Classifications

• • C—CHEMISTRY; METALLURGY
  • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
  • C22F—CHANGING THE PHYSICAL STRUCTURE OF NON-FERROUS METALS AND NON-FERROUS ALLOYS
  • C22F1/00—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working
  • C22F1/04—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of aluminium or alloys based thereon
  • C22F1/053—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of aluminium or alloys based thereon of alloys with zinc as the next major constituent
• • C—CHEMISTRY; METALLURGY
  • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
  • C22C—ALLOYS
  • C22C21/00—Alloys based on aluminium
  • C22C21/10—Alloys based on aluminium with zinc as the next major constituent

Definitions

• 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.
• 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.
• the alloys 7150, 7050 and 7349 are also used for the manufacture of fuselage stiffeners.
• 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.
• 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.
• Métallurgique de Gerzat discloses a process for obtaining such bottles.
• European patent application EP 257 167 Al 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.
• 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.
• 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
• 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
• 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.
• 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.
• 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.
• 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.
• damage tolerance is not a limiting factor, and we can therefore allow our to optimize the elastic limit and the breaking limit at the expense of damage tolerance, while taking care not to degrade corrosion resistance.
• 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.
• 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 .
• 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.
• 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.
• 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 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.
• addition elements capable of precipitating during maturation or tempering treatment, in order to be able to exhibit advantageous static mechanical characteristics.
• the content of these addition elements must fulfill the condition Mg + Cu> 6.4 - 0.4 Zn.
• a sufficient content of so-called anti-recrystallizing elements must be added.
• 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%.
• anti-recrystallizing elements in the form of fine precipitates formed during thermal or thermomechanical treatments, block the recrystallization.
• 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.
• 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%.
• 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.
• 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
• 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.
• 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.
• 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.
• 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.
• 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.
• 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.
• the flat test piece 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.
• 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.
• 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 sheet C presents a good compromise between mechanical strength and elongation.
• sheet D outside the invention, its mechanical strength is significantly better.
• 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.
• sheet B outside the invention, the elongation at break of sheet C is significantly better.
• sheet B 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.
• 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.
• 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.
• 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
• 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.
• 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 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.
• the T6511 and T76511 states have been tested.
• the values of the folding angle ⁇ (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
• 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).
• Lamination plates were produced by a process similar to that described in Example 1.
• the chemical composition is given in Table 11.
• 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.
• the resistance to stress corrosion resistance was characterized for several profiles produced according to example 3.
• Table 14 indicates the chemical compositions, Table 15 the mechanical characteristics obtained. Table 14
• 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.
• 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.
• 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.
• 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%.
• 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.
• 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).

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