AU2003260003A1 - Al-zn-mg-cu alloys welded products with high mechanical properties, and aircraft structural elements - Google Patents

Abstract

The present invention further relates to 7xxx alloys and products produced therewith that can be flat rolled, extruded or forged, as well as associated methods. Al-Zn-Mg-Cu alloys of the present invention preferably comprise (in mass percentage): 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 Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y and Yb, where the content of each of the elements, if selected, is between 0.02 and 0.7%, c) remainder aluminum and inevitable impurities. The present invention further is directed to products wherein Mg/Cu>2.4 and (7.9-0.4 Zn)>(Cu+Mg)>(6.4-0.4 Zn). The disclosed products can be used for example, as structural members in aeronautical construction, especially as stiffeners capable for use in fuselages of civilian and other aircrafts as well as in related applications.

Description

4JfS 1- _ C'est votre traduction! InO inwatique - Web Xeionau o que 0 Autombile Technique. -ainue d nuti stion Mj Idical -Phar.nacctiqu o I ,-- ri r Coin mercial -Marketing 0 VERIFICATION OF TRANSLATION I, Mrs. GONCALVEZ, of A.R.T. International - 26, rue Carnot 95410 Groslay, France declare as follows: 1. That I am well acquainted with both the English and French languages, and 2. That the attached document is a true and correct translation to the best of my knowledge and belief of: PCT/FR 03/01063 filed on April 4th 2003 Dated this 24th day of September 2004 (no witness required) 7- , ,. T1 S.A. au capital de 40 000 E - R.C S. B 392 83(0 337 ABSTRACT OF THE DISCLOSURE WROUGHT AL-ZN-MG-CU ALLOY PRODUCTS WITH VERY HIGH MECHANICAL CHARACTERISTICS AND STRUCTURAL MEMBERS FOR AERONAUTICAL CONSTRUCTION MADE THEREOF The invention relates to a product which is flat rolled, extruded or forged in Al-Zn-Mg-Cu alloy, characterised in that it contains (in mass percentage): a) Zn 8.3 - 14.0 Cu 0.3 - 4.0 Mg 0.5 - 4.5 5 Zr 0.03 - 0.15 Fe + Si < 0.25 b) at least one element selected from the group made up of Sc, 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%, 10 c) remainder aluminium and inevitable impurities, and in that it satisfies the conditions d) Mg/Cu > 2.4 and e) (7.7 - 0.4 Zn) > (Cu + Mg) > (6.4 - 0.4 Zn) The products according to the invention can be 15 used for manufacturing stiffeners of the fuselage of civilian aircraft. Figure 1 WROUGHT AL-ZN-MG-CU ALLOY PRODUCTS WITH VERY HIGH MECHANICAL CHARACTERISTICS AND STRUCTURAL MEMBERS FOR AERONAUTICAL CONSTRUCTION MADE THEREOF DESCRIPTION Technical field of the invention This invention relates to wrought products made of Al-Zn-Mg-Cu type alloys with very high mechanical characteristics, with a Zn content greater than 8.3%, as well as to structural members for aircraft 5 manufactured from these products. Description of the state of the art The alloys of type Al-Zn-Mg-Cu (belonging to the alloys of the 7xxx series) are currently used in 10 aeronautical construction, and particularly in the construction of civilian aircraft wings. For the extrados of the wings a skin made of plates with a high content of alloys such as 7150, 7055, 7449 is used, and optionally stiffeners made from profiles in alloys 15 7150, 7055, 7349 or 7449. 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 for example alloys 7075 and 7175 (zinc content 20 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 present a high tensile yield strength (TYS), as well as good fracture toughness and good resistance to 2 stress corrosion and to exfoliation corrosion. More recently, it has appeared that for certain applications the use of an alloy with a higher zinc content can have advantages, since this allows the tensile yield 5 strength to be increased further. Alloys 7349 and 7449 contain between 7.5 and 8.7% of zinc. Wrought alloys still higher in zinc have been described in the literature, but do not seem to be used in aeronautical construction. 10. The article "Microstructure and properties of a new super-high-strength Al-Zn-Mg-Cu alloy C912" by Y.L. Wu et al., published in the journal Materials & Design, vol 18, p. 211-215 (1998) describes an alloy of the composition Zn 8.7%, Mg 2.6%, Cu 2.5%, Si and Fe each < 15 0.05%, which is considered for the manufacture of structural members for wings and fuselage. US patent 5 560 789 (Pechiney Recherche) discloses an alloy comprising Zn 10.7%, Mg 2.84%, Cu 0.92% which is transformed by extrusion. The alloying elements of 20 this alloy with very high zinc, magnesium and copper contents are difficult to put into solid solution because the temperature of solution heat treatment is limited by the melting temperature of the phases with the lowest incipient melting point: this product has a 25 high mechanical strength, but a very low elongation at fracture due to the presence of coarse precipitates. Thus, such a product has a low formability. US patent 5 221 377 (Aluminum Company of America) discloses several alloys of type Al-Zn-Mg-Cu with a 30 zinc content of up to 11.4% and with a rather high copper content. These alloys are difficult to cast and 3 the alloying elements are difficult to put into solid solution, which favors the undesirable presence of coarse precipitates. Moreover, it has been proposed to utilise Al-Zn 5 Mg-Cu alloys with a high zinc content to manufacture hollow bodies intended to resist high pressures, such as for example compressed gas cylinders. European patent application EP 020 282 Al (Socit Metallurgique de Gerzat) discloses alloys with a zinc content of 10 between 7.6% and 9.5%. European patent application EP 081 441 Al (SociEtu Metallurgique de Gerzat) discloses a process for obtaining such cylinders. European patent application EP 257 167 Al (Soci6t Metallurgique de Gerzat) states that none of the known 15 alloys of type Al-Zn-Mg-Cu can safely and reproducibly satisfy the strict technical demands imposed by this specific application; it proposes moving towards a lower zinc content, namely between 6.25% and 8.0%. The teaching of these patents is specific to the 20 problem of compressed gas cylinders, particularly concerning maximisation of the bursting pressure of these cylinders, and cannot be transferred to other wrought products. In general, in alloys of type Al-Zn-Mg-Cu, a high 25 zinc content, but also a high Mg and Cu content are required in order to obtain good static mechanical characteristics (tensile yield strength (TYS), ultimate tensile strength,), but this is possible only if these elements can be put into solid solution. But it is also 30 well known (see for example US 5 221 377) that when in an alloy of the 7 xxx series the zinc content is 4 increased beyond around 7 to 8%, then problems associated with insufficient resistance to exfoliation corrosion and stress corrosion will arise. More generally, it is known that the most charged Al-Zn-Mg 5 Cu alloys are likely to cause corrosion problems. These problems are generally solved by means of particular thermal or thermo-mechanical treatments, especially by pushing the aging treatment beyond the peak, for example during T7 type treatment. But these treatments 10 can then cause a drop in the static mechanical characteristics. In other words, for a minimum level of resistance to corrosion envisaged, when optimising an alloy of type Al-Zn-Mg-Cu one must find a compromise between static mechanical characteristics (TYS Rp0.

2 , 15 UTS Rm, elongation at fracture A) and characteristics related to damage tolerance (fracture toughness, crack propagation rate etc.). According to the minimal level of resistance to corrosion envisaged, either a temper close to the aged peak is utilised (T6 tempers), which 20 generally offers a toughness - Rp0.

2 compromise favouring static mechanical characteristics, or ageing is increased beyond the peak (T7 tempers), by seeking a compromise favouring fracture toughness. Whichever approach is used, the manufacture and 25 use of such products cause two problems: on the one hand, these alloys with a high zinc and magnesium contents are difficult to cast and to transform, especially by extrusion, rolling or forging. For example, the maximum force that can be supplied by an 30 extrusion press can be a limiting factor. In particular, among the alloys of the 7xxx series, alloys 5 7349 and 7449 require very high extrusion forces. On the other hand, for certain applications, the formability of said rolled and extruded products is critical. This is true especially for fuselage 5 stiffeners. Therefore, the search for alloys having a mechanical strength still higher than 7349 and 7449 alloys is likely to lead to an alloy difficult to cast and to transform, with a low formability. 10 Summary of the invention A problem which the present invention attempts to solve is therefore to manufacture novel wrought products in alloy of Al-Zn-Mg-Cu type with a high zinc content, greater than 8.3%, and especially extruded 15 products, said products presenting a very high ultimate tensile strength, a very high tensile yield strength as well as adequate resistance to corrosion and a high formability, and which can be manufactured industrially under conditions of highest reliability compatible with 20 the severe requirements of the aeronautical industry. Subject matters of the invention The present inventors have found that the problem can be solved by finely adjusting the concentration of 25 the alloying elements Zn, Cu and Mg and of certain impurities (particularly Fe and Si), and by optionally adding other elements. One first subject matter of the present invention is directed to a product which is rolled, extruded or 30 forged in Al-Zn-Mg-Cu alloy, characterised in that it contains (in mass percentage): 6 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 5 made up of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, the content of each of said elements, if selected, being between 0.02 and 0.7%, c) remainder aluminium and inevitable impurities, and in that it satisfies the conditions 10 d) Mg/Cu > 2.4 and e) (7.9 - 0.4 Zn) > (Cu + Mg) > (6.4 - 0.4 Zn). Another subject matter of this invention is a rolled, extruded or forged product made of an Al-Zn-Mg Cu alloy, characterised in that it contains (in mass 15 percentage): 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 20 made up of Zr, Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, Cr, Mn, the content of each of said elements, if selected, being between 0.02 and 0.7%, c) remainder aluminium and inevitable impurities, and in that it satisfies the conditions 25 d) Mg/Cu > 2.4 and e) (7.9 - 0.4 Zn) > (Cu + Mg) > (6.4 - 0.4 Zn). In yet another subject matter of the present invention, a structural member for aeronautical construction is provided, incorporating at least one of 30 said products, and particularly a structural member utilised in the construction a fuselage of civilian 7 aircraft, such as a fuselage stiffener. Brief description of the figures Figure 1 shows the section of profile TI. 5 Figure 2 shows the section of profile T2. 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 are 10 approximate values expressed in millimetres. In figures 1, 2, 3 and 4, letter a designates the foot section (bottom flange), and letter b the top section (top flange) of the profile. Figure 6 diagrammatically illustrates the zone of 15 a fuselage stiffener which has been formed by joggling. The reference letters are as follows: a Joggling depth b Joggling width c Upper foot: important plane d deformations Lower foot: important plane deformations Figure 7 diagrammatically shows the location on profile T1 from which the sample is taken for the 3 point bending test. 20 Figure 8 diagrammatically shows the definition of the bending angle. Figure 9 diagrammatically shows geometric parameters that are important for the 3 -point bending test. 25 Figure 10 diagrammatically shows a crack with the 8 same length as two stiffeners, with the central stiffener broken. Figure 11 diagrammatically shows the buckling test. Figure (b) corresponds to a rotation A-A of 900. 5 Figure 12 compares buckling stresses for different Z-shaped stiffeners according to the invention (grey bars) and according to prior art (white bars), for the same geometry. 10 Detailed Description of the invention Unless indicated otherwise, the chemical compositions are given as mass percentages. Therefore, in a mathematical formula, "0.4 Zn" means "0.4 times the zinc content, expressed in mass percentage"; this 15 applies mutatis mutandis to the other chemical elements. The alloy designations follow the rules of The Aluminum Association. The metallurgical tempers are defined in the European standard EN 515. Unless indicated otherwise, the static mechanical 20 characteristics, i.e. the ultimate tensile strength Rm, the tensile yield strength Rp0.

2 and the elongation at fracture A, are determined by a tensile test according to the standard EN 10002-1. The term "extruded product" includes so-called "drawn" products obtained by 25 extrusion, followed by drawing. During preparatory studies, the present inventors arrived at the conclusion that a novel material exhibiting a significantly improved compromise between mechanical strength and formability should have in any 30 case a sufficiently high zinc content, typically above 8.3%, and preferably above 9.0%. However, this 9 condition is not sufficient. According to the present invention, the inventors have found a very specific domain of composition that can be used to produce wrought products, and especially 5 extruded products, which have at the same time high static mechanical properties, a sufficient resistance to corrosion, and a good formability. By this invention, the inventors have been able to develop extruded products which can be used advantageously as 10 fuselage stiffeners in civilian aircraft. For this use, damage tolerance is not a limiting factor; it is therefore admissible to optimise the UTS and TYS while losing some damage tolerance, if the corrosion resistance is not affected. However, increasing the UTS 15 and TYS to their maximum values so that the weight of the aircraft can be reduced, usually leads to a decrease in formability. It is known to one skilled in the art that fuselage stiffeners are subject to complex and very peculiar forming operations. When developing 20 an alloy for fuselage stiffeners which has a higher mechanical strength, it is therefore necessary that the formability is at least as good as, or preferably better than, in conventional alloys. According to the present invention, this problem 25 is solved by fine adjustment of the contents of the elements of the alloys and certain impurities, and by adding a controlled concentration of certain other elements to the alloy composition. The present invention applies to Al-Zn-Mg-Cu 30 alloys containing: Zn 8.3 - 14.0 Cu 0.3 - 2.0 Mg 0.5 - 4.5 10 as well as certain other elements specified herein below, and the rest being aluminium with its inevitable impurities. The alloys according to the present invention must 5 contain at least 0.5% magnesium, since it is not possible to obtain satisfactory static mechanical characteristics with a lower magnesium content. In alloys with a zinc content of less than 8.3%, the inventors have not found any improvement compared with 10 conventional alloys. Preferably, the zinc content is higher than 9.0%, and still more preferably higher than 9.5%. However, certain arithmetic relations between certain elements have to be respected, as will be explained below. In another advantageous embodiment of 15 the invention, the zinc content is between 9.0 and 11.0%. In any case, the zinc content should not exceed a value of about 14%, because the results are unsatisfactory beyond this value, regardless of the magnesium and copper contents. 20 The addition of at least 0.3% of copper improves resistance to corrosion. A minimum copper content of 0.6% is preferred. But to ensure satisfactory solution heat treatment, the Cu content should not exceed about 2%, and the Mg content should not exceed about 4.5% ; a 25 maximum content of 3.6% is preferred for magnesium. In a preferred embodiment, the copper content is between 0.6% and 1.2%, while the magnesium content is between 2.5% and 3.4%. In another preferred embodiment, the copper content is between 0.8% and 1.2%, while the 30 magnesium content is between 2.2% and 3.0%. As will be explained below, the ratio between the magnesium and 11 copper contents must comply with certain criteria. The present inventors have found that to solve the problem in question, in an alloy of type Al-Zn-Mg-Cu, several additional technical features must be 5 considered. First of all, the alloy must be sufficiently charged with alloying elements likely to precipitate during maturation or ageing treatment, in order to be capable of presenting advantageous static mechanical 10 characteristics. To do this, according to findings by the present inventors, in addition to the minimum and maximum limits for the zinc, magnesium and copper contents indicated hereinabove, the contents of these alloy additions must satisfy the condition Mg + Cu > 15 6.4 - 0.4 Zn. To reinforce this effect, a sufficient content of so-called anti-recrystallising elements must be added. More precisely, for alloys with more than 9.5% zinc, at least one element selected from the group comprising 20 the elements Zr, Sc, Hf, La, Ti, Y, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Yb, Cr, Mn must be added, with for each element present, a concentration of between 0.02 and 0.7%. It is preferred that the total concentration of the elements of said group does not exceed 1.5%. 25 These anti-recrystallising elements, in the form of fine precipitates formed during thermal or thermomechanical treatment, block recrystallisation. However, we have found that when the alloy is highly charged with zinc (Zn > 9.5%) excessive precipitation 30 has to be avoided when the wrought product is being quenched. A compromise then has to be found for the 12 anti-recrystallising elements content. According to the invention, for alloys with a zinc content of between 8.3% and 9.5%, zirconium with a content of between 0.03% and 0.15% must be added, and 5 at least one element selected from the group comprising the elements Sc, 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 a preferred embodiment, titanium is selected, alone or together 10 with one or more other elements selected from said group. The inventors have noted that for said anti recrystallising elements it is advantageous, regardless of the zinc content, not to exceed the following 15 maximum contents : Cr 0.40 ; Mn 0.60 ; Sc 0.50 ; Zr 0.15 ; Hf 0.60 ; Ti 0.15 ; Ce 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 ; La 0.35 and preferably 0.30. It is 20 preferred that the total concentration of these elements does not exceed 1.5%. The inventors have found that in order to optimise the UTS and TYS, a ratio Mg / Cu > 2.4 should be adopted, and preferably a ratio of at least 2.8, and 25 even more preferably of 3.5 or even 4.0. Another technical feature is associated with the need to be able to manufacture wrought products industrially under reliability conditions compatible with the severe requirements of the aeronautical 30 industry, as well as under satisfactory economic conditions. So it is necessary to choose a chemical 13 composition which minimises the appearance of hot cracks or splits during solidification of the plates or billets, said hot cracks or splits being crippling defaults leading to said plates or billets being 5 discarded. The inventors have noted during numerous tests that this appearance of hot cracks or splits was much more probable when the 7xxx alloys finished solidifying below 470 0 C. To significantly reduce the probability of hot cracks or splits occurring during 10 casting to an acceptable industrial 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). Within the scope of the present invention this 15 criterion is called the "castability criterion". The alloys produced according to this variant of the invention complete their solidification at a temperature of between 473 oC and 478 'C, and thus allow to reach an industrial reliability of metal 20 working processes (that is, a constant and excellent quality of the cast plates or billets) compatible with the severe requirements of the aeronautical industry. Another technical feature of the invention is associated with the need to minimise as much as 25 possible the quantity of insoluble precipitates (that are typically type S, M or T Al-Zn-Mg-Cu ternary or quaternary phases) following homogenisation and solution heat treatments, because this decreases the fracture toughness, the elongation at rupture, and 30 especially the formability. To do this, a Mg, Cu and Zn content such as Mg + Cu < 7.9 - 0.4 Zn is selected. The 14 inventors have found that there is no disadvantage to select a composition close to the borderline represented by the relation Mg + Cu < 7.9 - 0.4 Zn, but that beyond said borderline, the excellent aptitude to 5 deep formability by joggling (which is one of the main advantages of the products according to this invention) decreases rapidly. And finally, incorporating a small quantity, of between 0.02 and 0.15% per element, of one or more 10 elements chosen from the group made up of Sn, Cd, Ag, Ge and In enables the response of the alloy to the ageing treatment to be improved, and has beneficial effects on mechanical resistance and on resistance to corrosion of the product. A concentration between 0.05% 15 and 0.10% is preferred. Among these elements, silver is preferred. For profiles, adding one or more anti recristallising elements such as scandium is particularly advantageous; such an effect is also seen in thick plates. In profiles, an increase of mechanical 20 strength is observed which is higher when the width or thickness of the profile is low ; this so-called "press effect" is known to one skilled in the art. The inventors have found that when the added anti recrystallising element is scandium, a concentration 25 between 0.02% and 0.50% is preferred. The products according to the present invention are especially extruded products. They can be used advantageously to produce structural members suitable for use in aeronautical construction. A preferred 30 application of the products according to the present invention is their use as a structural member of 15 fuselage of civilian aircraft. These structural members, in particular stiffeners, are principally dimensioned for mechanical strength. Damage tolerance is not normally a property that affects the 5 dimensioning of such structural members, provided that it is equal to an acceptable minimum value; therefore, if needed and up to a certain point, the mechanical strength can be optimised while accepting a certain loss of damage tolerance, without decreasing the 10 usefulness of the final product. However, corrosion resistance must always be at an acceptable level. The increase of mechanical strength of said fuselage stiffeners allows, at the constructer's discretion, to decrease their weight, or to obtain, at constant 15 weight, a stronger fuselage structure. This may allow, by increasing the distance between two adjacent stiffeners (within the limits of bending of the fuselage sheets), to decrease the number of fuselage stiffeners, which leads to a decrease of the number of 20 assembly points between the stiffener and the wing skin. This can be very advantageous, because the assembly points, such as bolts or rivets, make a very significant contribution to the overall manufacturing cost of such structures. Thus, a particularly 25 advantageous use of the product according to the present invention is the use as structural member in aeronautical constructions, and in particular in the construction of aircraft comprising a fuselage assembled from a plurality of stiffeners and a 30 plurality of sheets, wherein at least part of said stiffeners are structural members according to the 16 present invention. Such an aircraft will either have a structure of less weight, which is at least as strong, or a stronger structure, which will not be heavier than that of aircraft according to the state of the art. 5 It is not only advantageous to minimise the number of assembly points between structural members of different type (such as fuselage stiffeners and fuselage skin), but also between structural members of the same type, and especially between two stiffeners. 10 In order to achieve this goal, it is advantageous to utilise sheets or extruded products with a relevant size parameter as large as possible; in the case of extruded profiles, this relevant size is essentially the length. The manufacture of very long profiles in 15 Al-Zn-Mg-Cu alloys with very high content of alloying elements requires an excellent process control during casting, extrusion and thermal treatments, and may require the adjustment of the chemical composition according to the invention. In particular, the present 20 inventors have found that the product according to the invention can be obtained by using a lower extrusion force than is possible with known products having a comparable zinc content, which makes it possible to manufacture longer profiles. 25 It is known that aviation authorities require a designed structure to resist limiting loads despite serious damage; the chosen damage is a crack with the same length as two stiffeners with the central stiffener broken (see figure 10). The inventors have 30 noted that the residual strength of fuselage panels stressed in tension can strengthen high strength 17 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 5 in the skin, which preventively avoids an unstable failure. The residual strength of the cracked panel is thus improved. This effect can be used either to improve the safety factor in constructions in which stiffeners are replaced by stiffeners according to the 10 invention, or to lower the weight of the construction using stiffeners with smaller sections and thinner fuselage sheets, and / or larger stiffener spacings. Aviation authorities require that the structure should be designed to resist an ultimate load for 3 15 seconds without excessive deformation. Nevertheless, plastic deformation is allowed. This results in post buckling designs for fuselage panels in locations critical for stability. Although the buckling of perfect columns (Euler's theory) or very thin real 20 structures is essentially an elastic phenomenon (governed by Young's modulus), post-buckling designs include plastic deformation and can benefit from an increase in the tensile yield strength. Figure 11 shows this buckling test. 25 The inventors have noted that the shear and compression stability of fuselage panels stressed in compression and / or shear can benefit from the high strength of stiffeners according to the invention. The use of stiffeners according to the invention as 30 structural elements in an aircraft fuselage panel can improve the shear and compression stability of fuselage 18 panels, because these stiffeners are more stable in buckling. This effect can be used either to increase the safety factor in constructions in which stiffeners are replaced by stiffeners according to the invention, 5 or to lower the weight of the construction by using stiffeners with smaller sections and thinner fuselage sheets and / or larger stiffener spacings. Rivet spacings can also be reduced, which reduces the assembly cost of the structure. 10 Table 17 shows parameters for different stiffener geometries used for the calculations. Figure 12 compares predicted buckling stresses for these different geometries from Zl to Z8 (from left to right). 15 Yet another problem which arises when using the said products as fuselage stiffeners is their formability. One forming technique used during the industrial manufacture of fuselage stiffeners from profiles is 20 joggling. This includes the introduction of a step localised over a zone of a few millimetres (see figure 6). This can be achieved, for profiles according to the present invention, either at elevated temperature (preferably at 130 oC), or at room temperature. When 25 joggling is performed at room temperature, it is preferred to proceed by a solution heat treatment of the profile delivered in the instable W temper, followed by quenching. Then forming is done by joggling. Joggling at room temperature does not allow a 30 forming as deep as joggling at elevated temperature, but when applicable, it is often more practical.

19 Joggling as an industrial process does not lend itself to the characterisation of materials under development. It is however known that defects appearing during joggling are related to maximum plane 5 deformation which can be supported by the material. Thus, it is possible to evaluate the aptitude of a material to forming by joggling by using the 3 point bending test. According to DIN standard 50111 (September 1987, see especially section 3.1), the 10 specimen width to thickness ratio must be sufficient with respect to its thickness to maintain conditions of plane deformation, in the centre of the test piece. In the present invention, in order to evaluate the formability at 130 oC (i.e. warm formability of the 15 product in its final temper), the flat test piece is deformed in a furnace at 130 oC until a drop of the applied force is detected ; this indicates that a crack has initiated. The temperature must be precisely controlled and maintained at 130 0 C during this test. 20 Since deformation takes place at elevated temperature, the rate of deformation is a parameter which influences the result. In the present case, this rate has been fixed at 50 mm/mn. The higher the bending angle (see definition in figure 8), the better the aptitude of the 25 material to forming by joggling. For mechanical reasons, it is important that all test pieces to be compared have the same thickness. When comparing two test pieces of different thicknesses, the face which will be under compressive stress has to be machined 30 down to the thickness of the thinnest test piece. In the case of a profile, the flat test piece is cut at a 20 representative location, as shown in figure 5 for the profile TI. The 3 point bending tests at 130 oC are applied to test pieces cut from products in T6x or T7x temper. It 5 is also possible to characterise formability in the as quenched condition W, if the period of time between quenching and execution of the three point bending test is controlled. In the case of extruded products, the bending angle at 130 0C is expressed as the average 10 value calculated from individual measurements on test pieces taken from different locations over the length of the profile. A product according to the present invention which is particularly preferred is an extruded product which 15 exhibits in the T6511 temper a bending angle, determined at 130 oC by the 3 point bending test according to DIN 50 111 (section 3.1) on a sample of 1.6 mm thickness cut from a plane area, of at least 340, and a TYS of at least 720 MPa, and preferably a 20 bending angle of at least 350 and a TYS of at least 750 MPa. For a thickness up to 60 mm, the static mechanical properties (Rp0.

2 , Rm, and A) do not depend much on the thickness of the section. Another preferred product is an extruded product 25 which exhibits in the T76511 temper a bending angle, determined at 130 oC by the 3 point bending test according to DIN 50 111 (section 3.1) on a sample of 1.6 mm thickness cut from a plane area, of at least 36 0 , and a TYS of at least 660 MPa and preferably of at 30 least 670 MPa. This product can be used for applications in which a corrosion resistance of a 21 rating of at least EB (EXCO test according to ASTM G34) is required for non-machined specimens. Both preferred products can be used advantageously as fuselage stiffeners in civilian aircraft. 5 An mentioned above, the present inventors have surprisingly found that compared to prior art products, and including prior art products with a comparable zinc content, the products according to the present invention exhibit a high formability when hot. On the 10 other hand, cold formability in the unstable W temper after solution heat treatment and quenching is slightly less good. For the manufacture of structural members of aircraft, such as fuselage stiffeners, we therefore prefer the warm forming process, if the said forming is 15 deep. The products according to the present invention can also be used as floorbeams for aircraft and in the form of extruded profiles, as seat tracks. In civilian aircraft, seat tracks are profiles, generally very long 20 and normally parallel to the length of the cabin, on which the seats are mounted. According to the present invention, seat tracks in T76511 temper can be obtained which exhibit a UTS in the area where the seats are fixed (i.e. the top of a "I" shaped profile) of 670 MPa 25 or more, and even of 680 MPa or more, and a TYS of 640 MPa or more, and even of 660 MPa or more. Seat tracks of commercial aircraft have to be resistant to corrosion by corrosive liquid foodstuff under high mechanical stress. Indeed, seat tracks according to the 30 present invention exhibit a good stress corrosion resistance as determined according to standard ASTM 22 G47. The use of structural members according to the present invention for aircraft construction leads to significant weight savings, which allows to increase 5 the load capacity of said aircraft, or to decrease fuel consumption. The invention will be better understood after reading the examples that are in now way limitative 10 EXAMPLES Example 1: Several Al-Zn-Mg-Cu alloys were prepared by semi continuous casting of rolling ingots, and were then subjected to a conventional transformation procedure, 15 comprising a homogenisation step, the parameters of which have been determined using the teaching of US patent 5,560,789, being followed by hot rolling, solution heat treatment followed by quenching and stress relieving operations, and finally an aging 20 treatment in order to obtain a product with temper T651. The result of this was plates in the T651 temper having a thickness of 20 mm. The compositions of the plates according to this example are specified in Table 1. 25 Table 1 Alloy Zn Mg Cu Fe Si Zr Ti Mn Sc Mg/Cu A 8,40 2,11 1,83 0,09 0,06 0,11 0,01 0 0 1,15 7 23 B 10.2 3.2 0.71 0.08 0.03 0.11 0.01 0 0 4.57 7 7 C 10.0 2.69 0.95 0.08 0.03 0.11 0.01 0 0 2.83 8 4 D 9.97 2.14 1.32 0.09 0.03 0.11 0.01 0 0 1.62 7 The static mechanical characteristics were determined by a tensile test according to standard EN 10002-1. The fracture toughness K 1 c was determined 5 according to standard ASTM E399. The results are specified in Table 2: Table 2 Alloy Tensile test L Tensile test TL Fracture direction direction toughness L T Rp 0

,

2 Rm A Rp 0 , 2 Rm A Kic MPa MPa % MPa MPa % MPaqm A 627 665 14.7 566 623 13.6 31.9 B 716 726.5 6.5 640 696 5.2 21.1 C 700 717 9.2 629 676 8.1 21 D 665 685 12.2 608 649 11 26.8 10 It appears that plate C according to the invention presents a good compromise between mechanical strength and elongation. Compared with plate D, which is outside the scope of the present invention, the mechanical 15 strength is significantly improved. Compared with plate A made of a 7449 alloy according to prior art, plate C exhibits a mechanical strength that is very significantly improved. The fact that fracture toughness is less good in plate C than in plate B 20 limits the possibilities of application of plate C to those applications for which fracture toughness is not 24 taken into account when dimensioning the structural members, but which require both a high mechanical strength and a good formability. Compared with plate B, which is outside the scope of the present invention, 5 the elongation at fracture of plate C is significantly improved. Moreover, in order to obtain the properties given in Table 2, plate B needs to be submitted to a rather long solution heat treatment, which does not lend 10 itself to the requirements of industrial production. And yet, coarse phases have been found in the product which have an adverse effect on the homogeneity of mechanical properties, both within the same production lot and within the same individual product (plate or 15 profile); this precludes the use of product B as structural member in aircraft. Example 2 : Several rolling ingots whose alloy composition is 20 specified in Table 3 were cast, with a silicon content approximately the same for all alloys, about 0.04%. Alloys Gl, G2, G3 and G4 and B are outside the present invention. The compositions of alloys B and D (outside the scope of the present invention) and of 25 alloy C (according to the invention) are described in example 1. During testing all these alloys exhibited satisfactory castability, that is, no splits or cracks were observed during casting tests performed on an industrial scale. 30 Alloys G5, G6, G7, G8 are outside the present invention, and alloy G9 is an alloy 7060 as per the 25 prior art ; these alloys exhibited cracks during casting tests. The difficulties showing up during casting of these alloys do not necessarily render the wrought 5 products from these plates unsuitable for use, but they are the cause of extra cost because their implementation (that is, the quantity of marketable metal relative to the quantity of charged metal, a parameter directly associated with the quantity of 10 discarded plates) will be greater than for the alloys corresponding to the preferred domain of the invention. In addition, the propensity of these alloys to form splits during their solidification makes it very difficult to render the casting process reliable within 15 the scope of a quality assurance program based on statistical process control. It is noted that all the 7xxx alloys having a very pronounced propensity to form splits or cracks in casting have a magnesium content lower than the 20 critical magnesium content ; this critical value was obtained by calculating the Mg limit value defined by the castability criterion. Table 3 25 Allo Zn Mg Cu Observed Critical Mg > y [%] [%] [%] crackabil Mg Critical ity content Mg G1 7.5 3 3 low 2.54 yes G2 8.5 3 2.3 low 2.35 yes G3 7.5 3 1.6 low 1.84 yes G4 6.5 3 2.3 low 2.03 yes B 10.27 3.2 0.71 low 1.82 yes C 10.08 2.69 0.95 low 1.91 yes D 9.97 2.14 1.32 low 2.08 yes 26 G5 8.5 2.3 3 high 2.7 no G6 6.5 2.3 3 high 2.38 no G7 8.5 1.6 2.3 high 2.35 no G8 7.5 1.6 1.6 high 1.84 no G9 7 1.65 2.1 high 2.01 no Example 3 : Extrusion ingots have been cast from alloys whose composition is summarised in Table 4. Homogenisation was carried out as follows : 5 Ingots Q1 and Q2 : 4 h at 465 oC + 20 h at 476 oC Ingots Q3 and Q4 : 4 h at 465 0C + 20 h at 471 oC Ingots P1 through P3 : 20 h at 471 oC. 10 M, T and S phases were completely dissolved during the homogenisation treatment; this was checked by differential enthalpic analysis (see US patent 5,560,789). Ingot diameter was 200 mm for ingots P3 and Q1 15 through Q4, and 155 mm for ingots P1 and P2. Table 4 Ingot Zn Mg Cu Cr Mn Si Fe Zr Ti Mg/Cu P1 8.10 2.48 1.65 0.14 0.17 0.01 0.08 0.15 0.03 1.50 P2 8.45 2.60 1.76 0.18 0.18 0.05 0.14 0.12 0.02 1.48 P3 8.39 2.55 1.71 0.18 0.16 0.04 0.15 0.11 0.02 1.49 Q1 10.2 3.10 0.68 0.17 0.17 0.07 0.08 0.13 0.04 4.56 0 Q2 10.2 2.84 0.95 0.18 0.17 0.06 0.11 0.13 0.03 2.99 0 27 Q3 9.98 2.10 1.24 0.18 0.17 0.06 0.14 0.12 0.03 1.69 Q4 10.0 2.15 1.25 0.18 0.17 0.07 0.14 0.12 0.03 1.72 0 R1 10.1 2.97 0.66 0.17 0.16 0.07 0.13 0.11 0.02 4.5 8 R2 10.1 3.12 0.70 0.17 0.16 0.07 0.13 0.11 0.02 4.46 6 From these homogenised and scalped ingots, five types of profiles TI, T2, T3, T4 and T5 were extruded. Their sections are represented on figures 1, 2, 3, 4 5 and 5. The temperature of the container and of the die was above 4000C , and the extrusion speed was below 0.50 m/min. Maximum extrusion forces are summarised in Table 5. It can be seen that surprisingly the alloys 10 according to the present invention do not require a higher extrusion force, and that the extrusion force surprisingly even decreases for certain types of profiles with increasing magnesium content. 15 Table 5 Prof Extrusio Extrusio Extrusio Extrusio Extrusio Extrusi ile n force n force n force n force n force on [bars] [bars] [bars] [bars] [bars] ratio for for for for for ingot P1 ingot Q1 ingot Q2 ingot Q3 ingot Q4 T1 179 175 170 164 164 58 T2 151 145 142 137 139 24 T3 203 208 200 193 195 13 Profiles Q1 through Q4 were solution heat treated at 471 oC, while profiles P1 through P3 were solution 20 heat treated at 472 oC (profiles TI, T2 and T3).

28 Profiles R1 and R2 were solution heat treated under comparable conditions. All profiles were water quenched and then stretched with a permanent set between 1.5% and 2%. Products in tempers T6511 or T76511 were 5 obtained. Their mechanical properties are summarised in Table 6 for specimens of three different thickness values in temper T6511, cut from a flat area of the profile. This temper has been obtained by artificial 10 aging under the following conditions: Alloys Q1 and Q2 : 18 h at 120 oC Alloys P1 through P3, Q3 and Q4 : 36 h at 120 oC. Table 6 15 Alloy Rm [MPa] Rp 0

,

2 [MPa] A [% ] profil T1 T3 T2 T1 T3 T2 T1 T3 T2 e Q1 755 753 788 743 736 783 8.4 7.0 4.7 Q2 746 750 778 731 729 771 9.8 8.7 6.0 Q3 698 699 728 674 673 712 13.6 12.3 9.3 Q4 697 696 723 673 670 704 13.3 11.7 10.7 P1 708 694 745 671 656 718 12.5 11.7 7.7 Sampling : T1 = foot of profile Tl. T2 = top of profile T2. T3 = top of profile T3. The properties in T76511 temper, obtained by artificial aging under the following conditions: Q1 through Q4 : 12 h at 120 oC + 8 h 20 at 1500C Pl: 12 h at 120 oC + 10 h at 15600 29 are summarised in Table 7. Table 7 Alloy Rm [MPa] Rp0, 2 [MPa] A [% ] profil T1 T3 T2 T1 T3 T2 T1 T3 T2 e Q1 694 706 712 674 687 696 9.9 7.7 8.3 Q2 694 704 708 675 686 693 10.3 9.0 8.3 Q3 674 676 697 662 664 684 9.6 9.7 10.0 Q4 673 677 687 657 663 672 11.1 9.7 10.0 P1 659 644 686 615 589 643 12.1 10.3 9.1 Sampling: T1 = foot of profile TI. T2 = top of profile T2. T3 = top of profile T3. 5 It can be seen that compared with alloy Pl, alloys Q1 and Q2 have a mechanical strength significantly higher. Corrosion resistance was evaluated by means of the 10 EXCO test (standard ASTM G34) of the products Ql and Q2 in temper T6511 (unmachined specimens taken from the beginning of the extrusion) as EA or EB, and was generally at least as good as or better than what was observed for samples P1 through P3, and Q3 and Q4. 15 For R1 and R2, the following mechanical properties were found: 30 Table 8 T6511 temper T76511 temper Rm Rp0,2 A Rm Rp0,2 A [MPa] [MPa] [% ] [MPa] [MPa] [% ] Alloy RI, profile 753 738 8 688 669 10 T4(a) Alloy R1, profile 756 743 8 686 667 9 T4(b) Alloy R2, profile T5 755 743 7 676 659 10 NOTE : profile T4(a) = samples cut from the top of the profile, see figure 4, mark (a). 5 Example 4: Formability of profiles of type T1 according to example 3 was studied by using the three point bending test according to DIN standard 50 111 (September 1987, section 3.1). The location of sampling, a flat area, is 10 shown on figure 7. The important parameters of the three point bending test are shown on figure 9. The test was performed at 130 oC. Both tempers T6511 and T76511 were tested. The resulting values for the bending angle a (as defined on 15 figure 8) are summarised in Table 9. These are mean values calculated from half a dozen individual determinations using specimens cut at different locations distributed over the length of the profiles.

31 Table 9 Bending angle alloy temper T76511 temper T6511 Q1 43.40 Q2 38.10 36.90 Q3 33.90 33.80 P1 41.50 35.20 5 In all cases, the profiles according to the invention (Ql and Q2) have a formability which was comparable with the formability of profiles according to prior art (Q3 and Pl). 10 Example 5: Cold formability of samples similar to those used in the example 4 (in the unstable W temper after solution heat treatment and quenching) was studied at room temperature by using the same three point bending 15 technique. The variation of the bending angle a (as defined on figure 8) over the length of the profiles is small. Table 10 refers to values obtained in the W temper. 20 Table 10 Sample Bending angle Q1 27.10 Q2 27.30 Q3 33.60 32 P1 34.50 Example 6: Rolling ingots were elaborated by a process similar to the one described in example 1. The chemical 5 composition is given in Table 11. Hot rolled plates with a thickness of 25 mm were obtained by a process similar to the one described in example 1. Said plates were solution heat treated at a temperature between 472 and 480 oC for 2 hours, quenched and stretched with a 10 permanent set comprised between 1.5% and 2%. Finally, the plates were artificially aged at a temperature of 135 oC. Table 11 15 Alloy Zn Mg Cu Fe Si Zr Ti Mn Sc Mg/Cu M 9.94 3.02 0.78 0.04 0.03 0.10 0.06 0 0 3.87 3 N 10.0 2.72 0.77 0.06 0.04 0.10 0.05 0 0.1 3.53 0 5 0 K 9.90 2.03 1.55 0.03 0.03 0.10 0.05 0 0.1 1.31 0 The following mechanical properties were obtained: Table 12 20 Plate Duration Rm(L) Rp0,2(L) Rp0,2(L) A Kzc (or of aging [MPa] (tensile) (compressive) [%] Kq) [h] [MPa] [MPa] [MPaqm]

(NOTE

33 1) N 6 711 687 678 10.4 16.9 N 12 702 695 696 9.7 14.5 M 6 691 676 662 10.0 21.2 M 12 684 675 660 8.9 20.4 K 6 694 666 620 12.9 23.2 K 12 692 674 685 11.7 19.7 NOTE 1 : measured with B = 1 inch and W = 3 inches. It was checked that for plates N and K, the aging treatment of 12 h leads to the T6 temper. For aging times significantly longer, Rp0,2(L) and Rm(L) decrease. 5 It can be seen that for the same Zn content and with a similar Mg/Cu ratio, plate N (containing 0,10% scandium) exhibits better static mechanical properties than plate M (no scandium). For the same zinc content and for the same 10 scandium content, plate N (high Mg/Cu ratio) exhibits better RpO,2(L) and Rm(L) values than plate K. Example 7: For several profiles elaborated according to 15 example 3, the resistance to stress corrosion was evaluated. The results are summarised in Table 13. Table 13 Sample Temper Stress Duration [MPa] of the test Alloy Q1, profile Tl, L T76511 530 > 30 days direction Alloy Qi, profile TI, L T6511 350 > 30 days 34 direction Alloy P1, profile T4, L T76511 430 > 30 days direction Alloy P1, profile T4, LT T76511 400 > 30 days direction Alloy Pl, profile T4, LT T6511 280 > 30 days direction Alloy RI, profile T4, LT T direction 76511 Alloy Ri, profile T4, LT T direction 76511 It can be seen that the products according to the invention show a satisfactory resistance to stress corrosion. 5 Example 8: Extruded profiles made of alloys 7349 or 7449 were produced with and without scandium, according to a process similar to the one described in example 3. 10 Table 14 lists the chemical compositions, Table 15 the obtained mechanical properties. Table 14 Alloy Zn Mg Cu Fe Si Zr Ti Mn Cr Sc Mg/Cu Xl 8.1 2.5 1.7 0.08 0.01 0.15 0.03 0.1 0.1 0 1.47 7 4 X2 8.4 2.1 1.9 0.06 0.02 0.10 0.02 0 0 0 1.11 15 Table 15 Temper T6 Temper T76 35 product Rm Rp 0

,

2 A Rm Rp0,2 A [MPa] [MPa] [% ] [MPa] [MPa] [% ] Alloy Xl, profile Ti, 713 681 15 650 606 13 measured at (a) Alloy Xl, profile T2, 711 678 11 654 614 10 measured at (a) Alloy Xl, profile T2, 740 708 7 670 628 8.5 measured at (b) Alloy X2, profile TI, 673 645 17 645 626 14 measured at (a) Alloy X2, profile T2, 680 653 12 646 623 11 measured at (a) Alloy X2, profile T2, 728 699 10 667 632 11 measured at (b) The comparison with the results of example 3 shows that the products according to the invention have increased mechanical strength (Rm, Rp0.

2 ) compared with 5 products X1 and X2 according to prior art. Example 9: Seat tracks for aircraft have been manufactured from billets of chemical composition R1 and Q1 10 according to the previous examples. These profiles are "I" type profiles including a foot section, a centre section (web), and a top section on which the seats are fixed. The thickness of the centre section was of the order of 2 mm, and the height of the profile was of the 15 order of 65 mm. Table 16 summarises the static mechanical properties in temper T76511.

36 Table 16 Alloy Sampling Rm [MPaI Rp0, 2 [MPa] R1 Foot 688 669 R1 Top 686 667 Q1 Foot 672 643 Q1 Top 683 660 Stress corrosion tests according to ASTM G47 shows 5 good resistance to stress corrosion (in the transverse direction). Example 10: Digital models simulating damage tolerance of high 10 strength stiffeners according to the invention have been evaluated in order to determine the residual strength of a fuselage panel. Aviation authorities require a designed structure to resist limiting loads despite serious damage; the chosen damage is a crack 15 with the same length as two stiffeners (14, 16) with the central stiffener (18) broken (see figure 10). The inventors have noted that the residual strength of fuselage panels stressed in tension can strengthen high strength stiffeners according to the invention. The use 20 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 preventively avoids an unstable failure. The residual strength of 25 the cracked panel is thus improved. This effect can be used either to improve the safety factor in constructions in which stiffeners are replaced by 37 stiffeners according to the invention, or to lower the weight of the construction by using stiffeners with smaller sections and thinner fuselage sheets, and / or larger stiffener spacings. 5 Skin failure is governed by the stress intensity factor at the crack tip. For extrados parts of the fuselage at which the stiffener spacing is 200 mm and the stiffening factor (stiffener section / total section) is equal to 0.25, the stress intensity factor 10 for a crack with a length equal to the length of two stiffeners with the central stiffener broken in a panel assembled with stiffeners according to the invention will be 25% less than it would be with a panel with stiffeners made using the widely used 2024 T3 alloy. 15 For longer cracks, the 2024 stiffener will be increasingly stressed in the plastic range compared with new stiffeners that have not reached the yield strength. The difference in the stress intensity factor can be as high as 15%. 20 Example 11: Digital models for fuselage panels stressed in compression and / or shear were evaluated to determine the stability in shear and in compression. Aviation 25 authorities require that the structure should be designed to resist an ultimate load for 3 seconds without excessive deformation. Nevertheless, plastic deformation is allowed. This results in post-buckling designs for fuselage panels in locations critical for 30 stability. Although the buckling of perfect columns (Euler's theory) or very thin real structures is 38 essentially an elastic phenomenon (governed by Young's modulus), post-buckling designs include plastic deformation and can benefit from an increase in the yield strength. Figure 11 shows this buckling test. A 5 fuselage skin (20) is fixed to the stiffeners (14, 16) at attachment points (22), for example rivets. The buckling test causes deformation of the panel, resulting in a gap (24) between the stiffener (14, 16) and the skin (22). 10 The applicant has noted that the shear and compression stability of fuselage panels stressed in compression and / or shear can benefit from the high strength of stiffeners according to the invention. The use of stiffeners according to the invention as a 15 structural element in an aircraft fuselage panel can improve the shear and compression stability of fuselage panels, because these stiffeners are more stable in buckling. This effect can be used either to increase the safety factor in constructions in which stiffeners 20 are replaced by stiffeners according to the invention, or to lower the total weight of the construction by using stiffeners with smaller sections and thinner fuselage plates and / or larger stiffener spacings. Rivet spacings can also be reduced, which reduces the 25 assembly cost of the structure. An estimate of the improved stability in buckling can be obtained by applying a general method given in [Michael C.Y. Niu, Airframe Stress Analysis and Sizing, 2 nd edition, chapter 10]. The inventors have noted.that 30 when this method is used, the increased stability of the stiffener according to the invention (with 700 MPa 39 yield strength in compression and a Young's modulus of 73 GPa in compression) compared with a 7150 T77511 stiffener (with 538 MPa typical yield strength in compression and a Young's modulus of 73 GPa in 5 compression) that is widely used in aircraft according to the state of the art, is 15% or more for typical use of Z >> shaped stiffeners. Table 17 shows the parameters of the different stiffener geometries used for the calculations. Figure 10 12 compares predicted buckling stresses for these different geometries from Z1 to Z8 (from left to right). Table 17 15 Small "Z" stiffener design: ZI Z2 Z3 Z4 Z5 Z6 Z7 Z8 Free foot section width [mm] 12.7 12.7 12.7 12.7 12.7 12.7 12.7 12.7 Free riveted section width [mm] 25.4 25.4 25.4 25.4 25.4 25.4 25.4 25.4 Height [mm] 38.1 38.1 38.1 38.1 38.1 38.1 38.1 38.1 Free foot section thickness [mm] 1.0 1.5 1.5 2.0 1.0 1.5 1.5 1.5 Riveted foot section 1.0 1.0 1.5 1.5 1.0 1.0 1.0 1.5 40 thickness [mm] Web thickness [mm] 1.0 1.0 1.0 1.0 1.5 1.5 1.5 1.5 Section [m 2] 76 83 95 102 95 102 102 114 Equivalent thickness [mm] 1.0 1.1 1.3 1.3 1.3 1.3 1.3 1.5

Claims (23)

1. Rolled, extruded or forged product in Al-Zn Mg-Cu alloy, characterised in that it comprises (in mass percentage): a) Zn 8.3 - 14.0 Cu 0.3 - 2.0 Mg 5 0.5 - 4.5 Zr 0.03 - 0.15 Fe + Si < 0.25 b) at least one element selected from the group made up of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, the content of each of said elements, if 10 selected, being between 0.02 and 0.7%, c) remainder aluminium and inevitable impurities, and in that it satisfies the conditions d) Mg/Cu > 2.4 and 15 e) (7.9 - 0.4 Zn) > (Cu + Mg) > (6.4 - 0.4 Zn).

2. Product according to Claim 1, wherein Mg / Cu > 2.8 and preferably > 3.5 and even more preferably >

4.0. 20 3. Product according to either claim 1 or 2, wherein the maximum content of the elements Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb does not exceed 1.5% in total. 25 4. Product according to any of claims 1 to 3, wherein from the group made up of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb only titanium is selected. 42

5. Rolled, extruded or forged product in Al-Zn Mg-Cu alloy, characterised in that it comprises (in mass percentage): 5 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 made up of Zr, Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, 10 Ho, Er, Y, Yb, the content of each of said elements, if selected, being between 0.02 and 0.7%, c) remainder aluminium and inevitable impurities, and in that it satisfies the conditions 15 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, wherein the 20 maximum content of the elements Zr, Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb does not exceed 1.5% in total. 7. Product according to any of claims 1 to 4, 25 wherein Zn > 9.0%. 8. Product according to claim 7, wherein Zn > 30 9. Product according to claim 7, wherein the zinc content is comprised between 9.0% and 11.0%. 43

10. Product according to any of claims 1 to 9, wherein Cu > 0.6%. 5 11. Product according to any of claims 1 to 10, wherein Cu 0.6 - 1.2% and Mg 2.5 - 3.4%.

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

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

14. Product according to any of claims 1 to 13, 15 wherein Mg > 1.95 + 0.5 (Cu - 2.3) + 0.16 (Zn - 6) + 1.9 (Si - 0.04).

15. Product according to any of claims 1 to 14, wherein the maximum mass percentage of the following 20 elements is not exceeded: Cr 0.40 Mn 0.60 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 Dy 0.40 Ho 0.40 Er 0.40 25 Yb 0.40 Y 0.20.

16. Product according to any of claims 1 to 15, characterised in that it further comprises at least one element selected from the group made up of Ag, Sn, Cd, 30 Ge, In, the content of each of these elements, if selected, being comprised between 0.02% and 0.15%, and 44 preferably between 0.05% and 0.10%.

17. Extruded product according to any of claims 1 to 16, characterised in that said product exhibits in 5 temper T6511, measured on test pieces cut from a plane area of the profile, a) a bending angle, determined at 130 oC by means of a three point bending test according to DIN 50 111 (section 3.1) on a specimen of 1.6 mm thickness and 10 expressed as the mean value calculated from several individual measurements performed on specimens cut from different locations over the length of the profile, of at least 340, and b) a tensile yield strength (Rp0. 2 ) of at least 15 720 MPa, and preferably a bending angle of at least 350 and a TYS of at least 7 50 MPa.

18. Extruded product according to any of claims 1 20 to 16, characterised in that said product exhibits in temper T76511, measured on test pieces cut from a plane area of the profile, a) a bending angle, determined at 130 oC by means of a three point bending test according to DIN 50 25 111 (section 3.1) on a specimen of 1.6 mm thickness and expressed as the mean value calculated from several individual measurements performed on specimens cut from different locations over the length of the profile, of at least 370, and 30 b) yield strength (Rpo. 2 ) of at least 670 MPa. 45

19. Extruded product according to claim 18, wherein the stress corrosion resistance, determined by an EXCO test according to ASTM G34 in the T6511 temper on unmachined test pieces, is at least of level EB. 5

20. Structural aircraft member manufactured from a product according to any of claims 1 to 19.

21. Structural aircraft member according to claim 10 20, wherein said structural member is a fuselage stiffener.

22. Structural aircraft member according to claim 20, wherein said structural member is a seat track. 15

23. Seat track according to claim 22, having a ultimate tensile strength in the T76511 temper in the zone of fixation of the seats of at least 670 MPa, and preferably of at least 680 MPa. 20

24. Seat track according to either claim 22 or 23, characterised in that it has a tensile yield strength in the T76511 in the zone of fixation of the seats of at least 640 MPa and preferably of at least 25 660 MPa.

25. Structural aircraft member according to claim 20, wherein said structural member is a floor beam. 30

26. Aircraft comprising a fuselage assembled from a plurality of stiffeners and a plurality of sheets, 46 wherein at least part of said stiffeners are structural members according to any of claim 20.

27. Fuselage structure assembled starting from a 5 plurality of stiffeners according to claim 21 and a plurality of sheets, characterised in that for a spacing of said stiffeners of 200 mm and a stiffening ratio of 0.25, the stress intensity factor for a crack with a length of two stiffeners with the central 10 stiffener broken is at least 5% lower than it would be with stiffeners made using. the 2024 T3 alloy.

28. Fuselage structure assembled starting from a plurality of stiffeners according to claim 21 and a 15 plurality of sheets, characterised in that the buckling stability of said stiffeners is improved by at least 15% compared with an identical structure including Z shaped stiffeners with the same geometry made using the 7150 T77511 alloy.