DE102009057018A1 - Aircraft fuselage structure has skin field, where diagonal stiffening elements are formed by skin-field polygons, and stiffening frames are designed for limiting fuselage section - Google Patents

Abstract

The aircraft fuselage structure (1a) has a skin field (2), where diagonal stiffening elements (6) are formed by skin-field polygons (4). Stiffening frames (8) are designed for limiting a fuselage section (10), where the stiffening elements are connected with one of the stiffening frames. The stiffening element is connected at a corner area (12) of the stiffening frame.

Description

The invention relates to an aircraft fuselage structure according to the preamble of patent claim 1.

Currently, it is customary to manufacture aircraft fuselage structures from fuselage shells to which individual substantially rectangular skin panels forming stiffening structures are assigned. In many cases, the fuselage shells are carried out by stiffening profiles, which are connected directly to the inner surface of the shell and extend in the longitudinal direction, so-called stringers, and stiffening profiles which cross over them in the circumferential direction, so-called frames. The respective stiffening profiles are usually joined in a differential assembly process, preferably with riveted joints.

In addition to this conventional construction is for example from the US 4,086,378 B1 an aircraft fuselage structure with an integral stiffening grid is known, which has in a plane extending in shape and shape similar stiffening profiles that limit skin panels in the shape of isosceles triangles. This lattice structure is called isogrid. The individual skin fields are preferably assigned stiffening elements in the circumferential direction instead of ribs and additional diagonal stiffening elements instead of stringers. In the spaces between the diagonal stiffening elements of the fuselage structure window frames can be arranged. A disadvantage of such an aircraft fuselage structure is that the load-bearing stiffening elements limit the choice of the size and shape of fuselage cutouts such as window, door and maintenance cutouts. Furthermore, the load paths are disturbed when interrupted by the fuselage sections stiffening elements and require local reinforcements or thickening in the outer skin and / or the stiffening elements.

Isogrid structures are commonly used in cylindrical fuselage structures of constant radius, and in particular in non-internal pressure applications such as space rocket stages. In the case of an application in an aircraft fuselage structure, the lattice structure of an isogrid which is constant in shape and width proves to be not optimal in terms of weight, since substantially different main load directions and load amplitudes prevail in an aircraft fuselage over the circumference and in the longitudinal direction. A constant lattice structure in shape and width can not lead to a weight-optimized solution.

In contrast, the invention has for its object to provide an aircraft fuselage structure, in which a high structural rigidity with low manufacturing complexity and less limited size and shape of the stiffening structure of the fuselage shells and fuselage cutouts while reducing the structural weight is made possible.

This object is achieved by an aircraft fuselage structure having the features of patent claim 1.

The aircraft fuselage structure according to the invention is provided with skin panels, the stiffening elements in the circumferential direction, obliquely to the longitudinal axis employed diagonal stiffening elements, so-called stiffeners, and stiffening frames, in particular window and / or door frames are assigned, the fuselage cutouts for windows, doors, gates and / or maintenance openings limit, in particular fuselage cutouts are arranged to form windows in a longitudinally extending so-called window band. According to the invention, the stiffening elements are connected to at least one of the stiffening frames. This allows a power flow over the window or door frame, so that overall a load-optimized structure is achieved, which has a lower structural weight compared to conventional structures with the same or higher rigidity. In this way it is further achieved that the selection of the cutout size and shape can be made more freely, since the stiffening frame of the fuselage cutouts continue the load path of the stiffening elements respectively. Accordingly, an otherwise usual reinforcement of the outer skin and / or the stiffening elements in the circumferential direction in the region of the cutout or in particular in the region of a window strip can be reduced, so that a further reduction in weight is made possible.

According to a particularly preferred embodiment of the invention, at least one of the stiffening elements with at least one corner region of the stiffening frame z. B. a window and / or door frame connected. The stiffening elements preferably follow this radially to the stiffening frame.

The stiffening elements are preferably arranged in accordance with the prevailing main load direction in the respective fuselage area and dimensioned so that overall a lightweight structure with high rigidity and a low sensitivity to dents is achieved. The aforementioned isogrid structure represents a special case for skin fields in the form of isosceles triangles.

In a specific embodiment, at least two mutually diagonally opposite corner regions of the stiffening frame are each assigned at least one stiffening element connected thereto. At a According variant of the invention is associated with each corner region of the stiffening frame at least one stiffening element connected thereto. As a result, an advantageous introduction of force into the stiffening frame of the cutout and thus the most uniform possible load flow is achieved. In this case, for example, can be dispensed with diagonal stiffening elements between window openings in the window hinge.

The angles of the diagonal stiffeners to the longitudinal axis and the width of the triangular skin fields formed by the stiffeners preferably determine the cutout size and shape as a function of the structural mechanical requirements. The stiffening frame can be formed according to the invention substantially rectangular, since the frames are mitbefill. As a result, an enlarged compared to conventional solutions with oval or round cabin windows window surface is made possible.

Between at least two stiffening frames connected to at least one stiffening element, preferably a window frame, at least one stiffening frame, preferably a window frame, can be arranged, to which neither a diagonal nor a stiffening element in the circumferential direction is assigned. The intermediate frames executed at least in the corner regions without stiffening elements are in this case arranged within a skin panel.

The distances and angles of the stiffening elements to each other are preferably formed load-dependent. In this case proves to be advantageous in a circumferential direction gradually changing the spacing of the stiffening elements. Particularly advantageous here is a change in the spacing at positions with significant changes in the main load direction and load amplitude. In one embodiment of the aircraft fuselage structure, the spacings of the stiffening elements relative to one another below a floor, for example of the cabin floor, are reduced in relation to the distances above the floor. For this purpose, the mesh width is preferably provided by halving additional reinforcing elements.

In a further advantageous embodiment, the aforementioned change in the distances is performed by additional stiffening elements immediately below the window band. In the region of the stiffening frames, in particular of the window and / or door frames, the skin panel can have at least one planar reinforcing panel, in particular a reinforcing skin panel.

According to a further advantageous embodiment, the employment of the stiffening elements is adapted to the longitudinal axis along the circumference according to the changes of the main load direction. This leads to lower angles to the longitudinal axis of the aircraft in the upper and lower regions of the aircraft fuselage, in particular in the upper and lower shell, which are primarily loaded with tension and pressure in the longitudinal direction. In the lateral areas of the fuselage, the side shells, which are essentially loaded with thrust, the angles of the stiffening elements are set larger, so that they are optimized for a thrust load. Advantageously, this adjustment of the angle is carried out together with the aforementioned stepwise change in the spacing of the stiffening elements.

In an aircraft fuselage structure with at least one upper fuselage shell and at least one lower fuselage shell, it has proved to be advantageous to arrange the shell joints such that they do not extend through the region of the window and / or door frames. To be particularly advantageous, it has been shown to arrange the window and / or door frame below upper shell joints. As a result, enforcement in the hindleg structure and a technically sophisticated connection of the rear structure with the shells in the frame areas avoided or significantly reduced by the smaller number of stiffening elements in the upper fuselage shell.

According to a preferred embodiment of the invention, skin panels with the stiffening elements form an integral or integrally fabricated hull shell. This achieves a reduction of the process steps in the hull assembly and a further weight optimization of the fuselage structure.

The stiffening frame delimiting the fuselage cutout preferably forms a circumferential profile, which is a further component of the lightweight structure and, compared to previous designs, improves the structural mechanical performance since it is possible to dispense with joints in this highly stressed area. As a result, an additional potential for reducing assembly costs and weight can be achieved.

The hull shells are preferably designed as a fiber composite component, in particular as a CFRP component, highly integrated in order to optimally utilize material-specific advantages. In particular, the hull shells may be made by a resin infusion process.

In one exemplary embodiment of the aircraft fuselage structure, in each case a fuselage bulkhead, preferably an omega fuselage bulkhead, is arranged between two adjacent window frames as stiffening element in the circumferential direction. Such omega fuselage frames are advantageous to produce and reduce the number of components required for Connection to the fuselage shell and for lateral stabilization, so-called clips and cleats. The load is distributed over a larger area due to the two leg feet of the Omega trunk frame. The load can be transferred by connecting floor cross members and support profiles via connecting elements with the omega frames. Local skin enhancements and Doppler are also possible.

The connecting elements can be integrally formed or joined by gluing and / or riveting to the floor cross members, support profiles and fuselage frames. For already mounted on the skin field omega frames, the connection can be done by means of so-called blind rivets.

Other advantageous developments of the invention are part of the further subclaims.

In the following preferred embodiments of the invention will be explained in more detail with reference to schematic drawings. Previously mentioned stiffening frames of fuselage cutouts and their connection to the stiffening elements are shown below by means of window cutouts. Show it:

1 a spatial representation of an aircraft fuselage structure according to a first embodiment of the invention;

2 an enlarged view of the detail Y of the section of the aircraft fuselage structure 1 in the area of a window frame;

3 an enlarged view of the detail X from 1 ;

4 a spatial representation of another aircraft fuselage structure according to a second embodiment of the invention;

5 an enlarged view 4 in the area of a window band;

6 a schematic cross-sectional view of the fuselage structure 1 ;

7 a schematic representation of an aircraft fuselage structure according to a third embodiment of the invention;

8th a schematic representation of an aircraft fuselage structure according to another embodiment of the invention;

9 a schematic representation of an aircraft fuselage structure according to another embodiment of the invention;

10 a schematic representation of an aircraft fuselage structure according to another embodiment of the invention;

11 a schematic representation of an aircraft fuselage structure according to another embodiment of the invention and

12 a schematic representation of an aircraft fuselage structure according to another embodiment of the invention.

1 shows an aircraft fuselage structure 1a formed from at least one hull shell 2 , the dermal polygon 4 forming diagonal stiffening elements 6 so-called stiffeners and stiffening frames 8th , in particular window frames, are assigned to the fuselage cutouts 10 , in particular window cutouts, limit. The stiffening elements 6 are with the window frames 8th in corner areas 12 connected. Between in corner areas 12 with stiffening elements 6 connected window frames 8th is each a window frame 14 arranged, the no stiffening element 6 assigned. These in the corner areas 12 without stiffening elements 6 running in-between window frames 14 are here within a skin field polygon 4 ie within a mesh.

In particular 2 it can be seen that an enlarged section of the fuselage structure 1a out 1 in the area of a window frame 8th shows, is in each case a stiffening element 6 with a corner area 12 of the window frame 8th connected. The stiffening elements designed as T-profiles 6 close radially to the window frame 8th at. The window frame 8th forms a circumferential profile, which is another component of the lightweight structure and compared to previous designs improves the structural mechanical performance, as can be dispensed with joints in this highly loaded area. This results in an overall optimum power flow with reduced structural weight of the fuselage structure 1a reached. The power flow takes place in sections over the window frame 8th , so that overall a load-optimized structure is achieved, which has a lower structural weight compared to conventional structures with the same or higher rigidity. This also ensures that the choice of window size and window shape can be made more freely because the window frame 8th the load path of the stiffening elements 6 continue each. The stiffening elements 6 are arranged and dimensioned depending on the structural load. Accordingly, a reinforcement of the outer skin in the area a ribbon window 42 be reduced, so that a further weight reduction is possible.

The window frames 8th are formed substantially rectangular. As a result, an enlarged compared to conventional solutions with oval or round cabin windows window surface is made possible, which improves the passenger comfort.

The skin fields 4 form with the stiffening elements 6 integral hull shells 2 respectively. 16 out. The hull shells 16 are designed as a CFRP fiber composite component in a highly integral manner in order to optimally exploit the material-specific advantages. In particular, the hull shells 16 be prepared by a resin infusion process. This results in a simplified production and a further weight optimization of the hull structure 1 reached.

The angles and width of the Hautfeld polygons 4 determine the window size depending on the structural mechanical requirements. The distances between the stiffening elements 6 to each other are formed load dependent. At the in 1 illustrated embodiment of the aircraft fuselage structure 1a are the spacings of the stiffening elements 6 to each other below a cabin floor 18 opposite the distances above the cabin floor 18 reduced. For this purpose, the mesh size is halved additional diagonal stiffening elements 20 intended. Between two adjacent window frames 8th is each as an omega fuselage frame 22 trained fuselage bulkhead arranged as a stiffening element in the circumferential direction. The load is due to the two leg feet of the omega frames 22 advantageously distributed over a larger area. The radial load transfer takes place by connecting floor cross members 24 and support profiles 26 over fasteners 28 . 30 (see. 3 ) with the omega frames 22 ,

As 3 It can be seen, the enlarged view of the detail X from 1 shows are the floor cross member 24 about trained as tabs fasteners 28 with the fuselage frames 22 riveted. Advantageously, the floor cross member 24 on both sides of the Omega hull frame 22 either with two connecting elements designed as tabs 28 connected, or with an integral connection element the same function. In the same way, further support elements of the floor, in particular support profiles 26 be connected to the hull bulkhead.

4 shows a further embodiment of an aircraft fuselage structure 1b with T-profiled stiffening elements 6 both in the circumferential direction, as well as in a diagonal, angled to the aircraft longitudinal axis direction. The floor structure and its connection is not shown in this figure. Instead of the omega fuselage frames 22 in 1 is central to each window 10 and between each pane 10 a circumferential, circumferentially extending stiffening element 6 arranged. Both the centrally arranged to the cutouts frames, and the diagonally arranged stiffening close to the end of the window frame 8th on, whereby a continuation of the load paths is ensured. In contrast to the embodiment in 1 be through the stiffening elements 6 exclusively triangular skin fields formed with substantially the same shape and size. Nevertheless, the diagonal stiffening element 6 along the hull circumference in their employment to the longitudinal axis continuously adjusted according to the prevailing main load direction.

According to 5 , which is an enlarged view of the window band 42 out 4 it can be seen close to the center of the window openings arranged, designed as a hull ribbing stiffening elements 6 end to the oval window frame 8th in the middle, whereby a continuation of the load paths is ensured. As already in 2 The diagonal stiffening elements also close 6 end in the corner areas 12 of the window frame 8th at right angles.

Advantageously, the embodiment is made 1 with the embodiment 2 combined in such a way that at a suitable point a reduction of the stiffening elements 6 and the fuselage frames in 1 he follows. In particular, this reduction may be directly above the window band 42 be achieved. As a result, the centrally arranged hull bulkhead closes only at the bottom of the window opening 10 on the window frame 8th at the end. Accordingly, the diagonal stiffening elements 6 arranged such that they end only at one of the corner areas 12 the upper half of a window cutout 10 connect.

As 6 it can be seen, which is a schematic cross-sectional view of the fuselage structure 1a respectively. 1b out 1 respectively. 4 shows it is at a hull structure 1a or 1b with several hull shells 16a -D advantageous when the shell joints 36a Are arranged so that they do not run through the area of the window frame. In the illustrated embodiment find an upper hull shell 16a and two middle hull shells 16b . 16d and a lower hull shell 16c Use. It has proved to be advantageous in this case, the window hinge 42 below upper shell joints 36a . 36d to arrange.

7 shows a schematic representation of an aircraft fuselage structure 1c according to a second embodiment of the invention, that essentially differs from the variant described above in that each window frame 8th in the four corners 12 each a stiffening element 6 assigned. In other words - the window frames 8th are so relative to the skin field polygons 4 positioned that every corner areas 12 with stiffening elements 6 are provided and alternately a fuselage frame in the middle of the window frame 8th tethered and between two window frames 8th is arranged.

8th shows an aircraft fuselage structure 1d According to another embodiment of the invention, in accordance with the embodiment of 7 every window frame 8th in the four corners 12 each a stiffening element 6 is assigned, wherein the centrally arranged hull frame over the previous example accounts.

9 shows an aircraft fuselage structure 1e according to another embodiment of the invention, in which every second fenestration 8th at all corners 12 stiffeners 6 are assigned (see. 1 ).

10 shows an aircraft fuselage structure 1f According to a further embodiment of the invention that essentially differs from the variants described above in that only two of a side of the window frame 8th associated corner areas 12 each one connected to these stiffening element 6 assigned. Here are the stiffening elements 6 each on adjacent sides of the window frame 8th arranged.

11 shows an aircraft fuselage structure 1g According to a further embodiment of the invention that essentially differs from the variants described above in that two diagonally opposite corner regions 12 the window frame 8th each one connected to these stiffening element 6 assigned. The stiffening elements 6 are each on adjacent sides of the window frame 8th arranged.

12 shows an aircraft fuselage structure 1h according to a last embodiment of the invention, in which the stiffening elements 6 are not interrupted by window frames but a plurality of window frame forming triangular structures 38 limit. The stiffening profiles 6 are also in this variant as with the skin 2 formed integral or integral T-profiles, wherein triangular window elements 40 inserted in the T-profiles and attached to these. This will make a window band 42 formed with a large axial extent in a highly rigid structure.

The aircraft fuselage structure according to the invention 1a -H is not on the described variants with window frames 8th limited, but the structure according to the invention can be provided in alternative, not shown embodiments with door frame of a car door, a cargo door, an emergency exit hatch or a manhole etc.

Disclosed is an aircraft fuselage structure 1a -H with at least one skin area 2 , the dermal polygon 4 forming diagonal stiffening elements 6 and stiffening frame 8th for hull cuts 10 are assigned, wherein the stiffening elements 6 with at least one of the stiffening frames 8th connected to form a highly rigid structure.

LIST OF REFERENCE NUMBERS

1a-h

Aircraft fuselage structure

2

fuselage shell

4

Skin panel polygon

6

stiffener

8th

stiffening frame

10

hull Extract

12

corner

14

stiffening frame

16a-d

fuselage shell

18

cabin floor

20

stiffener

22

bulkhead

24

Floor beams

26

support profile

28

connecting element

30

connecting element

32

crossing area

34

profile reinforcement

36a-d

shell shock

38

triangular structure

40

element window

42

window strip

QUOTES INCLUDE IN THE DESCRIPTION

This list of the documents listed by the applicant has been generated automatically and is included solely for the better information of the reader. The list is not part of the German patent or utility model application. The DPMA assumes no liability for any errors or omissions.

Cited patent literature

• US 4086378 B1 [0003]

Claims (13)

Aircraft fuselage structure with at least one skin panel ( 2 ), the skin field polygon ( 4 ) forming diagonal stiffening elements ( 6 ) and a fuselage cutout ( 10 ) limiting stiffening frameworks ( 8th ), characterized in that the stiffening elements ( 6 ) with at least one of the stiffening frames ( 8th ) are connected. An aircraft fuselage structure according to claim 1, wherein at least one of the stiffening elements ( 6 ) with at least one corner area ( 12 ) of the stiffening frame ( 8th ) connected is. An aircraft fuselage structure according to claim 2, wherein each corner region ( 12 ) of the stiffening frame ( 8th ) at least one stiffening element connected thereto ( 6 ) assigned. Aircraft fuselage structure according to one of the preceding claims, wherein the stiffening elements ( 6 ) radiating on the stiffening frame ( 8th ) connect. Aircraft fuselage structure according to one of the preceding claims, wherein the stiffening frames ( 8th ) are substantially rectangular. An aircraft fuselage structure according to any one of the preceding claims, wherein the fuselage cutouts ( 10 ) and the limiting stiffening frameworks ( 8th ) a window band ( 42 ) train. An aircraft fuselage structure according to any one of the preceding claims, wherein the spacings of the stiffening elements ( 6 ) to each other below a floor ( 18 ) with respect to the distances above the floor ( 18 ) and / or below the window band ( 42 ) with respect to the distances above the window band ( 42 ) are reduced. Aircraft fuselage structure according to one of the preceding claims, wherein the angles of the stiffening elements ( 6 ) are variable to the longitudinal axis in the circumferential direction of the fuselage. Aircraft fuselage structure according to one of the preceding claims, with at least one upper fuselage shell ( 16a ) and at least one lower hull shell ( 16c ) whereby the shell joints ( 36a -D) are arranged such that they are not penetrated by the region of the stiffening frame ( 8th ). Aircraft fuselage structure according to one of the preceding claims, wherein the skin panel ( 2 ) with the stiffening elements ( 6 ) integral or integrally manufactured hull shells ( 16a -D) is formed. An aircraft fuselage structure according to claim 10, wherein at least one fuselage shell ( 16a -D) is a fiber composite component, in particular a CFRP component. Aircraft fuselage structure according to one of the preceding claims, comprising a plurality of stiffening frames ( 8th ), in particular window frames, wherein between at least two with at least one stiffening element ( 6 ) stiffening frames ( 8th ) at least one window frame ( 14 ) is arranged, which no stiffening element ( 6 ) assigned. Aircraft fuselage structure according to one of the preceding claims, wherein between two adjacent stiffening frames ( 8th ), in particular window frames of a window hinge ( 42 ), in each case a fuselage frame ( 22 ) and at least one bulkhead ( 22 ) centrally to at least one of these stiffening frames ( 8th ) is arranged.

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