AIRCRAFT STRUCTURE FOR FLOW CONTROL

Abstract

An aircraft structure (11) for flow control including 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). Weight reduction and maintaining required fatigue strength of the structure may be achieved because 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).

Claims

1. An aircraft structure for flow control comprising a perforated panel having an inner surface directed to a structure interior, an outer surface configured to be contact with an ambient flow, and a plurality of micro pores connecting the inner and outer surfaces, and at least one elongate crack stopper element attached to the inner surface of the perforated panel, wherein the at least one elongate crack stopper element is configured to inhibit crack propagation within the perforated panel.

2. The aircraft structure according to claim 1, wherein the at least one elongate crack stopper element extends in a main load direction of the aircraft structure.

3. The aircraft structure according to claim 1, wherein a material forming the at least one elongate crack stopper element has a fatigue strength higher than a material forming the perforated panel.

4. The aircraft structure according to claim 1, wherein the at least one elongate crack stopper element includes at least one strip of fiber reinforced plastic (FRP) material.

5. The aircraft structure according to claim 1, wherein the at least one elongate crack stopper element includes two adjacent crack stopper elements and each of the two adjacent crack stopper elements has a width (w) in a range of 1/100 to 1/1 of a distance (d) between the two adjacent crack stopper elements.

6. The aircraft structure according to claim 1, wherein the micro pores in the perforated panel extend through the at least one elongate crack stopper element.

7. The aircraft structure according to claim 1, further comprising an inner panel mounted to the perforated panel via stiffeners that are attached to the inner surface of the perforated panel.

8. The aircraft structure according to claim 7, wherein the at least one elongate crack stopper element includes a plurality of crack stopper elements between the inner surface of the perforated panel and at least some of the stiffeners.

9. The aircraft structure according to claim 7, wherein at least some of the stiffeners are configured as crack stopper elements of the at least one elongate crack stopper element.

10. The aircraft structure according to claim 9, wherein the stiffeners configured as crack stopper elements are formed of a material having a fatigue strength higher than a material forming the perforated panel.

11. The aircraft structure according to claim 9, wherein the stiffeners are formed as crack stopper elements are shaped are shaped for crack stopping.

12. The aircraft structure according to claim 11, wherein the stiffeners have an increased thickness at least at a head portion which is attached to the inner surface of the perforated panel.

13. An aircraft comprising a fuselage, wings, a vertical tail plane and a horizontal tail plane, wherein the aircraft structure according to claim 1, is arranged at the wings and/or at the vertical tail plane and/or at the horizontal tail plane.

14. An aerodynamic structure on an aircraft comprising: a perforated skin panel included an outer surface configured to be in contact with an ambient airflow and inner surface opposite to the outer surface and facing an interior of the aerodynamic structure, and micro pores extending through the perforated skin panel and connecting the inner and outer surfaces, and crack stopper elements bonded to the inner surface of the perforated skin panel such that the crack stopper elements overlap at least some of the micro pores, wherein a plurality of the crack stopper elements are oriented in a spanwise direction of the aerodynamic structure and the crack stopper elements each have a width narrower than a gap between adjacent ones of the crack stopper elements.

15. The aerodynamic structure of claim 14, wherein the perforated skin panel is metallic or a fiber metal laminate, and the crack stopper elements is a fiber reinforced plastic material.

16. The aerodynamic structure of claim 14, wherein the crack stopper elements include micro pores aligned with the micro pores of the perforated skin panel.

17. The aerodynamic structure of claim 14, further comprising stiffeners extending in the spanwise direction and the crack stopper elements are sandwiched between the stiffeners and the perforated skin panel.

18. The aerodynamic structure of claim 14, further comprising stiffeners extending in the spanwise direction, and the crack stopper elements are integral with the stiffeners.

Description

SUMMARY OF DRAWINGS

[0020] Embodiments of the present invention are illustrated in the drawings which are:

[0021] FIG. 1 is a perspective view of an aircraft according to the invention,

[0022] FIGS. 2A and 2B are schematic illustrations (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of low width FRP strips,

[0023] FIGS. 3A and 3B are schematic illustrations (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements in the form of high width FRP strips,

[0024] FIGS. 4A and 4B are schematic illustrations (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged between -stiffeners and the perforated panel,

[0025] FIGS. 5A and 5B are schematic illustrations (top and cross sectional view) of an embodiment of an aircraft structure having crack stopper elements arranged between Z-stiffeners and the perforated panel,

[0026] FIGS. 6A and 6B are schematic illustrations (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,

[0027] FIGS. 7A and 7B are 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,

[0028] FIGS. 8A and 8B are schematic illustrations (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,

[0029] FIGS. 9A and 9B are schematic illustrations (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.

DETAILED DESCRIPTION

[0030] In FIG. 1 shows an aircraft 1 comprising 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.

[0031] In FIGS. 2A and 2B 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.

[0032] 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 FIGS. 2A and 2B is about 1/25 of the distance d between each two adjacent crack stopper elements 25.

[0033] The embodiment shown in FIGS. 3A and 3B differs from the embodiment of FIGS. 2A and 2B 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.

[0034] In FIGS. 2A and 2B, and FIGS. 3A and 3B, 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.

[0035] In FIGS. 4A and 4B to 9A and 9B 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. 4A and 4B, 6A and 6B, 7A and 7B, and 9A and 9B, the stiffeners 31 have -profiles integral with the inner panel 29 and attached to the perforated panel 13. In the embodiments of FIGS. 5A and 5B and 8A and 8B 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.

[0036] In the embodiments shown in FIGS. 4A and 4B, to 6A and 6B, 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 FIGS. 6A and 6B, 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 FIGS. 7A and 7B, the crack stopper elements 25 are also in the form of FRP strips but are provided only at every other stiffener 31.

[0037] In the embodiments shown in FIGS. 8A and 8B and FIGS. 9A and 9B, 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 FIGS. 8A and 8B, 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.

[0038] In the embodiment of FIGS. 9A and 9B, 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.

[0039] While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms comprise or comprising do not exclude other elements or steps, the terms a or one do not exclude a plural number, and the term or means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.