferably, the crack stopper elements extend within an angle range of 0° to 45°, further preferred 0° to 30°, most preferred 0° to 15° relative to the main load direction, preferably to the span direction. In such a way, the crack stopper elements also extend transversely, preferably perpendicularly, to the expected direction of crack propagation.
    [0008] According to another preferred embodiment, the crack stopper elements have considerably higher fatigue strength than the perforated panel, in particular than the related adjacent parts of the perforated panel, specifically with respect to loads applied in the longitudinal direction of the crack stopper elements. In such a way, cracks propagating in the perforated panel towards crack stopper elements will be stopped, deflected in a direction parallel to the crack stopper elements, or at least decelerated due to the higher fatigue strength of the crack stopper elements.
    [0009] According to yet another preferred embodiment, the crack stopper elements are formed as strips of fiber reinforced plastic (FRP) material. In case of the perforated panel made of a metal material a local fiber metal laminate (FML) is thereby formed by the FRP strips together with the perforated panel in the area of the strips. The volume ratio of metal and FRP might be e.g. 1/1 or similar. Such FML due to its high crack growth resistance is specifically suitable for crack stopping.
    [0010] According to a further preferred embodiment, the crack stopper elements have a width, perpendicular to the longitudinal extension and along the inner surface of the perforated panel, ranging from 1/100 to 1/1, preferably 1/25 to 1/15, most preferred 1/20, of the distance between each two adjacent crack stopper elements. A lower width of the crack stopper elements might be advantageous as less micro pores are covered by the crack stopper elements, while a higher width might provide more efficient crack stopping. Thus, the width is ideally adapted depending on the requirements of the specific application.
    [0011] According to a further preferred embodiment, the micro pores in the perforated panel extend also through the crack stopper elements. In such a way, the micro pores are not covered by the crack stopper elements and air might be sucked in and blown out through the micro pores in a homogeneous way. This is specifically advantageous in the case of crack stopper elements having a higher width, where consequently a higher number of micro pores would be covered. However, non-perforated crack stopper elements might also be preferred due to more simple production and higher strength, specifically in the case of lower width crack stopper elements where only few micro pores are covered.
    [0012] In another preferred embodiment, the aircraft structure further comprises an inner panel mounted to the perforated panel via stiffeners that are attached to the inner surface of the perforated panel, e.g. by bonding. The inner panel and the stiffeners are preferably made of a FRP material. The stiffeners might be part of the inner panel or might be separate parts attached to the outer side of the inner panel, e.g. by co-curing, molding, or bonding. For example, the stiffeners might have Z-profiles mounted to both the inner panel and the perforated panel, or might have Ω-profiles or trapezoid-shaped profiles that are integral with or mounted to the inner panel and mounted to the perforated panel. By adding such an inner panel a double-walled, stiffener reinforced aircraft structure is formed, which is specifically advantageous for use in flow control systems as the outer perforated panel is additionally supported.
    [0013] In particular, it is preferred that at least some of the crack stopper elements are provided between the inner surface of the perforated panel and at least some of the stiffeners. For example, crack stopper elements might be provided at each stiffener or might be provided only at some of the stiffeners, e.g. at every other stiffener. Additionally, crack stopper elements might be attached to the inner surface of the perforated panel at locations between the stiffeners, e.g. centrally between each two adjacent stiffeners. In such a way, the stiffeners can be used to fix the crack stopper elements to the perforated panel and no additional bonding or fixation needs to be applied for the related crack stopper elements. Additionally or advantageously, it is preferred that at least some of the stiffeners are formed as crack stopper elements. In such a way, no or less separate crack stopper elements need to be provided and fixed to the perforated panel.
    [0014] In particular, it is preferred that the stiffeners are formed as crack stopper elements by the material of the stiffeners being configured for crack stopping. Specifically, the material of the stiffeners is selected such that the stiffeners have higher fatigue strength than the perforated panel, in particular than the portion of the perforated panel to which the respective stiffeners are attached. In such a way, crack stopping is provided only by respectively adjusting the material of the corresponding stiffeners, which are provided, anyway. Separate crack stopper elements can be saved.
    [0015] Additionally or alternatively, it is preferred that the stiffeners are formed as crack stopper elements by the shape, in particular the profile, of the stiffener being configured for crack stopping. Specifically, the shape of the stiffeners is selected such that the stiffeners have higher fatigue strength than the perforated panel, in particular than the portion of the perforated panel to which the respective stiffeners are attached. In such a way, crack stopping is provided only by respectively adjusting the shape of the corresponding stiffeners, which are provided, anyway. Separate crack stopper elements can be saved.
    [0016] In particular, it is preferred that the stiffeners have an increased thickness at least at the portion attached to the inner surface of the perforated panel. Such an increased thickness of the stiffeners provides increased fatigue strength that serves to stop or at least decelerate crack propagation within the perforated panel.
    [0017] A further aspect of the present invention relates to an aircraft. The aircraft comprises a fuselage, wings, a vertical tail plane, and a horizontal tail plane. The aircraft structure according to any of the embodiments described above is arranged at the wings and/or at the vertical tail plane and/or at the horizontal tail plane of the aircraft. The features and advantages described above in connection with the aircraft structure apply vis-à-vis to the aircraft.
    [0018] Hereinafter, preferred embodiments of the present invention are described in more detail by means of a drawing. The drawing shows in Fig. 1 a perspective view of an aircraft according to the invention, Fig. 2 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of low width FRP strips, Fig. 3 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of high width FRP strips, Fig. 4 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged between Ω-stiffeners and the perforated panel, Fig. 5 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged between Z-stiffeners and the perforated panel, Fig. 6 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged in connection with the stiffeners and in addition between each two adjacent stiffeners, Fig. 7 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged in connection with every other stiffener, Fig. 8 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of stiffeners, wherein every other stiffener has a specific material adapted for crack stopping, Fig. 9 a schematic illustration (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of stiffeners, wherein every other stiffener has a thickened shape adapted for crack stopping.
    [0019] In Fig 1 an embodiment of an aircraft 1 according to the present invention is shown. The aircraft 1 comprises a fuselage 3, wings 5, a vertical tail plane 7, and a horizontal tail plane 9. At the wings 3, at the vertical tail plane 7, and at the horizontal tail plane 9 the aircraft 1 comprises an aircraft structure 11 according to any of the embodiments described hereinafter.
    [0020] In Fig. 2 a first embodiment of the aircraft structure 11 according to the invention is shown. The aircraft structure 11 comprises a perforated panel 13 having an inner surface 15 directed to a structure interior 17, an outer surface 19 in contact with an ambient flow 21, and a plurality of micro pores 23 distributed over the perforated panel 13 and connecting the inner and outer surfaces 15, 19. The perforated panel 13 is formed of titanium material. A plurality of elongate crack stopper elements 25 are attached to the inner surface 15 of the perforated panel 13 in parallel to one another. The crack stopper elements 25 are configured to inhibit crack propagation within the perforated panel 11 past the crack stopper elements 25.
    [0021] The crack stopper elements 25 extend, longitudinally, in a main load direction corresponding to the span direction of the aircraft structure 11. Further, the crack stopper elements 25 are formed as strips of fiber reinforced plastic (FRP) material having considerably higher fatigue strength than the related adjacent parts of the perforated panel 13, thereby forming a local fiber metal laminate (FML) together with the perforated panel 13 in the area of the strips. The width w of the crack stopper elements 25 of the embodiment shown in Fig. 2 is about 1/25 of the distance d between each two adjacent crack stopper elements 25.
    [0022] The embodiment shown in Fig. 3 differs from the embodiment of Fig. 2 by the width w of the crack stopper elements 25 being about 1/10 of the distance d between each two adjacent crack stopper elements 25. Further, due to the high width w of the crack stopper elements 25 the micro pores 23 in the perforated panel 13 extend also through the crack stopper elements 25, so that the micro pores 23 are not blocked by the crack stopper elements 25.
    [0023] In both Fig. 2 and Fig. 3 it is indicated how the crack stopper elements 25 inhibit crack propagation by a) stopping the crack 27, b) decelerating propagation of the crack 27, and c) deflect propagation of the crack 27 to an uncritical direction in parallel to the crack stopper elements 25.
    [0024] In Figs. 4 to 9 various embodiments of the aircraft structure 11 are shown where the aircraft structure 11 further comprises an inner panel 29 mounted to the perforated panel 13 via stiffeners 31 that are attached to the inner surface 15 of the perforated panel 13. The inner panel 29 and the stiffeners 31 are made of FRP material. In the embodiments of Figs. 4, 6, 7 and 9, the stiffeners 31 have Ω-profiles integral with the inner panel 29 and attached to the perforated panel 13. In the embodiments of Figs. 5 and 8 the stiffeners 31 have Z -profiles attached to both the inner panel 29 and the perforated panel 13. At least some of the crack stopper elements 25 are provided between the inner surface 15 of the perforated panel 13 and at least some of the stiffeners 31.
    [0025] In the embodiments shown in Figs. 4 to 6, crack stopper elements 25 in the form of FRP strips are provided between each stiffener 31 and the perforated panel 13. In the embodiment shown in Fig. 6, crack stopper elements 25 are additionally attached to the inner surface 15 of the perforated panel 13 at locations centrally between each two adjacent stiffeners 31. In the embodiment shown in Fig. 7, the crack stopper elements 25 are also in the form of FRP strips but are provided only at every other stiffener 31.
    [0026] In the embodiments shown in Figs. 8 and 9, instead of being in the form of FRP strips, the crack stopper elements 25 are formed by the stiffeners 31 themselves. Specifically, at least some of the stiffeners 31 are formed as crack stopper elements 25. In the embodiment of Fig. 8, the stiffeners 31 are formed as crack stopper elements 25 by the material of the stiffeners 31 being configured for crack stopping, i.e. the material of the stiffeners 31 is selected such that the stiffeners 31 have higher fatigue strength than the perforated panel 13.
    [0027] In the embodiment of Fig. 9, the stiffeners 31 are formed as crack stopper elements 25 by the shape of the stiffeners 25 being configured for crack stopping, i.e. the shape of the stiffeners 31 is selected such that the stiffeners 31 have higher fatigue strength than the perforated panel 13. This is done by providing the stiffeners 31 with an increased thickness at the head portion 33 attached to the inner surface 15 of the perforated panel 13.
    权利要求:
    Claims (13)
    [0001] An aircraft structure (11) for flow control, comprising
    a perforated panel (13) having an inner surface (15) directed to a structure interior (17), an outer surface (19) in contact with an ambient flow (21), and a plurality of micro pores (23) connecting the inner and outer surfaces (15, 19),
    characterized in that
    one or more elongate crack stopper elements (25) are attached to the inner surface (15) of the perforated panel (13), and
    the crack stopper elements (25) are configured to inhibit crack propagation within the perforated panel (13).
    [0002] The aircraft structure according to claim 1, wherein the crack stopper elements (25) extend in a main load direction of the aircraft structure (11).
    [0003] The aircraft structure according to claim 1 or 2, wherein the crack stopper elements (25) have a higher fatigue strength than the perforated panel (13).
    [0004] The aircraft structure according to any of claims 1 to 3, wherein the crack stopper elements (25) are formed as strips of fiber reinforced plastic (FRP) material.
    [0005] The aircraft structure according to any of claims 1 to 4, wherein the crack stopper elements (25) have a width (w) ranging from 1/100 to 1/1, preferably 1/25 to 1/15, most preferred 1/20, of the distance (d) between each two adjacent crack stopper elements (25).
    [0006] The aircraft structure according to any of claims 1 to 5, wherein the micro pores (23) in the perforated panel (13) extend also through the crack stopper elements (25).
    [0007] The aircraft structure according to any of claims 1 to 6, further comprising an inner panel (29) mounted to the perforated panel (13) via stiffeners (31) that are attached to the inner surface (15) of the perforated panel (13).
    [0008] The aircraft structure according to claim 7, wherein at least some of the crack stopper elements (25) are provided between the inner surface (15) of the perforated panel (13) and at least some of the stiffeners (31).
    [0009] The aircraft structure according to claim 7 or 8, wherein at least some of the stiffeners (31) are formed as crack stopper elements (25).
    [0010] The aircraft structure according to claim 9, wherein the stiffeners (31) are formed as crack stopper elements (25) by the material of the stiffeners (31) being configured for crack stopping.
    [0011] The aircraft structure according to claim 9 or 10, wherein the stiffeners (31) are formed as crack stopper elements (25) by the shape of the stiffeners (31) being configured for crack stopping.
    [0012] The aircraft structure according to claim 11, wherein the stiffeners (31) have an increased thickness at least at the head portion (33) attached to the inner surface (15) of the perforated panel (13).
    [0013] An aircraft (1) comprising a fuselage (3), wings (5), a vertical tail plane (7), and a horizontal tail plane (9), wherein the aircraft structure (11) according to any of claims 1 to 12 is arranged at the wings (5) and/or at the vertical tail plane (7) and/or at the horizontal tail plane (9).
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    同族专利:
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    引用文献:
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