EP1492896A1

EP1492896A1 - Al-zn-mg-cu alloys welded products with high mechanical properties, and aircraft structural elements

Info

Publication number: EP1492896A1
Application number: EP03740569A
Authority: EP
Inventors: Frank Eberl; Christophe Sigli; Timothy Warner; Sjoerd Van Der Veen
Original assignee: Pechiney Rhenalu SAS
Current assignee: Constellium Issoire SAS
Priority date: 2002-04-05
Filing date: 2003-04-04
Publication date: 2005-01-05
Anticipated expiration: 2023-04-04
Also published as: ATE415498T1; WO2003085146A1; DE60324903D1; AU2003260003A1; FR2838135A1; DE03740569T1; FR2838135B1; ES2316779T3; US20060182650A1; US20050072497A1; JP2005530032A; EP1492896B1

Abstract

The invention concerns a rolled, extruded or stamped Al-Zn-Mg-Cu alloy product, characterized in that it contains (in wt. %): a) Zn 8.3 14.0 Cu 0.3 2.0 Mg 0.5 4.5 and preferably 0.5 3.6 Zr 0.03 0.15 Fe + Si < 0.25; b) at least one element selected from the group consisting of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, the content of each of said elements, if selected, ranging between 0.02 and 0.7 %; c) the rest being aluminium and unavoidable impurities, and in that it satisfies the following conditions d) Mg / Cu > 2.4 and e) (7.9 0.4 Zn) > (Cu + Mg) > (6.4 0.4 Zn), and preferably, Mg > 1.95 + 0.5 (Cu 2.3) + 0.16 (Zn - 6) + 1.9 (Si 0.04). The inventive products can be used in particular for manufacturing civil aircraft fuselage stiffeners.

Description

CORROYED ALLOY PRODUCTS AL-ZN-MG-CU

WITH VERY HIGH MECHANICAL CHARACTERISTICS,

AND ELEMENTS OF AIRCRAFT STRUCTURE

Technical field of the invention

The present invention relates to wrought products of Al-Zn-Mg-Cu type alloys with very high mechanical characteristics, with a Zn content greater than 8.3%, as well as aircraft structural elements incorporating such products.

State of the art

Alloys of the Al-Zn-Mg-Cu type (belonging to the family of 7xxx alloys) are commonly used in aeronautical construction, and in particular in the construction of the wings of civil aircraft. For the upper surfaces of the wings, for example a skin made of heavy plates made of alloys 7150, 7055, 7449, and optionally stiffeners made of profiles made of alloys 7150, 7055, 7349 or 7449. The alloys 7150, 7050 and 7349 are also used for the manufacture of fuselage stiffeners.

Some of these alloys have been known for decades, such as alloys 7075 and 7175 (zinc content between 5.1 and 6.1% by weight), 7050 (zinc content between 5.7 and 6.7%) , 7150 (zinc content between 5.9 and 6.9%) and 7049 (zinc content between 7.2 and 8.2%). They have a high elastic limit, as well as good toughness and good resistance to stress corrosion and exfoliating corrosion. More recently, it has become apparent that for certain applications, the use of an alloy with a higher zinc content may have advantages since this makes it possible to further increase the elastic limit. Alloys 7349 and 7449 contain between 7.5 and 8.7% zinc. Wrought alloys richer in zinc have been described in the literature, but do not seem to be used in aircraft construction. The article “Microstructure and properties of a new super-high-strength Al-Zn-Mg-Cu alloy C912” by YL Wu et al., Published in the journal Materials & Design, Vol 18, p. 211-215 (1998) presents an alloy Zn 8.7%, Mg 2.6%, Cu 2.5%, Si and Fe <0.05% (each) considered for the manufacture of structural elements for wing and fuselage.

US patent 5,560,789 (Pechiney Research) discloses an alloy with a composition Zn 10.7%, Mg 2.84%, Cu 0.92% which is transformed by spinning. The addition elements of this alloy, which is very charged with zinc, magnesium and copper, are difficult to dissolve because the dissolving temperature of the product is limited by the melting temperature of the phases having the lowest melting point: this product has a high mechanical resistance, but a very low elongation at break, due to the presence of coarse precipitates; said product is not very formable.

US Patent 5,221,377 (Aluminum Company of America) discloses several alloys of the Al-Zn-Mg-Cu type with a zinc content up to 11.4% and sufficiently loaded with copper. They are difficult to pour, and the additives are difficult to dissolve, which promotes the undesirable presence of coarse precipitates.

Furthermore, it has been proposed to use Al-Zn-Mg-Cu alloys with a high zinc content for the manufacture of hollow bodies intended to withstand high pressures, such as, for example, compressed gas cylinders. European patent application EP 020 282

Al (Société Métallurgique de Gerzat) discloses alloys with a zinc content of between 7.6% and 9.5%. European patent application EP 081 441 Al (Company

Métallurgique de Gerzat) discloses a process for obtaining such bottles. European patent application EP 257 167 Al (Société Métallurgique de Gerzat) notes that none of the known Al-Zn-Mg-Cu type alloys can safely and reproducibly meet the severe technical requirements imposed by this specific application; it proposes to move towards a lower zinc content, namely between 6.25% and 8.0%. The teaching of these patents is specific to the problem of compressed gas cylinders, in particular with regard to maximizing the bursting pressure of these cylinders, and cannot be transferred to other wrought products. In general, in alloys of the Al-Zn-Mg-Cu type, a high content of zinc, but also of Mg and Cu is necessary to obtain good static mechanical properties (yield strength, breaking strength) , provided that these elements can be solved. But it is also well known (see for example US 5,221,377) that when the zinc content in an alloy of the family 7xxx is increased beyond about 7 to 8%, problems are encountered with resistance to insufficient exfoliating corrosion and stress corrosion. More generally, it is known that the most loaded Al-Zn-Mg-Cu alloys are liable to cause corrosion problems. These problems are generally resolved using specific thermal or thermomechanical treatments, in particular by pushing the tempering treatment beyond the peak, for example during a T7 type treatment. However, these treatments can then cause a drop in static mechanical characteristics. In other words, for a minimum level of corrosion resistance targeted, the optimization of an alloy of the Al-Zn-Mg-Cu type must seek a compromise between the static mechanical characteristics (elastic limit R _p o, ₂ , breaking limit R _m , elongation at break A) and the damage tolerance characteristics (toughness, speed of crack propagation, etc.). Depending on the minimum level of corrosion resistance targeted, a state close to the tempered peak is used (states-T6), which generally offers a tenacity compromise - R _p o, 2 favoring static mechanical characteristics, or the tempering is pushed to beyond the peak (T7 states), seeking a compromise favoring tenacity.

Whatever approach is chosen, the development and use of such products poses two problems: on the one hand, these alloys heavily loaded with zinc and magnesium are difficult to cast and transform, in particular by spinning, rolling or forging . For example, the maximum effort that a spinning press can provide can be a limiting factor. In particular, among the alloys of the 7xxx family, the alloys 7349 and 7449 require very high spinning forces. On the other hand, there are applications in which the aptitude of said extruded and rolled products for shaping is an important factor. This is particularly the case for fuselage stiffeners. Consequently, it is to be feared that the search for an alloy with an even higher mechanical resistance than those of alloys 7349 and 7449 will result in an alloy which is difficult to cast and to transform, and difficult to shape.

Problem

The problem to which the present invention attempts to respond is to propose new wrought products of high-grade Al-Zn-Mg-Cu type alloy. zinc, greater than 8.3%, and in particular of extruded products, which are characterized by a very high breaking strength, a very high yield strength, sufficient resistance to corrosion, good formability , and which can be manufactured industrially under conditions of reliability compatible with the high requirements of the aeronautical industry.

Objects of the invention

The Applicant has found that the problem can be solved by adjusting the concentration of the Zn, Cu and Mg addition elements and certain impurities (in particular Fe and Si) in a fine manner, and possibly adding other elements.

A first object of the present invention consists of a rolled, extruded or forged product of Al-Zn-Mg-Cu alloy, characterized in that it contains (in percent by mass): a) Zn 8.3 - 14.0 Cu 0.3 - 2.0

Mg 0.5 - 4.5 and preferably 0.5 - 3.6 Zr 0.03 - 0.15 Fe + Si <0.25 b) at least one element selected from the group consisting of Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, the content of each of the said elements, if selected, being between 0.02 and 0.7%, c) the aluminum and impurities inevitable, and in that it satisfies the conditions d) Mg / Cu> 2.4 and e) (7.9 - 0.4 Zn)> (Cu + Mg)> (6.4 - 0, 4 Zn). A second object of the present invention consists of a rolled, extruded or forged product of Al-Zn-Mg-Cu alloy, characterized in that it contains (in percent by mass): a) Zn 9.5 - 14.0 Cu 0.3 - 2.0 Mg 0.5 - 4.5 and preferably 0.5 - 3.6

Fe + Si <0.25 b) at least one element selected from the group consisting of Zr, Se, Hf, La, Ti,

Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, Cr, Mn, the content of each of the said elements, if selected, being between 0.02 and 0.7%, c ) the remaining aluminum and unavoidable impurities, and in that it satisfies the conditions d) Mg / Cu> 2.4 and e) (7.9 - 0.4 Zn)> (Cu + Mg)> (6.4 - 0.4 Zn).

A third object of the present invention is an aircraft structural element which incorporates at least one of the said products, and in particular a structural element used in the construction of the fuselage of civil aircraft, such as a fuselage stiffener.

Description of the figures

Figure 1 shows the section of the TI profile.

Figure 2 shows the section of the T2 profile.

Figure 3 shows the section of profile T3.

Figure 4 shows the section of profile T4. Figure 5 shows the section of profile T5.

In Figures 1, 2, 3, 4 and 5, the dimensions shown are approximate and expressed in millimeters.

In Figures 1, 2, 3 and 4 the letter a denotes the "sole" of the profile, and the letter b the "heel". FIG. 6 schematically shows the area of a fuselage stiffener which has undergone shaping by routing. The benchmarks are as follows: a Routing depth b Routing width c Upper flange: appearance of large plane deformations d Lower flange: appearance of large plane deformations

Figure 7 shows schematically the location on the TI profile where the sample is taken for the 3-point bend test.

Figure 8 shows schematically the definition of the folding angle. Figure 9 schematically shows the important geometric parameters for the three-point bending test.

Figure 10 shows schematically a crack with a length of two stiffeners with the central stiffener broken.

Figure 11 schematically shows the buckling test. Figure (b) corresponds to a rotation A- A of 90 °.

FIG. 12 compares the buckling stresses for different types of Z-shaped stiffeners according to the invention (gray bars) and according to the prior art (white bars), for the same geometry.

Detailed description of the invention

Unless otherwise stated, all information relating to the chemical composition of the alloys is expressed in percent by mass. Consequently, in a mathematical expression, "0.4 Zn" means: 0.4 times the zinc content, expressed in percent by mass; this applies mutatis mutandis to other chemical elements. The designation of alloys follows the rules of The Aluminum Association. The metallurgical states are defined in European standard EN 515. Unless otherwise stated, the static mechanical characteristics, that is to say the tensile strength R _m , the elastic limit R _p o, ₂ , and the elongation at the rupture A, are determined by a tensile test according to standard EN 10002-1. The term “spun product” includes so-called “drawn products”, that is to say products which are produced by spinning followed by drawing. The applicant, in the course of a number of preparatory studies, came to the conclusion that a new material having a compromise between mechanical strength and significantly better formability should in any event have a sufficient zinc content, typically higher around 8.3%, and preferably above 9.0%. This condition is however not sufficient.

In the context of the present invention, the Applicant has found a very particular field of composition which allows the production of wrought products, and in particular of spun products, which have both very high static mechanical characteristics, corrosion resistance acceptable, and good formability. The Applicant has thus been able to develop spun products which can be used very advantageously as stiffeners in the fuselage of civil aircraft. In this application, damage tolerance is not a limiting factor, and we can therefore allow ourselves to optimize the elastic limit and the breaking limit at the expense of damage tolerance, while taking care not to degrade corrosion resistance. On the other hand, the fact of pushing the elastic limit and the breaking limit as far as possible, making it possible to lighten the structure of the aircraft, usually results in a deterioration in the fitness for shaping. However, the fuselage stiffeners are subjected to complex and very specific shaping operations. To develop a more resistant alloy for fuselage stiffeners, it is therefore necessary to ensure that the formability does not deteriorate compared to known alloys, or, preferentially, is better than that of known alloys .

According to the invention, the problem is solved by fine adjustment of the contents of the alloying elements and certain impurities, and by adding a controlled concentration of certain other elements to the composition of the alloy.

The present invention applies to Al-Zn-Mg-Cu alloys containing: Zn 8.3 - 14.0 Cu 0.3 - 2.0 Mg 0.5 - 4.5 as well as certain other elements specified below , and the rest being aluminum with its inevitable impurities. The alloys according to the invention must contain at least 0.5% magnesium, since it is not possible to obtain satisfactory static mechanical characteristics with a lower magnesium content. According to the Applicant's observations, with a zinc content of less than 8.3%, no result is obtained which is better than those obtained with known alloys. Preferably, the zinc content is greater than 9.0%, and even more preferably greater than 9.5%. However, it is necessary to respect certain relationships between certain elements, as explained below. In another advantageous embodiment, the zinc content is between 9.0 and 11.0%. In any event, it is not desired to exceed a zinc content of approximately 14%, because above this value, whatever the magnesium and copper content, the results are not satisfactory.

The addition of at least 0.3% copper improves the corrosion resistance. A content of at least 0.6 is preferred. But to ensure satisfactory dissolution, the Cu content should not exceed approximately 2%, and the Mg content should not exceed approximately 4.5%; a maximum content of 3.6% is preferred for magnesium. In an advantageous embodiment, the copper content is between 0.6% and 1.2 while the magnesium content is between 2.5% and 3.4%. In another advantageous embodiment, the copper content is between 0.8% and 1.5 while the magnesium content is between 2.2% and 3.0%. As will be explained below, the ratio between the magnesium and copper contents must meet certain criteria.

The Applicant has found that in order to solve the problem posed, it is necessary to take into account, in an alloy of the Al-Zn-Mg-Cu type, several additional technical characteristics.

First of all, the alloy must be sufficiently loaded with addition elements capable of precipitating during maturation or tempering treatment, in order to be able to exhibit advantageous static mechanical characteristics. For this, according to the Applicant's observations, in addition to the minimum and maximum limits for the zinc, magnesium and copper contents indicated above, the content of these addition elements must fulfill the condition Mg + Cu> 6.4 - 0.4 Zn. To reinforce this effect, a sufficient content of so-called anti-recrystallizing elements must be added. More specifically, for alloys with more than 9.5% zinc, at least one element selected from the group comprising the elements Zr, Se, Hf, La, Ti, Y, Ce, Nd, Eu, Gd, must be added. Tb, Dy, Ho, Er, Yb, Cr, Mn with, for each element present, a concentration of between 0.02 and 0.7%. It is preferable that the concentration of all of the elements of said group does not exceed 1.5%.

These anti-recrystallizing elements, in the form of fine precipitates formed during thermal or thermomechanical treatments, block the recrystallization. However, the applicant has found that excessive precipitation must be avoided during the quenching of the wrought product, and especially when the alloy is heavily loaded with zinc (Zn> 9.5%). A compromise must therefore be found with regard to the content of anti-recrystallizing elements.

According to the invention, for alloys with a zinc content of between 8.3% and 9.5%, it is necessary to add zirconium with a content of between 0.03% and 0.15%, and in addition at least an element selected from the group comprising the elements Se, Hf, La, Ti, Y, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Yb, with, for each element present, a concentration of between 0, 02 and 0.7%. In an advantageous embodiment, titanium is chosen, alone or associated with one or more other elements of said group. The Applicant has found that for said anti-recrystallizing elements, it is advantageous, whatever the zinc content, not to exceed the following maximum content: Cr 0.40; Mn 0.60; Se 0.50; Zr 0.15; Hf 0.60; Ti 0.15; This 0.35 and preferably 0.30; Nd 0.35 and preferably 0.30; Eu 0.35 and preferably 0.30; Gd 0.35; Tb 0.35; Ho 0.40; Dy 0.40; Er 0.40; Yb 0.40; Y 0.20; The 0.35 and preferably 0.30. Advantageously, the total of these elements does not exceed 1.5%.

The Applicant has found that to improve the breaking limit and the elastic limit, it is preferable to respect a Mg / Cu ratio> 2.4, and preferably at least 2.8, even more preferably 3.5 or even 4 , 0. Another technical characteristic is linked to the need to be able to produce industrially wrought products under conditions of reliability compatible with the high requirements of the aeronautical industry, as well as under satisfactory economic conditions. It is therefore necessary to choose a chemical composition which minimizes the occurrence of cracks or slots during the solidification of the plates or billets, said cracks or slots being unacceptable defects leading to the scrapping of said plates or billets. The Applicant has noted during numerous tests that this occurrence of cracks or slots was much more likely when the 7000 alloys finished solidifying below 470 ° C. To significantly reduce the probability of cracks or cracks occurring to an industrially acceptable level, it is better to choose a chemical composition such as

Mg> 1.95 + 0.5 (Cu - 2.3) + 0.16 (Zn - 6) + 1, 9 (Si - 0.04). This criterion is called in the context of the present invention the "flowability criterion". The alloys produced according to this variant of the invention complete their solidification at a temperature of between 473 ° C and 478 ° C, and make it possible to achieve industrial reliability in the processes for preparing the metal (that is to say a consistency of the quality of the cast plates or billets) compatible with the high requirements of the aeronautical industry.

Another technical characteristic of the invention is linked to the need to minimize as much as possible the quantity of insoluble precipitates (which are typically ternary or quaternary phases Al-Zn-Mg-Cu of type S, M or T) after the homogenization and dissolution treatments, as this reduces the toughness, the elongation at break and above all the aptitude for shaping; for this, we choose a content of Mg, Cu and Zn such that Mg + Cu <7.9 - 0.4 Zn. According to the Applicant's observations, there is no disadvantage in positioning itself very close to this limit represented by the relationship Mg + Cu <7.9 - 4.4 Zn, but its exceeding leads to a rapid deterioration of the aptitude for deep shaping by crimping, which is one of the advantages of the products according to the invention. And finally, the incorporation of a small amount, between 0.02 and 0.15% per element, of one or more elements chosen from the group composed of Sn, Cd, Ag, Ge, In makes it possible to improve the response of the alloy to the tempering treatment, and has beneficial effects on the mechanical strength and on the corrosion resistance of the product. A content of between 0.05 and 0.10% is preferred. Among these elements, money is the preferred element. In the case of a profile, the addition of one or more anti-recrystallizing elements, such as scandium, is particularly advantageous; such an effect is also observed in the case of heavy plates. The profiles also benefit from an increase in their mechanical strength, which is all the greater when the width or thickness of the profile is small; this so-called "press effect" is well known to those skilled in the art. The Applicant has found that when the anti-recrystallizing element added is scandium, a content of between 0.02 and 0.50% is advantageous.

The products according to the invention are in particular spun products. They can be used advantageously for the manufacture of structural elements in aircraft construction. A preferred application of the products according to the invention is the application as a structural element in the fuselage of a civil aircraft. These elements, in particular the stiffeners, are firstly dimensioned in mechanical strength. For this use, damage tolerance is not usually a property which enters into the dimensioning, insofar as this is of a reasonable level: we can, if necessary and up to a certain point, optimize mechanical resistance to the detriment of damage tolerance, and without fear of reducing the usefulness of the product. Corrosion resistance should always remain at an acceptable level. The increase in the mechanical strength of said fuselage stiffeners makes it possible, at the choice of the manufacturer, to reduce their weight, or to have, at equal weight, a more rigid fuselage structure. This can make it possible, by increasing the spacing between two neighboring stiffeners (within the limit of the crease resistance of the fuselage sheets), to reduce the number of stiffeners, which leads to a reduction in the number of fasteners or assembly points between stiffener and wing skin. This can prove to be very advantageous, since fasteners or assembly points, such as rivets or bolts, are an important part of the cost of manufacturing such structures. A particularly advantageous use of the product according to the invention is therefore the application as a structural element in the field of aeronautical construction, and more precisely in the construction of aircraft comprising a fuselage assembled from a plurality of stiffeners and a plurality of sheets, at least part of said stiffeners being structural elements according to the invention. Such an aircraft is characterized by a lighter structure, but at least as rigid, or by a more rigid structure, but not heavier than existing aircraft.

Like the fasteners between structural elements of different type (for example stiffener and fuselage skin), it is desirable to minimize the number of assemblies between two structural elements of the same type, and in particular between two stiffeners. To this end, it is advisable to use sheets or extruded products of relevant size as large as possible; in the context of spun products, this relevant dimension is essentially the length. However, the manufacture of very long sections of highly charged Al-Zn-Mg-Cu alloys requires excellent control of the casting, spinning and heat treatment processes, and may require an adaptation of the chemical composition according to the invention. More particularly, the Applicant has found that the product according to the invention can be obtained with a reduced spinning pressure compared to known products with comparable zinc content, which makes it possible to manufacture profiles of greater length.

It is known that the aviation authorities require a structure designed to withstand limit loads despite significant damage; the damage chosen is a crack with a length of two stiffeners with the central stiffener broken (see figure 10). It has been noted by the applicant that the residual strength of the fuselage panels working in traction can benefit from the high-resistance stiffeners according to the invention. The use of stiffeners according to the invention as a structural element in fuselage panels can improve the residual strength of the structure, because they close the crack in the skin, which prevents unstable rupture as a preventive measure. . The residual strength of the cracked panel is thus improved. This effect can be used either to increase the safety margin in constructions where stiffeners are replaced by stiffeners according to the invention, either to lower the weight of the construction by using stiffeners with reduced sections and thinner fuselage sheets, and / or larger stiffener spacings.

Furthermore, the aviation authorities request that the structure be designed to withstand an ultimate load for 3 seconds without excessive deformation. However, plastic deformation is allowed. This leads to post-buckling designs for fuselage panels in critical stability locations. Although buckling of perfect columns (Euler theory) or very thin real structures is essentially an elastic phenomenon (governed by Young's modulus), post-buckling designs show plastic deformation and can benefit from a increased yield strength. This buckling test is shown in Figure 11.

It has been noted by the applicant that the shear and compression stability of the fuselage panels working in compression and / or shear can benefit from the high resistance of the stiffeners according to the invention. The use of stiffeners according to the invention as a structural element in an aircraft fuselage panel can improve the shear and compression stability of the fuselage panels, because these stiffeners show higher buckling stability. This effect can be used either to increase the safety margin in constructions where stiffeners are replaced by stiffeners according to the invention, or to lower the weight of the construction by using stiffeners with reduced sections and fuselage plates more thinner, and / or larger stiffener spacings. It is also possible to obtain an increase in the spacing of the rivets, which decreases the cost of assembling the structure.

Table 17 shows parameters of different geometries of stiffeners used for the calculations. Figure 12 compares the predicted buckling stresses for these different geometries from Zl to Z8 (from left to right). Another problem which arises in particular when using said products as fuselage stiffeners is their aptitude for shaping.

One form of shaping used during the industrial manufacture of fuselage stiffeners from profiles is the joggle. This is an introduction to a step localized over an area of a few millimeters (see Figure 6). This can be done, in the case of profiles according to the invention, either hot (preferably at 130 ° C), or cold. In the case of cold routing, the profile delivered in the W state (unstable) will advantageously be dissolved, followed by quenching. Then the shaping is carried out by routing. Cold jogging does not allow shaping as deep as hot jogging, but when it is applicable, it is often more practical.

Routing as an industrial shaping process is not suitable for use in the study of materials under development. However, it is known that the failure of the material during routing is directly linked to maximum plane deformations that can be supported by the material. This makes it possible to evaluate the suitability of a material for shaping by routing using the 3-point bending test. According to standard DIN 50111 (September 1987, in particular section 3.1), the sample must be sufficiently large compared to its thickness to be in conditions of plane deformations in the center of the test pieces.

In the context of the present invention, in order to evaluate the formability at 130 ° C (lukewarm formability of the product in the final state), the flat test piece is deformed in an oven at 130 ° C until the start of the fall applied force (which means crack initiation), always ensuring that the temperature of the sample is at 130 ° C. Since the deformation is done hot, the speed of deformation is a parameter which influences the result. It was fixed by a cross speed of 50 mm / min. The higher the folding angle (see definition in FIG. 8), the higher the aptitude for shaping by drilling. For mechanical reasons, it is important that the samples to be compared have the same thickness. If two samples of different thickness are to be compared, the face is machined in compression to the necessary thickness. In the case of a profile, the sampling of the sample from which the flat specimen is prepared is done at a representative location as shown in Figure 7 for the TI profile.

The 3-point bending tests at 130 ° C are carried out on the T6x state or on the T7x state of the product. Nevertheless, it is possible to characterize the formability in the raw quenching state W with this test, provided that the time between the tensile stress relieving which follows the quenching and the execution of the three-point bending test is controlled. In the case of extruded products, the bending angle at 130 ° C is expressed as an average value calculated from individual measurements carried out on samples taken at different locations distributed over the length of the profile.

A particularly preferred product according to the present invention is a spun product which, in the T6511 state, measured on test pieces taken from a flat area, has a bending angle, measured at 130 ° C. by a 3-point bending test according to DIN 50. 1 11 (section 3.1) on a sample 1.6 mm thick, at least 34 °, and an elastic limit R _p o, ₂ of at least 720 MPa, and preferably a bending angle of at least 35 ° and an elastic limit of at least 750 MPa. The static mechanical characteristics (R _p o.2, Rm and A) depend very little on the thickness of the section for thicknesses up to around 60 mm.

Another particularly advantageous product according to the invention is a spun product which has in the T76511 state, measured on test specimens taken from a flat area, a bending angle, measured at 130 ° C. by a 3-point bending test according to DIN. 50 111 (section 3.1) on a sample 1.6 mm thick, at least 36 °, and an elastic limit R _p o, ₂ of at least 660 MPa, and preferably at least 670 MPa. This product can be used in cases where the corrosion resistance must be at least EB level during an EXCO test (ASTM G34 standard) carried out on non-machined samples.

These two preferred products are particularly suitable for the manufacture of fuselage stiffeners for civil aircraft. As indicated above, the Applicant has surprisingly found that, compared with known products, including those with a comparable zinc content, the products according to the invention show good aptitude for hot forming. On the other hand, the aptitude for cold forming in an unstable state W after re-solution and quenching is slightly less good. For the manufacture of structural elements of aircraft, such as fuselage stiffeners, the Applicant therefore prefers the hot forming method, if said shaping is deep.

The products according to the invention can also be used as a structural element for a floor, and in particular as a floor, aircraft profile, as well as, in the form of profiles, as seat rails. The seat rails are generally very long profiles, generally arranged parallel to the length of the cabin, on which the rows of seats are fixed in a commercial aircraft. According to the invention, it is possible to obtain seat rails in the T76511 state with a breaking strength of the seat fixing zone (ie the heel of a type “I” profile) whose breaking strength reaches 670 MPa and even 680 MPa, and the elastic limit of which reaches 640 MPa and even 660 MPa. The seat rails of commercial aircraft must resist corrosion by corrosive food liquids under high mechanical stresses, and the seat rails according to the invention effectively show good resistance to corrosion under stress determined according to standard ASTM G47.

The use of structural elements according to the invention for the construction of aircraft makes it possible to significantly lighten the structure of said aircraft, which makes it possible to increase their payload capacity, or to reduce their fuel consumption.

The invention will be better understood with the aid of the examples, which however are not limiting. Examples

Example 1:

Several Al-Zn-Mg-Cu alloys were prepared by semi-continuous plate casting, and they were subjected to a conventional transformation range, comprising a homogenization step, the parameters of which were determined according to the teaching of the patent. US 5,560,789, followed by hot rolling, a solution dissolving step followed by quenching and stress relieving operations, and finally an income in the state T651. 20 mm thick sheets were thus obtained in the T651 state. The compositions of the sheets making up this test are shown in Table 1.

Table 1

The static mechanical characteristics were determined by a tensile test according to standard EN 10002-1. The Kic toughness was determined according to ASTM E399. The results are shown in Table 2:

Table 2

It is found that the sheet C, according to the invention, presents a good compromise between mechanical strength and elongation. Compared to sheet D, outside the invention, its mechanical strength is significantly better. Compared to sheet A, made of alloy 7449 according to the state of the art, alloy C has a very improved mechanical resistance. The fact that the toughness of sheet C is less good than that of sheet B limits its application to certain uses for which the toughness is not dimensioning, but which require both excellent mechanical strength and good suitability for shaping. Compared to sheet B, outside the invention, the elongation at break of sheet C is significantly better. Furthermore, in order for sheet B to be able to achieve the results indicated in Table 2, it must be subjected to a fairly long solution treatment which does not lend itself to the requirements of industrial production. And even, we note that there are too many coarse phases in the product which have a detrimental effect on the homogeneity of the mechanical properties, both within the same batch and within the same product (sheet or profiled); this could prohibit the use of product B as an aircraft structural element.

Example 2:

Several alloy plates were cast, the composition of which is indicated in Table 3, with an Si content approximately equal to 0.04% for all the alloys.

The alloys Gl, G2, G3 and G4 and B are outside the present invention. The composition of alloys B and D, outside the invention, is indicated in Example 1, as well as that of Example C (according to the invention). All of these alloys showed satisfactory flowability during the tests, that is to say that cracks or cracks were not observed during the casting tests on an industrial scale.

The alloys G5, G6, G7, G8 are outside the present invention, and the alloy G9 is an alloy 7060 according to the state of the art; these alloys presented cracks during the casting tests. The difficulties appearing during the casting of these alloys do not necessarily make the wrought products obtained from these plates unfit for use, but are at the origin of additional costs because the implementation (i.e. the quantity of salable metal compared to the quantity of metal charged, a parameter which is directly linked to the quantity of discarded plates) will be greater than for the alloys corresponding to the preferred field of the invention. In addition, the propensity of these alloys to form slots during their solidification makes it very difficult to make the casting process reliable as part of a quality assurance program by statistical control of the processes.

It is found that the alloys of the 7xxx series having a very pronounced propensity for the formation of cracks or cracks in casting have a magnesium content lower than the critical magnesium content; this critical value was obtained by calculating the limit value in Mg defined by the flowability criterion.

Table 3

Example 3

Spinning billets were prepared with alloys the composition of which is summarized in Table 4. The alloys were homogenized as follows: Samples Ql and Q2: 4 h at 465 ° C + 20 h at 476 ° C

Samples Q3 and Q4: 4 h at 465 ° C + 20 h at 471 ° C

PI samples at P3: 8 p.m. at 471 ° C.

The M, T and S phases were completely dissolved during the homogenization treatment; this has been verified by differential enthalpy analysis (see US Patent 5,560,789).

The diameter of the billets was 200 mm for the P3 and Ql to Q4 billets, and 155 mm for the PI and P2 billets. Table 4

From these homogenized and peeled billets, five types of sections TI, T2, T3, T4 and T5 were developed, the sections of which are shown in Figures 1, 2, 3, 4 and 5. The temperature of the container and the the tool was greater than 400 ° C, the spinning speed was less than 0.50 m / min. The maximum extrusion pressures are summarized in Table 5. It is surprisingly found that for the alloys according to the invention, the extrusion pressure does not increase, and, surprisingly, even decreases, for certain types of profiles, when the magnesium content increases.

Table 5

Profiles Q1 to Q4 were dissolved at 471 ° C, profiles PI to P3 at 472 ° C (profiles TI, T2 and T3). The RI and R2 profiles were treated under comparable conditions. All the profiles were soaked in water and fractionated with a permanent elongation of between 1.5 and 2%. Products are obtained in the T6511 or T76511 state.

The static mechanical characteristics are summarized in Table 6, for three different thicknesses of the specimen in the T6511 state, taken from a flat area of the profile. These states were obtained by artificial aging under the following conditions:

Alloys Ql and Q2: 18 hours at 120 ° C

PI alloys at P3, Q3 and Q4: 36 hours at 120 ° C. Table 6

The characteristics in the T76511 state, obtained by artificial aging under the following conditions:

QlàQ4: 12 h at 120 ° C + 8 h at 150 ° C PI: 12 h at 120 ° C + 10 h at 156 ° C are summarized in table 7.

Table 7

It can be seen that compared to the PI alloy, the Ql and Q2 alloys have a significantly higher mechanical resistance.

The corrosion resistance was characterized according to the EXCO test (standard ASTM G34) of the products Ql and Q2 in the state T6511 (samples not machined at the start of spinning) was of level EA or EB and overall at least as good or better than that of samples PI to P3 and Q3 and Q4.

For RI and R2, we find the following mechanical characteristics:

Table 8

Example 4

The formability of the TI type profiles of Example 3 was studied using the 3-point bending test according to DIN 50 111 of September 1987 (section 3.1). The sample collection location, a flat area, is shown in the figure

7. The important parameters of the 3-point bending device are shown in the figure

9. The test was carried out at 130 ° C.

The T6511 and T76511 states have been tested. The values of the folding angle α (defined in Figure 8) are presented in Table 9. These are average values calculated from half a dozen individual measurements made on samples taken at different locations distributed along the length of the profiles. Table 9

In all cases, the profiles according to the invention (Q1 and Q2) have a formability comparable to that of the profiles according to the state of the art (Q3 and PI).

Example 5:

The aptitude for cold forming of samples similar to those of example 4 was studied (in the unstable state W after re-solution and quenching), with the same 3-point bending technique, but at Room temperature. There is a small dispersion of the folding angle α (defined in Figure 8) as a function of the length of the profiles. Table 10 refers to the values measured in state W.

Table 10

Example 6

Lamination plates were produced by a process similar to that described in Example 1. The chemical composition is given in Table 11. By a process similar to that described in example 1, plates with a thickness of 25 mm were prepared by hot rolling. They were dissolved for 2 hours at a temperature between 472 and 480 ° C, quenched and fractionated with a permanent elongation between 1.5 and 2%. Then, the sheets were subjected to a tempering treatment at a temperature of 135 ° C.

Table 11

The following mechanical characteristics were obtained

Table 12

It has been verified that for sheets N and K, the 12 hr income leads to state T6. For significantly longer incomes, the parameters R _p o, 2 (L) and R _m (D degrade. We note that with equal Zn content, and with a neighboring Mg / Cu ratio, the sheet N (with 0 , 10% scandium) shows better static mechanical characteristics than sheet M, without scandium. With equal zinc content, and with equal scandium content, N sheet with a high Mg / Cu ratio shows better values of R o, 2 (D and R _m ( _L >) than K sheet.

Example 7:

The resistance to stress corrosion resistance was characterized for several profiles produced according to example 3.

Table 13 collates the results obtained:

Table 13

It is found that the products according to the invention show resistance to corrosion under stress which is entirely satisfactory.

Example 8:

Extruded sections of 7349 or 7449 alloys were prepared, with and without the addition of scandium, according to a process similar to that of Example 3.

Table 14 indicates the chemical compositions, Table 15 the mechanical characteristics obtained. Table 14

Table 15

By comparing with the results of Example 3, it is found that the products according to the invention have a mechanical strength (R _m , R _p o, 2) improved compared to products XI and X2 according to the state of the art.

Example 9: Aircraft seat rails were manufactured from billets of chemical composition RI and Ql according to the preceding examples. These profiles are of type "I" and include a sole, a central zone (core) and a heel (on which the seats are fixed). The thickness of the central zone was of the order of 2 mm, the height of the profile of the order of 65 mm.

Table 16 shows the static mechanical characteristics in the T76511 state.

Table 16

Stress corrosion tests (in the cross direction) show good resistance during a test according to ASTM G47.

Example 10:

Numerical models simulating the damage tolerance of the high resistance stiffeners according to the invention were evaluated in order to determine the residual resistance of a fuselage panel. The aviation authorities are asking for a structure designed to withstand limit loads despite significant damage; the damage chosen is a crack with a length of two stiffeners (14, 16) with the central stiffener (18) broken (see Figure 10). It has been noted by the applicant that the residual strength of the fuselage panels working in traction can benefit from the high-resistance stiffeners according to the invention. The use of stiffeners according to the invention as a structural element in fuselage panels can improve the residual strength of the structure, because they close the crack (12) in the skin (18), which makes it possible to avoid to prevent unstable rupture. The residual strength of the cracked panel is thus improved. This effect can be used either to increase the safety margin in constructions where stiffeners are replaced by stiffeners according to the invention, or to lower the weight of the construction by using stiffeners with reduced sections and fuselage plates more thinner, and / or larger stiffener spacings. The rupture of the skin is governed by the stress intensity factor at the tip of the crack. For upper surface fuselages with a stiffener spacing of 200 mm and a stiffening rate (stiffener section / total section) of 0.25, the stress intensity factor for a crack with a length of two stiffeners with the stiffener central broken in a panel assembled with stiffeners according to the invention will be reduced by 5% in comparison with a panel with stiffeners made with the widely used 2024 T3 alloy. For longer cracks, the stiffener in 2024 will be used more and more in the plastic field in comparison with new stiffeners which have not even reached the elastic limit. The difference in the stress intensity factor can be as high as 15%.

Example 11:

Numerical models for fuselage panels working in compression and / or shear were evaluated in order to determine the stability in shear and compression. Aviation authorities are requesting that the structure be designed to withstand an ultimate load for 3 seconds without excessive deformation. However, plastic deformation is allowed. This leads to post-buckling designs for fuselage panels in critical stability locations. Although buckling of perfect columns (Euler theory) or very thin real structures is essentially an elastic phenomenon (governed by Young's modulus), post-buckling designs show plastic deformation and can benefit from a increased yield strength. This buckling test is shown in FIG. 11. A fuselage skin (20) is fixed to the stiffeners (14, 16) using fixing points (22), for example rivets. The buckling test leads to a deformation of the panel which is manifested by a gap (24) between the stiffener (14, 16) and the skin (22).

It has been noted by the applicant that the shear and compression stability of the fuselage panels working in compression and / or shear can benefit from the high resistance of the stiffeners according to the invention. The use of stiffeners according to the invention as a structural element in an aircraft fuselage panel can improve the shear and compression stability of the fuselage panels, because these stiffeners show higher buckling stability. This effect can be used either to increase the safety margin in constructions where stiffeners are replaced by stiffeners according to the invention, or to reduce the weight of the construction by using stiffeners with reduced sections and thinner fuselage sheets, and / or larger stiffener spacings. It is also possible to obtain an increase in the spacing of the rivets, which decreases the cost of assembling the structure.

An estimate of buckling stability gain can be obtained by applying a general method given in in [Michael CY Niu, Airframe Stress Analysis and Sizing, 2 ^nd edition, chapter 10]. The Applicant noted, using this method, that the increase in the stability of the stiffener according to the invention (with 700 MPa elastic limit in compression and a Young's modulus in compression of 73 GPa) compared with a stiffener in 7150 T77511 (with 538 MPa typical limit value in compression and a Young's modulus in compression of 73 GPa), which is widely used in aircraft according to the state of the art, is greater than or equal to 15% for typical use stiffeners in the form of a "Z".

Table 17 shows parameters of different geometries of stiffeners used for the calculations. Figure 12 compares the predicted buckling stresses for these different geometries from Zl to Z8 (from left to right).

Table 17

Claims

1. Rolled, extruded or forged product of Al-Zn-Mg-Cu alloy, characterized in that it contains (in percent by mass): a) Zn 8.3 - 14.0 Cu 0.3 - 2.0 Mg 0.5 - 4.5

Zr 0.03 - 0.15 Fe + Si <0.25 b) at least one element selected from the group consisting of Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er , Y, Yb, the content of each of the said elements, if selected, being between 0.02 and 0.7%, c) the remaining aluminum and inevitable impurities, and in that it satisfies the conditions d) Mg / Cu> 2.4 and e) (7.9 - 0.4 Zn)> (Cu + Mg)> (6.4 - 0.4 Zn).

2. Product according to claim 1, characterized in that Mg / Cu> 2.8 and preferably> 3.5 and even more preferably> 4.0.

3. Product according to one of claims 1 or 2, characterized in that the mass concentration of the elements Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, does not exceed not 1.5% in total.

4. Product according to any one of claims 1 to 3, characterized in that one selects from the group consisting of Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho,

Er, Y, Yb only titanium.

5. Rolled, extruded or forged product of Al-Zn-Mg-Cu alloy, characterized in that it contains (in percent by mass): a) Zn 9.5 - 14.0 Cu 0.3 - 2.0

Mg 0.5 - 4.5 Fe + Si <0.25 b) at least one element selected from the group consisting of Zr, Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, Cr, Mn, the content of each of the said elements, if selected, being between 0.02 and 0.7%, c ) the remaining aluminum and unavoidable impurities, and in that it satisfies the conditions d) Mg / Cu> 2.4 and e) (7.9 - 0.4 Zn)> (Cu + Mg)> (6.4 - 0.4 Zn).

6. Product according to claim 5, characterized in that the mass concentration of the elements Zr, Se, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, does not exceed 1.5% in total.

7. Product according to one of claims 1 to 4, in which Zn> 9.0%.

8. Product according to claim 7, in which Zn> 9.5%.

9. Product according to claim 7, in which the zinc content is between 9.0% and 11%.

10. Product according to any one of claims 1 to 9, in which Cu> 0.6%.

11. Product according to any one of claims 1 to 10, in which Cu 0.6 - 1.2% and Mg 2.5 - 3.4%.

12. Product according to any one of claims 1 to 10, in which Cu 0.8 - 1.5% and Mg 2.2 - 3.0%.

13. Product according to any one of claims 1 to 12, in which Mg 0.5% - 3.6%.

14. Product according to any one of claims 1 to 13, in which Mg> 1.95 + 0.5 (Cu - 2.3) + 0.16 (Zn - 6) + 1.9 (Si - 0.04).

15. Product according to any one of claims 1 to 14, in which the following maximum concentrations are not exceeded:

Cr 0.40 Mn 0.60 Se 0.50 Zr 0.15 Hf 0.60 Ti 0.15

Ce, Nd, La and Eu 0.35 each and preferably 0.30 each Gd 0.35 Tb 0.35 Ho 0.40 Dy 0.40 Er 0.40 Yb 0.40 Y 0.20

16. Product according to any one of claims 1 to 15, characterized in that it additionally contains at least one element selected from the group consisting of Ag, Sn, Cd, Ge, In, the content of each of these elements, s 'it is selected, being between 0.02% and 0.15% and preferably between 0.05% and 0.10%.

17. A spun product according to any one of claims 1 to 16, characterized in that it has in the state T6511, measured on test pieces taken from a flat area, a) a bending angle, measured at 130 ° C by a 3-point bending test according to DIN 50 111 (section 3.1) on a 1.6 mm thick sample, and expressed as an average value calculated from individual measurements carried out on samples taken from different locations distributed over the length of the profile, at least 34 °, and b) an elastic limit R _p o, 2 of at least 720 MPa, and preferably a folding angle of at least 35 ° and an elastic limit of at least 750 MPa.

18. A spun product according to any one of claims 1 to 16, characterized in that it has in the T76511 state, measured on test pieces taken from a flat area, a) a bending angle, measured at 130 ° C by a test 3 point bending according to DIN

50 111 (section 3.1) on a sample 1.6 mm thick, and expressed as an average value calculated from individual measurements carried out on samples taken from different locations distributed over the length of the profile, at least 37 ° and preferably at least 40 °, and b) an elastic limit R _p o, 2 of at least 670 MPa.

19. A spun product according to claim 18, characterized in that the corrosion resistance, determined according to the EXCO test (ASTM standard G34) state T6511 on non-machined samples, is at least level EB.

20. Element of aircraft structure, produced in a product according to any one of claims 1 to 19.

21. Structural element according to claim 20, characterized in that said element is a fuselage stiffener.

22. Structural element according to claim 20, characterized in that said element is a seat rail.

23. A seat rail according to claim 22, characterized in that its resistance to rupture in the T76511 state in the seat attachment zone is at least 670 MPa and preferably at least 680 MPa.

24. Seat rail according to claim 22 or 23, characterized in that its elastic limit in the state T76511 in the seat fixing area is at least 640 MPa and preferably at least 660 MPa.

25. structural element according to claim 20, characterized in that said element is a floor profile.

26. An aircraft comprising a fuselage assembled from a plurality of stiffeners and a plurality of sheets, characterized in that at least part of said stiffeners are structural elements according to claim 20.

27. Fuselage structure assembled from a plurality of stiffeners according to claim 21 and a plurality of sheets, characterized in that for a spacing of said stiffeners of 200 mm and a stiffening rate (section of the stiffener / total section ) 0.25, the stress intensity factor for a crack of a length of two stiffeners with the broken central stiffener is reduced by at least 5% compared to stiffeners made of 2024 T3 alloy.

28. Fuselage structure assembled from a plurality of stiffeners according to claim 21 and a plurality of sheets, characterized in that the buckling stability of said stiffeners is improved by at least 15% compared to an identical structure incorporating Z-shaped stiffeners of the same alloy geometry

7150 T77511.