WO2011066950A1

WO2011066950A1 - Aircraft fuselage component

Info

Publication number: WO2011066950A1
Application number: PCT/EP2010/007269
Authority: WO
Inventors: Andreas Knote; Steffen Niemann; Henrik Borgwardt; Robert Kaps; Tobias Ströhlein; Christian Hühne
Original assignee: Deutsches Zentrum für Luft- und Raumfahrt e.V.
Priority date: 2009-12-03
Filing date: 2010-12-01
Publication date: 2011-06-09
Also published as: DE102009056533B4; DE102009056533A1

Abstract

The invention relates to an aircraft fuselage component (4) having an exterior wall (5) with reinforcing elements and having an opening (3, 9) in the exterior wall (5). The aircraft fuselage component (4) is designed as an integral fibre composite element, in which a reinforcing fibre lattice (7) for forming the reinforcing elements is arranged adjacent to the opening (3, 9) and the exterior wall (5), wherein the reinforcing fibre lattice (7) is integrally connected to the exterior wall (5) and wherein the fibre orientation of the reinforcing fibre lattice (7) is adapted to redirect loads around the opening (3, 9) and to divert loads acting at load application points at the opening (3, 9) into the lateral surface regions of the aircraft fuselage component (4).

Description

PLANE HULL COMPONENT

The invention relates to an aircraft fuselage component with an outer wall with reinforcing elements and with an opening in the outer wall.

Aircraft door environment structures are conventionally constructed by making a framework of frames with main and secondary frames and side rails to which the skin is attached. The structure of the fuselage is defined by large cutouts, such as e.g. noticeably weakened for the door opening. Static and dynamic loads acting on the door environment structure must therefore be redirected around the large area using a specially designed and enhanced environment structure. Internal pressurization of an aircraft fuselage during flight, caused by the differential pressure between the aircraft interior side and the aircraft exterior side, is initiated at discrete points from the door into the aircraft fuselage. The forces occurring at these discrete points provide for localized load overshoot, which must be suitably transmitted through the door environment structure into the fuselage. The door environment structure thus supports the large door cutout and introduces the loads from the door into the fuselage.

Conventionally, the door environment structure is made in differential construction. The functions of the individual components, in particular the main and secondary frames, the door side rails, the intercostal and the fuselage skin are clearly separated. The differential design requires a high assembly cost to assemble the items. This leads to a large amount of time and high costs. Furthermore, created by the joining of many items a long tolerance chain, so that the production of the individual parts and the

CONFIRMATION COPY Assembly must be done with high precision. Furthermore, in the area of the door environment structure, to the left and right of the door cut-out, the rib of the fuselage is continued in the radial direction, and intercostals are used at the level of the door's load introduction points. This gives the door environment structure its typical ladder shape.

A disadvantage of the differential design is also the high cost of repair and replacement of damaged items in case of failure. DE 10 2004 009 020 A1 describes intercosts for frames in aircraft, with which a load can be derived from a first frame to a second frame and / or to the outer skin of the aircraft. With the help of these separate intercostal a targeted power transmission, especially in the field of door openings, possible.

DE 10 2007 021 076 AI also discloses intercostals or intermediate ribs for an aircraft for discharging a load from a bulkhead to another bulkhead or skin. With the aid of a movable cross-member, lead-through areas for system lines are easily accessible without affecting the truss structure.

Even with the use of fiber composite material is traditionally a replica of the metallic differential construction with ribs, side members and Z wi schenkostale.

The disadvantage of conventional aircraft fuselage components is the high assembly costs due to the number of individual parts and the necessary compliance with tolerances. In addition, damaged items are very difficult to replace. The optical detectability of damage is difficult or impossible with fiber composite materials. This increases the risk of discovering a damaged structure only after many flight cycles during extensive maintenance. Another problem is that repairs, especially on the fuselage skin in the door area are very expensive, since the outer skin undergoes very high shear stress at this large section. The object of the present invention is therefore to provide an improved aircraft fuselage component with outer wall, reinforcing elements and an opening in the outer wall.

The object is achieved with the aircraft fuselage component of the aforementioned type in that the aircraft fuselage component is formed as an integral fiber composite element, in which adjacent to the opening and the outer wall, a reinforcing fiber webbing for forming the reinforcing elements is arranged, wherein the reinforcing fiber web is integrally connected to the outer wall and wherein the fiber orientation of the reinforcing fiber scrim is adapted to divert loads around the aperture and divert loads applied to load introduction points at the aperture into the side surface regions of the aircraft fuselage.

It is proposed to form an aircraft fuselage component as an integral fiber composite element, in which the reinforcing elements are not designed, as usual, discretely from individual components each having an assigned function. Rather, the pronounced anisotropy of fiber composites is exploited, which can accommodate significantly higher tensile loads in the fiber longitudinal direction, as transverse to it. This makes it possible to form the aircraft fuselage as an integral fiber composite element with integrated reinforcing fiber, in which the fiber orientation is adapted to optimize the environmental structure of the opening. The fiber orientation of the reinforcing fiber fabric makes it possible to divert the cutting loads which are introduced from the aircraft fuselage and, for example, a large opening into the environmental structure of the opening, around the large cutout, ie the opening. Explicitly shaped frames and intercostals in the vicinity of the opening according to the prior art are no longer necessary because they are replaced by a load-optimized alignment of the reinforcing fiber fabric. The aircraft fuselage component therefore also leads to a weight reduction. Since the fiber alignment of the reinforcing fiber webs takes into account the cutting and door loads, the aircraft body component has comparable mechanical properties to a conventional structure of differential construction.

The opening is preferably provided for receiving an aircraft door.

For the purposes of the present invention, aircraft are to be understood as meaning any aircraft, but in particular, commercial aircraft and helicopters, in which the aircraft fuselage component can be advantageously used.

It is particularly advantageous if the reinforcing fiber fabric is covered with a sacrificial layer for detecting damage, for example by impact stress.

By applying such a sacrificial layer on the outside of the aircraft fuselage, damage to the environmental structure of the aperture may be avoided, or at least e.g. be detected optically. The sacrificial layer may e.g. a polymer foam with a closed-pored surface. The polymer foam absorbs impact stress to a certain extent. In case of overuse, the pores collapse and the damage becomes visible on the surface by a duckling. A similar principle can also be achieved with a honeycomb core structure with a thin cover layer as the sacrificial layer on the outer skin.

In the event that the environmental structure of the opening is damaged, a complex repair can be avoided. Due to the fully integral fuselage component, only the fuselage component must be detached from the surrounding fuselage area and replaced with a new one.

It is particularly advantageous if the aircraft fuselage component has a separate supply element that is detachable from the outer wall and the reinforcing fiber web formed integrally with the outer wall. Not all To disassemble supply lines for electrical, air, water and hydraulics of the aircraft around the opening during an exchange of the fuselage component, these supply lines are mounted on this separate carrier element decoupled from the aircraft fuselage component. When the aircraft fuselage component is replaced, the carrier element only has to be detached from it and remains in the aircraft. For this purpose, it is releasably connected to only a few defined and easily dismantled bearings with the aircraft fuselage.

This separate detachable support element for supply lines thus has a significant advantage over conventional door environment structures in which the supply lines are nested housed in the door environment structure and removable only with great effort.

The main direction of extension of the fibers of the reinforcing fiber fabric is preferably adapted to the direction of the load acting on the aircraft fuselage, such that the fibers redistribute the flows of load distributed from the opening into the peripheral regions of the aircraft fuselage and the flows from the peripheral regions of the aircraft fuselage and loads acting on the outer wall the opening around in other edge regions of the fuselage component derives.

Thus, it is particularly advantageous if a first laminate of a reinforcing fiber fabric with a fiber orientation in the main direction of the fibers is provided such that reinforcing fibers extend from the side edges of the aircraft body component diagonally past the opening to the upper edges and / or lower edges of the aircraft fuselage. This first layer of reinforcing fiber fabric thus provides for the load redirection of loads on the side edges of the aircraft fuselage and on the outer skin around the opening. It is particularly advantageous if a second structure, which is produced integrally with the first laminate, of a reinforcing fiber scrim having a fiber orientation in the main extension direction of the fibers is provided such that reinforcing fibers extend from the region adjacent to the opening from the opening extend away in the direction of the side edges, wherein a plurality of reinforcing fiber bundles are arranged at a distance from each other.

This second structure of the reinforcing fiber fabric serves to dissipate it onto the opening, e.g. loads acting on a door in the opening into the aircraft fuselage component and for as even as possible a planar discharge of this load from the opening to the fuselage in the vicinity of the aircraft fuselage component.

By selectively aligning the fibers of the reinforcing fiber web along the load paths, a smooth transition of the load from point initiation, e.g. reached at the door to the introduction to the outer skin. In a corresponding manner, a surface load application, e.g. from the door in the opening into the surrounding structure possible. With the aid of the first laminate for diverting the load around the opening and through the second structure for discharging loads from the opening to the outside, a weight-saving design of the aircraft fuselage component which is optimally adapted to the load distribution succeeds in meeting the requirements for load transfer of load introduction points the opening is taken into account as well as the requirements for a uniform as possible load distribution in the fuselage component with redirection of loads around the opening.

The derivation of the load from a load introduction point at the opening to the side edges of the aircraft fuselage component preferably does not take place simply by means of a reinforcing fiber layer directed outwardly from the load introduction point. Rather, it is advantageous if the reinforcing fiber bundles extend with part of their length along an edge region of the opening. Thus, the reinforcing fiber bundles of the second layer of the reinforcing fiber scrim should preferably completely enclose the opening altogether so as not only to dissipate loads from certain load transfer points, but to pick up the loads around the circumference of the opening and dissipate outwardly therefrom over comb-like reinforcing fiber bundles. The opening is advantageously reinforced by the reinforcing fiber bundles extending along edge regions of the opening in this way. The invention will be described below with reference to the accompanying drawings with an embodiment. Show it:

Figure 1 - sketch of an aircraft with openings in the fuselage;

FIG. 2 shows a sketch of an aircraft fuselage component with a door opening and integrally formed reinforcing fiber layers;

3 shows a three-dimensional view of a section of an aircraft fuselage with an aircraft fuselage component inserted therein for a door opening.

FIG. 1 shows a sketch of an aircraft 1 in whose fuselage 2 openings 3 for doors and windows are provided.

These areas of the fuselage 2 of the aircraft 1 with the openings 3 require due to the increased loads acting on them and the weakening through the opening 3 a special structural design.

FIG. 2 shows a sketch of an aircraft fuselage component 4 according to the invention for use in the region of such an opening 9, in particular a door opening. It is clear that the aircraft fuselage component 4 is designed as an integral fiber composite element and has an outer wall 5 (outer skin). Adjacent to this outer wall 5, a first laminate 6 of a reinforcing fiber scrim 7 is integrally formed therewith. This first layer 6 of the reinforcing fiber fabric 7 essentially serves to dissipate loads from the side edges 8a, 8b including the upper edge 8c and lower edge 8d around the opening 9 into the aircraft fuselage as evenly as possible and to guide them out of them. The fiber orientation of the fibers of this reinforcing fiber scrim 7 of the laminate 6 is optimized such that its main extension direction is directed diagonally past the opening 9 from the side edges 8a, 8b to the upper edges 8c and lower edges 8d of the aircraft fuselage component 4, respectively. Diagonal reinforcing fiber bundles can furthermore be provided at the four corners, with the result that reinforcing fibers flow from one side edge 8a or 8b into the outer area of the adjacent upper edge 8c or lower edge 8d or Reinforcing fibers from the central region of the upper edge 8c and lower edge 8d in the outer region of the respective upper edge 8c and lower edge 8d are guided past the opening 9. A second structure 10 with reinforcing fiber webs 7 is provided to guide load application points at the opening 9 away from the opening 9 into the aircraft fuselage component 4 and to distribute it in an areally possible manner. The main direction of extension of the fibers of this second structure 10 of the reinforcing fiber fabric 7 is optimized such that reinforcing fibers extend from the region adjacent to the opening 9 from the opening 9 in the direction of the side edges 8a, 8b or possibly the upper edges 8c and lower edges 8d as side edges , In this case, a plurality of reinforcing fiber bundles are arranged at a distance from one another, so that the reinforcing fiber bundles extend in a finger-like manner away from the opening 9. These fingers of the reinforcing fiber bundles are spaced apart from each other and extend to each other so that a comb-like structure is formed, is derived by the load surface of the edge region of the opening 9 in the aircraft fuselage component 4. From there, the load can then be discharged relatively evenly over the side edges 8a, 8b to the body regions adjacent to the aircraft fuselage component 4.

The area of the illustrated door environment structure is thus considered as an arbitrary structure to be optimized. For this purpose, in a vicinity of the opening 9, such as the door opening, around the cutting loads are considered, which act from the rest of the fuselage structure on the illustrated environment structure. In addition, the loads from the door to the door environment structure are considered. In particular, shear loads are considered as cutting loads.

The reinforcing fiber scrims 7 now exploit the pronounced anisotropy of fiber composites, which can absorb significantly higher tensile loads in the fiber longitudinal direction than transversely thereto. The aircraft fuselage component 4 is thus designed as an optimized surface structure with ideal fiber orientation. Incidentally, the cutting loads introduced from the fuselage and the door into the door environment structure may be along the ideal fiber orientation the load paths are passed around the large opening 9. It can be seen that the aircraft fuselage component is not bound to the geometric boundary conditions of existing frames and intercostals, so that the weight of the aircraft fuselage component can be reduced compared to conventional door opening structures. The exact shape of the optimized structure depends on the type of aircraft and the structural design of a door or tailgate in the opening 9 and the hull design and must be optimally designed accordingly for the prevailing loads. It is important to exploit the optimized fiber orientation so that the tensile loads are optimally absorbed in the fiber longitudinal direction and passed around the opening 9.

The fully integral aircraft fuselage component 4 can be manufactured either in resin injection process in one shot. It is also conceivable, however, for the one or more layers of reinforcing fiber ply 7 to be glued to the outer skin or riveted to the outer skin. In this case, the reinforcing fiber scrims would be made as sub-modules which are then integrally bonded to the outer skin.

FIG. 3 shows a representation of a section of an aircraft fuselage 2 with the usual frames 1 1 and stringers 12. It is clear that the aircraft fuselage component 4 shown in FIG. 2 is inserted into a section of the fuselage 2. In this case, via the ribs 1 1 and stringer 12 and optionally via the outer wall 5 introduced loads on the reinforcing fiber scrim 7 around the opening 9 around. This load redirection is preferably such that the hull 2 is still sufficiently strong despite the opening 9 and the aircraft fuselage component 4 used therein. The weakening of the hull 2 caused by the opening 9 is compensated by the reinforcing fiber webs 7, which are formed integrally with the outer wall 5 and adapted to the opening 9 and the load flow.

FIG. 3 shows the view of a half-shell of the fuselage 2 from the inside on the inside of the fuselage component 4. Support elements 13 for supply lines (not shown) can be seen, which are configured as separate components for the fuselage component 4. These carrier elements 13 are connected to as few bearings as possible with the aircraft fuselage component 4 and carry Supply lines for eg electrical, air, water and hydraulics of the aircraft 1. By decoupling the carrier elements 13 of the aircraft fuselage component 4, the fuselage component 4 can be easily replaced after releasing the storage without the supply lines must be removed.

Furthermore, an optional sacrificial layer (not shown) may be provided on the outer side of the fuselage in the vicinity of the opening 9 or the door, which protects the aircraft fuselage component 4 from damage and damages e.g. optically visible. Impact damage on the outer wall 5 and the sacrificial layer arranged on the outer side of the outer wall 5 are visually visible through deformation of the sacrificial layer due to the plastic deformation of the sacrificial layer. Such a damage can then be easily detected and lead to the replacement of the aircraft fuselage component or only the sacrificial layer 4. The sacrificial layer may e.g. a polymer foam with a closed-pored surface. The polymer foam absorbs impact stress to a certain extent. In case of overuse, the pores collapse and the damage becomes visible on the surface by a duckling. In a corresponding manner, the sacrificial layer can also be achieved with the aid of a honeycomb core structure with a thin cover layer as the sacrificial layer on the outer skin.

Claims

claims

A fuselage fuselage component (4) with an outer wall (5), with reinforcing elements and with an opening (3, 9) in the outer wall (5), characterized in that the fuselage trunk component (4) is formed as an integral fiber composite element adjacent to the opening (3, 9) and the outer wall (5) being a reinforcing fiber web (7) for forming the reinforcing elements, the reinforcing fiber web (7) being integrally connected to the outer wall (5) and the fiber orientation of the reinforcing fiber web (7) adapted for the diversion of loads around the opening (3, 9) and for discharging loads acting on load introduction points at the opening (3, 9) into the side surface areas of the aircraft fuselage component (4)

Second fuselage fuselage component (4) according to claim 1, characterized in that the

Opening (3, 9) is provided for receiving an aircraft door.

Third fuselage fuselage component (4) according to claim 1 or 2, characterized in that the Verstärkungsfasergelege (7) is covered with a sacrificial layer for detecting damage, for example by impact stress.

4. fuselage fuselage component (4) according to claim 3, characterized in that the sacrificial layer is formed from a polymer foam with closed-pore surface.

A fuselage fuselage component (4) according to any one of the preceding claims, characterized in that the main direction of extension of the fibers of the reinforcing fiber fabric (7) is adapted to the direction of the load flows acting on the fuselage fuselage component (4) such that the fibers in the main direction extend the load flows from the Opening (3, 9) away distributed in the edge regions of the aircraft fuselage component (4) and the load flows from the edge regions of the aircraft fuselage component (4) and on the outer wall (5) acting loads around the opening in other edge regions of the aircraft fuselage component (4) derives.

Aircraft fuselage component (4) according to one of the preceding claims, characterized in that a first laminate (6) of a reinforcing fiber scrim (7) with a fiber orientation in the main direction of extension of the fiber is provided such that reinforcing fibers from the side edges (8a, 8b) of the aircraft fuselage component (Fig. 4) extend diagonally past the opening (3, 9) to the upper edges (8c) and lower edges (8d) of the aircraft fuselage component (4).

Aircraft fuselage component (4) according to one of the preceding claims, characterized in that a second structure (10) of a reinforcing fiber fabric (7) is provided with a fiber orientation in the main direction of extension of the fibers such that reinforcing fibers are adjacent to the opening (3, 9) Area from the opening (3, 9) away in the direction of the side edges (8a, 8b) extend, wherein a plurality of reinforcing fiber bundles are spaced from each other.

Aircraft fuselage component (4) according to claim 7, characterized in that the reinforcing fiber bundles extend with a part along an edge region of the opening (3, 9).

Aircraft fuselage component (4) according to any one of the preceding claims, characterized in that the aircraft fuselage component (4) has a separate supply element for the supply lines detachable from the outer wall (5) and integrally formed with the outer wall (5).