WO2005007910A1

WO2005007910A1 - Welded aluminium structural component with a metal induced tear deviation

Info

Publication number: WO2005007910A1
Application number: PCT/DE2004/001459
Authority: WO
Inventors: Frank Palm
Original assignee: Eads Deutschland Gmbh
Priority date: 2003-07-14
Filing date: 2004-07-07
Publication date: 2005-01-27
Also published as: WO2005007910B1; DE10331990A1

Abstract

The invention relates to a welded aluminium structural component (1) provided with a metal induced tear deviation, especially for aeroplanes. Said structural component comprises a film field (2) and at least one reinforcing element (3) which is secured to the film field (2) by welding, optionally, by using an additional welding substance. The film field (2), the reinforcing element (3) and the additional welding substance are made of aluminium materials and the aluminium material contains at least one of the film field components (2), a reinforcement element (3) and an additional welding substance, in addition to 0.05 2 wt.- % of one or several elements from the group made up of zircon (Zr), scandium (Sc), yttrium (Y), titanium (Ti), terbium (Tb), hafnium (Hf), niobium (Nb), tantalum (Ta), vanadium (V) and lanthanides. A metallurgical fine grain area is formed between the film field (2) and the reinforcing element (3).

Description

Welded aluminum structural component with metal induced crack deviation

The present invention relates to a welded aluminum structural component with a metal-induced crack deviation according to the preamble of claim 1, and to an aircraft pressure fuselage according to claim 12, which at least partially consists of such an aluminum structural component.

In aircraft construction, aluminum structural components for pressure fuselage shells, which essentially consist of skin-field stringer connections, have previously been produced using the riveting or adhesive production process. Not only are stringers riveted to a skin field, but also angle elements, so-called clips, which are used to fasten frames running in the circumferential direction of the pressure hull. As is well known, the clips are riveted to both the skin field and the stringer.

Recently, laser beam welding for producing skin field stringer connections has become established for the production of large-area aluminum structural components for aircraft pressure fuselage shells. Such welded skin-stringer connections are advantageously distinguished from the previous connections made by riveting or gluing by a lower weight and significantly reduced manufacturing times.

DE 196 39 667, for example, describes a method for laser beam welding of profiles on large-format aluminum structural components, in which two laser beams are used which are guided to the weld seam from two sides simultaneously by means of a seam search system. An aircraft shell part produced by means of laser beam welding is known, for example, from DE 19844 035. A disadvantage of the welded and thus integral or monolithic aluminum structural components is the crack propagation. A crack that progresses in the skin area of the welded skin-stringer connection not only cuts through the skin area but also the stringer and thus permanently weakens the overall structure. In contrast to riveted structures, in which a crack passes under the riveted stringer without extending to it, the welded aluminum structural component is to be regarded as a completely monolithic structure, with cracks in the skin area of the welded-on reinforcement elements stringer, clips and the like immediately and unchecked cut. This problem currently exists with all structural aluminum components welded using laser beams. Because of this disadvantage, the Airbus 318 currently only uses monolithic structural components welded in the area of the pressure fuselage lower shell, e.g. made of AlMgSiCu materials (AI material type 6xxx, according to the American alloy designation), since there are no tensile stresses in this area of the pressure fuselage and fail-safe concepts ( so-called "fail safe" concepts) are not required.

The present invention is therefore based on the object of creating a welded aluminum structural component in which the propagation of cracks onto a reinforcing element fastened to the skin field of the structural component is effectively prevented, so that it is suitable for a fail-safe "fail-safe" construction.

This object is achieved by a welded aluminum structural component that has a skin field and at least one reinforcing element that is attached to the skin field by means of welding, if necessary using a welding additive, whereby skin field ner reinforcing element and welding additive consist of aluminum materials and the aluminum material at least one of the components skin field, reinforcing element and welding additive additionally 0.05 - 2 wt .-% of one or more elements of the group consisting of zircon (Zr), scandium (Sc), yttrium (Y), titanium (Ti), terbium (Tb), hafnium (Hf), niobium (Nb), tantalum (Ta), vanadium (V) and the lanthanides , whereby a metallurgical fine grain zone is formed between the skin field and the reinforcing element.

This metal-induced fine-grain zone is formed in the area of the melted and re-solidified weld seam zone and thus preferably between the skin field and the weld metal (= the material melted during welding and solidified again after welding, which either consists of base material or of base material and welding additive) or between reinforcing element and weld metal. This advantageously effectively prevents the spread of a tear present in the skin field on the reinforcing element.

Due to the targeted alloying of the above-mentioned elements, a narrow seam is produced from very fine grains in the weld metal along the melting line (s) if the above-mentioned alloying elements are present in the appropriate amount in the melted and melted base material of the reinforcing element, skin field or welding additive , Because of this fine grain zone, a crack progressing in the skin field forms one or two secondary cracks in the weld metal instead of a crack in the reinforcing element when it hits the fine grain zone, which progress very slowly along the joint plane of the welded joint. This has the great advantage that the welded-on reinforcement element remains intact and is therefore available as a load-bearing path for a longer period of time.

The formation of the metallurgical fine grain zone on the basis of the alloyed elements is essentially based on the fact that during the welding process certain intermetallic phases (stoichiometric connections of the above-mentioned elements with the AI base material (generally of the type AI ₃ X)), which are present in the AI materials alloyed with these elements, are not completely melted in the area of the melting interface and are dissolved in the melt. At the welded interface between the weld metal and the reinforcing element or at the interface between the weld metal and the skin field, there remain, for example, non-melted, i.e. solid AlZr or AlSc / Zr phases, which at the interface in the melting zone as so-called heterogeneous nuclei (crystallization nuclei) during solidification of the weld metal. The resulting narrow seam of very fine, equidistant grains with its special cohesive properties causes the desired, changed crack propagation or crack deflection and thus prevents the undesired, premature damage to the reinforcing element.

It is advantageous here that the crack propagation is achieved effectively both in a welded connection without using a welding additive and in a connection welded using a welding additive by appropriately alloying the reinforcing element and / or skin field or alloying the reinforcing element, skin field and / or welding additive. This enables a flexible use of the idea according to the invention.

It is useful that aluminum alloys for the skin field and reinforcing element are aerospace alloys with the basic chemistry AICu ... (American alloy designation AA 2xxx), AlMg ... (American alloy designation AA 5xxx, which also includes AlMgSc alloys, for example), AlMgSi. .. (American alloy designation AA 6xxx), AlZn ... (American alloy designation AA 7xxx) and / or other Al material systems (American alloy designations AA 8xxx and AA 4xxx) as can be selected for AlLi alloys. These are typical aerospace certified alloys that are commonly used in the manufacture of aircraft pressure hulls. Thus, the present invention is for fen can be used. Thus, the present invention is applicable to all alloys commonly used in aviation or developed in the future, provided that they are modified or adapted by adding the above-mentioned elements. As a material form, this applies to forged materials (e.g. sheets, plates, etc.) and to cast materials (e.g. investment casting, die casting, but also rheo or thixo casting).

It is particularly advantageous that any combination of the aerospace alloy types listed above can be selected for the skin field and reinforcing element if at least one of the components skin field, reinforcing element and

Welding filler is modified according to alloy technology. In the event that the skin field and reinforcing element are not modified, a welding filler material, which is often necessary from a welding point of view, can alternatively be modified as a third element in the overall system, so that the desired fine grain zone is created, as explained above.

The skin field and reinforcing element can consist, for example, of an identical or similar type of aviation alloy (for example AA 6xxx). The skin field can consist, for example, of alloy 6013 and the reinforcement element of alloy 6110 A, the chemistry of the reinforcement element being adapted in accordance with the invention by adding Ti or Zr. Alternatively, the skin field and reinforcing element can also consist of different types of aviation alloy. The skin field can consist, for example, of an alloy of type 6xxx (eg 6056) and the reinforcing element can be of type 2xxx (eg 2195). In this case, the alloy 2195 is already defined with Zr due to its classification in the American AI alloy key. Alternatively, an AIMg6.3MnZrSc welding filler can be used, for example. Because a variety of combinations of aviation alloy types is possible is, the present invention is universally applicable in aviation, without major restrictions to certain types of alloy.

According to an alternative embodiment, a plurality of bores are provided in the reinforcement element along the joining plane of the welded connection. A bore diameter of 2 to 10 mm and a spacing of the bore center points of 4 to 1000 mm are expedient. The center of the drill holes are at a distance from the welded skin field of preferably> 1 mm to approx. 15 mm. The arrangement of bores along the joint plane reduces the spread of the crack deflected by the metallurgical fine-grain zone in the joint plane, since the side cracks are stopped by the holes when they strike bores.

It is particularly advantageous that the aluminum structural component according to the invention cannot be used only for the lower shell of an aircraft pressure fuselage, as before, but that a "fail-safe" construction is ensured due to the prevented crack propagation, so that the aluminum structural component also for the upper shell and The advantage of this is that the welded, integral construction can be extended to all areas of the aircraft fuselage, which in turn enables further manufacturing cost savings. In addition, the overall structural weight can be further reduced, which results in improved operating costs for aircraft.

Another advantage is that the reinforcing element welded to the skin field can be not only a stringer but also clips and / or push combs. In addition, the reinforcing element can also be a frame, which is known to be attached directly to the skin field in small aircraft. In the case of large aircraft, however, the frames are arranged on corresponding clips. net, which in turn are attached to both the skin field and the stringer. The reinforcing elements can alternatively be wire, matrix, fiber reinforced or the like. The present invention is therefore not restricted to specific reinforcement elements or partial areas in the aircraft, nor to a specific aircraft type.

It is also expedient that the reinforcing element can be fastened to the skin field by laser beam welding, arc welding as well as by other fusion welding methods. There is therefore no restriction with regard to the welding process, which makes the present invention universally and flexibly applicable.

Alternatively, after welding, the welded component can be subjected to a heat aftertreatment made up of various sub-steps in order to optimize properties. This can be done, for example, at a temperature of 50 - 450 ° C for a period of 15 - 1500 minutes.

The welded aluminum structural component according to the invention is used in particular in aircraft, where it at least partially forms lower, upper and / or side shells of the pressure fuselage.

The invention is described in more detail below with reference to the attached figures. The same components are provided with the same reference symbols in the figures.

Show it:

1 shows a schematic representation of the crack propagation in previous welded aluminum structural components known from the prior art; 2 shows a schematic representation of the propagation of cracks in an aluminum structural component according to the present invention;

3 shows an aluminum structural component according to FIG. 2 with bores arranged in the reinforcing element;

4 shows a micrograph of a melting interface of an aluminum structural component with an unmodified base material according to the prior art;

5 shows a micrograph of a melting interface of a laser-welded aluminum structural component with a base material (material 1) modified according to the invention;

6 shows a micrograph of a reflow interface of a laser-welded aluminum structural component with a base material (material 2) modified according to the invention; and

7 shows a micrograph of a melting interface of an arc-welded (TIG process) aluminum structural component with a base material (material 3) modified according to the invention.

1 shows a schematic three-dimensional representation of an aluminum structural component 1 which has a skin field 2 and at least one reinforcing element 3. For the sake of clarity, only a single reinforcing element 3 is shown in FIG. 1. The skin field 2 forms part of an aircraft pressure fuselage in a known manner, the reinforcing element 3 being a stringer which is attached to the skin field in the longitudinal direction of the aircraft. Hautfeld- Stringer connections are well known in the art, so further detailed description can be omitted.

In Fig. 1, the reinforcing element 3 is exemplified as a stringer, of course, the following explanation also applies to other reinforcing elements that are welded to the skin, such as clips or push combs. As is known, clips are used to fasten frames running in the circumferential direction of the aircraft pressure fuselage and are welded to both the stringer and the skin field. Instead of several individual clips, a push comb can also be used in a known manner. Since, in the case of small aircraft, the frames are known to be attached directly to the skin field, the reinforcing element shown schematically in FIG. 1 can also be a frame.

The skin area 2 of the aluminum structural component 1 shown in FIG. 1 consists of an aluminum material which is selected from one of the aviation-certified alloy types 2xxx, 4xxx, 5xxx, 6xxx, 7xxx and 8xxx. Alloy 6013 (skin field) is an example of a 6xxx skin field alloy. The reinforcement element 3, which is designed as a stringer in FIG. 1, likewise consists of an aluminum material which is selected from the same types of aviation alloys as the skin field alloy (e.g. 6110A).

FIG. 1 also shows a crack that spreads in the skin area 2 and that spreads in the direction of the reinforcing element 3. This is shown schematically by the arrow A running from right to left. In the case of the welded structural components known from the prior art, the crack A divides into two partial cracks B, C when the progressive crack A meets the reinforcing element 3, the one partial crack B (shown in dashed lines) under the reinforcing element 3 in the skin area 2 continues. The other partial tear leads to cutting the reinforcing element 3, which is represented by the arrow C. In other words, in the previous monolithic aluminum structural components produced by welding, a crack in the skin area 2 continues on the welded-on reinforcing elements 3 without braking and cuts through them. As a result, the reinforcing element 3, which is no longer available as a load-bearing path, can no longer maintain the strength of the overall structure, which necessitates the replacement of the structural component. As a result, this skin-field reinforcement element connection cannot be used where fail-safe ("fail safe") concepts (eg aircraft side and upper shells) are required.

FIG. 2 shows in an analogous manner a welded aluminum structural component 1 which, as in FIG. 1, has a skin field 2 and at least one reinforcing element 3, again only a single reinforcing element 3 in the form of a stringer being shown. As in the known welded, monolithic structural components according to FIG. 1, the skin field 2 and the reinforcing element 3 consist of aluminum materials (for example alloys of the type 2xxx, 4xxx, 5xxx, 6xxx, 7xxx and 8xxx), but now with the modified or adapted chemistry. The skin field 2 and the reinforcing element 3 can consist of identical or different alloy types. The following material pairing may be mentioned as an example: the skin field 2 consists of the 6xxx alloy 6013 and the reinforcing element consists of the 2xxx alloy 2195. In this case the alloy 2195 contains the alloy component Zr.

According to the invention, the aluminum material of the reinforcing element 3 and / or the skin field 2 is additionally 0.05-2% by weight of one or more elements of the group consisting of zirconium (Zr), scandium (Sc), yttrium (Y), titanium (Ti) , Terbium (Tb), hafnium (Hf), niobium (Nb), tantalum (Ta), vanadium (V) and the lanhanides. This special modification of the reinforcing element Material is caused that after welding at the interface between the weld metal and the reinforcing element 3 and / or at the interface between the weld metal and the skin field 2, a metallurgical fine grain zone is formed directly from the welding process.

The crack propagation is again shown schematically in FIG. 2, the crack propagating in the skin field 2 being represented by the arrow A ′ running from right to left. The crack A 'progresses in the skin field 2, strikes the reinforcing element 3 or rather the fine grain hem and is split there into one or more secondary cracks C, C ". The original crack (shown in broken lines; arrow B') progresses in the skin field 2 , whereby it passes under the reinforcing element 3, as in a riveted structure. However, the side cracks C, C "now progress very slowly along the joining plane of the welded connection between the skin field and the reinforcing element. A transmission or spreading of the crack A 'to the reinforcing element 3, which previously led to a severing of the reinforcing element 3, does not occur here. Consequently, the reinforcing element 3 remains intact as a reinforcing element here, i.e. the reinforcement element is unaffected by the crack and therefore remains load-bearing and therefore redundant, which corresponds to the requirements of a "fail-safe" design. Thus, the welded, monolithic structural component can not only be used longer, but can also be used in the aircraft pressure fuselage where such " Fail-safe "concepts are indispensable, such as in the pressure fuselage upper shell or in the pressure fuselage side shells.

In order to further reduce or stop the progression of the secondary cracks C, C "along the joint plane, additional bores 4 can be provided in the reinforcing element 3, which are arranged along the joint plane, adjacent to the welded skin field 2 (FIG. 3). The distance a between the drilling center points and the skin field 2 is typically> 1 mm to approx. 15 mm, at a bore diameter D of 2 - 10 mm and a distance d of the bore centers of 4 - 1000 mm.

According to a further embodiment, the welded structural component shown in FIGS. 2 and 3 can be produced in a known manner by welding the reinforcing element 3 to the skin field 2 using a welding additive. The crack propagation described in connection with FIG. 2 can be effected in an analogous manner by alloying the above-mentioned elements. It is sufficient here that only the welding filler is modified in accordance with the invention. Alternatively, reinforcement element 3 and / or skin field 2 can also be added. Of course, it is also possible to use an unmodified welding filler (e.g. AISM2). Then, as described above, the reinforcing element 3 and / or the skin field 2 must be modified in accordance with the invention in order to effectively prevent the crack from spreading.

For a more detailed explanation of the metallurgical effect that occurs at the interface between the skin field 2 and the welded-on reinforcing element 3, reference is made to FIGS. 4 to 7.

FIG. 4 shows a micrograph of a melting interface of an aluminum structural component with an unmodified base material according to the prior art. The aluminum structural component is a laser-welded connection made of material 6013, using a welding filler material of the type AISM2. The skin area 2 can be seen in the lower area of FIG. 4 and the weld metal can be seen in the upper area. There is no fine grain boundary along the melting line, which is marked with arrows in FIG. 4, but only a stalk-like (dendritic) solidification in the direction of the center of the weld seam. 5 shows a micrograph of a reflow interface of a laser-welded aluminum structural component with a base material modified according to the invention. The reinforcing element consists of the material 2195, which additionally contains approximately 0.12% by weight of zirconium (Zr) as an alloy component. An AlSil 2 filler material was used for welding. The weld metal can be seen in the lower area of FIG. 5 and the reinforcing element can be seen in the upper area. Furthermore, a more or less narrow hem of the fine grain zone (marked with arrows) can be seen in FIG. 5, which extends along the melting line.

Analogously, FIG. 6 shows the formation of a fine grain zone along a melting line in a connection of the material 2098 (reinforcing element) welded by means of a welding filler. In this case too, the reinforcing element material contains approximately 0.12% by weight of zirconium (Zr) as an alloy component. The formation of the fine grain zone is again marked with arrows, the weld metal being visible in the lower area of FIG. 6 and the reinforcing element in the upper area.

Fig. 7 shows a cross section of an arc welded connection of the material AlMgLiZnSc (1424). A welding material of the type AIMg6.3MnZrSc was used for welding. In this case, the base material of the skin field contains the elements Sc (approx. 0.25% by weight) and Zr (approx. 0.08% by weight) as an alloy component. The filler metal also contains these alloying elements. Again, a fine com zone forms in the area of the melting line.

State of the art example:

The combination AA is a common material combination for skin field and stringer

6013 T6 (skin field) and AA 6110A T6 (stringer). The alloy AA 6013 T6 the The Alcoa company has the following composition: 0.90% by weight of magnesium (Mg), 0.72% by weight of copper (Cu), 0.36% by weight of manganese (Mn), 0.27% by weight Iron (Fe) and the rest aluminum. The alloy AA 6110A T6 from Otto Fuchs is very similar to alloy 6013. Both alloys contain no additions according to the invention. There is only a low content of Ti («0.05%), because a so-called grain refiner (TiB2 wire) is added to these alloys with the aim of setting the grain size in the casting material as small as possible for processing reasons. This has been an established practice in the manufacture of semi-finished AI products for decades. So-called stringer-stiffened skin fields have now been produced from the two materials mentioned. Three stringers were welded on using Nd-YAG laser beam welding (stringer No. 1, 2 and 3, respectively). The following process parameters were used: Nd-YAG laser with 400 μm optical fiber focal length f = 150 mm => 0 focus = 300 μm - laser power: 2300 watts welding speed V _SC h = 2100 mm / min welding filler metal SZW = AlSil 2

The panels produced in this way were then tested on a tensile testing machine with regard to their cyclical crack propagation behavior. It was of primary interest how the crack behaves when it hits the stringer. For this purpose, a crack was made in the middle of the sample (e.g. by sawing it in), the middle stringer (No. 2) having also been severed. Under cyclic loading, the cracks then lengthened (transversely to the load direction). Depending on the stringer production, the crack (or both partial cracks of the left and right half of the sample) met stringer No. 1 and No. 3. The progressive crack severed stringer No. 1 and No. 3, and then the component collapsed in the testing machine. Example according to the present invention:

A material pairing according to the invention is e.g. a standard aerospace material for the skin field (alloy type AA 2524 according to the American alloy key, an AICu4Mg2Mn) and a Zr-containing (approx. 0.12% by weight) stringer material AA 2195 (AICu4Li1 MgAgZr). In an analogous manner to the previous example, a stringer-stiffened skin field with three stringers (No. 1, 2 and 3) was produced with identical process parameters without using a welding additive. The presence of Zr in the stringer alloy creates a narrow seam with fine, equidistant grains at the interface between the stringer base material and the weld metal, thus redirecting the attacking cracks. The crack does not grow into the stringer and the stringer can still maintain its load-bearing function. With corresponding investigations of the crack progress behavior in a tensile testing machine, whereby again the middle stringer (No. 2) e.g. was cut by sawing, it was found that stringers No. 1 and No. 3 are not cut.

The alloys known from DE 198 38 017, DE 198 38 018 and DE 198 38 015 are listed as further examples of the base materials of the reinforcing element and / or skin field, which alloys are combined with base materials from one of the typical aviation alloy types mentioned at the beginning can be.

Claims

claims

1.Welded aluminum structural component (1) with a metal-induced crack deviation, in particular for airplanes, comprising a skin field (2) and at least one reinforcing element (3) which is attached to the skin field (2) by means of welding, optionally using a welding additive, whereby the skin field (2), the reinforcing element (3) and the welding additive consist of aluminum materials and the aluminum material at least one of the components skin field (2), reinforcing element (3) and welding additive additionally 0.05-2% by weight of one or more Elements of the group consisting of zircon (Zr), scandium (Sc), yttrium (Y), titanium (Ti), terbium (Tb), hafnium (Hf), niobium (Nb), tantalum (Ta), vanadium (V) and contains the lanthanides, whereby a metallurgical fine grain zone is formed between the skin field (2) and the reinforcing element (3).

2. Welded aluminum structural component (1) according to claim 1, characterized in that the metallurgical fine grain zone forms in the region of the melted and solidified weld zone.

3. Welded aluminum structural component (1) according to claim 1 or 2, characterized in that the aluminum materials for skin field (2) and reinforcing element (3) are aviation alloys of the type 2xxx, 4xxx, 5xxx, 6xxx, 7xxx and / or 8xxx ,

4. Welded aluminum structural component (1) according to claim 3, characterized in that the skin field (2) and the reinforcing element (3) consist of an identical aerospace alloy type.

5. Welded aluminum structural component (1) according to claim 3, characterized in that the skin field (2) and the reinforcing element (3) consist of different types of aviation alloy.

6. Welded aluminum structural component (1) according to one of claims 1 to

5, characterized in that the lanthanides are in particular cerium (Ce), neodymium (Nd), europium (Eu), gadolinium (Gd), dysprosium (Dy), holmium (Ho) and / or erbium (Er).

7. Welded aluminum structural component (1) according to one of claims 1 to

6, characterized in that the reinforcing element (3) and / or the skin field (2) consist of forged, extruded and / or cast aluminum materials.

8. Welded aluminum structural component (1) according to one of claims 1 -

7, characterized in that a plurality of bores (4) is provided in the reinforcing element (3) along the joining plane of the welded connection.

9. Welded aluminum structural component (1) according to one of claims 1 to

8, characterized in that it forms at least part of a lower shell, an upper shell and / or a side shell of an aircraft pressure fuselage.

10. Welded aluminum structural component (1) according to one of claims 1 to 9, characterized in that the reinforcing element (3) is a stringer,

Frame, clip and / or push comb is.

11. Welded aluminum structural component (1) according to one of claims 1 to 10, characterized in that the reinforcing element (3) by laser beam welding, arc welding or another fusion welding process is attached to the skin field (2).

12. Aircraft pressure fuselage, characterized in that its lower, upper and / or side shells at least partially from the aluminum structural component

(1) according to one of claims 1 to 11.

13. Aircraft pressure fuselage according to claim 12, characterized in that a plurality of reinforcing elements (3) on the skin field (2) of the aluminum structural component (1) is attached by means of laser beam welding.