EP1382698A1 - Wrought product in Al-Cu-Mg alloy for aircraft structural element - Google Patents

Abstract

Weldable product, notably rolled, drawn or forged, is made in an alloy with the following composition (by weight %): (a) Cu: 3.8 - 4.3; (b) Mg: 1.25 - 1.45; (c) Mn: 0.2 - 0.5; (d) Zn: 0.4 - 1.3; (e) Fe less than 0.15; (f) Si less than 0.15; (g) Zr at most 0.05; (h) Ag less than 0.01; (i) other elements each less than 0.05 and less than 0.15 in total; (j) remainder Al; (k) treated by putting into solution, tempering and cold drawing, with a permanent deformation of between 0.5 and 15 %, and preferably of between 1.5 and 3.5 %. The cold drawing is achieved by controlled traction and/or cold transformation, such as rolling or drawing. Independent claims are also included for: (i) a plated sheet of this alloy; (ii) an aircraft structural element of this alloy; (iii) a method for the fabrication of a weldable product of this alloy.

Description

Field of the invention

The invention relates to aircraft structural elements, in particular sheet metal for large commercial aircraft fuselage made from rolled, spun or forged alloy AlCuMg in the treated state by dissolving, quenching and cold work hardening, and having, compared to the products of the prior art used for the same application, an improved compromise between the different properties required.

State of the art

The fuselage of commercial aircraft of large capacity is typically made of an AlCuMg alloy sheet skin, as well as longitudinal stiffeners and circumferential frames. A type 2024 alloy is used which, according to the designation of the Aluminum Association or the standard EN 573-3, has the following chemical composition (% by weight):
If <0.5 Fe <0.5 Cu: 3.8 - 4.9 Mg: 1.2-1.8 Mn: 0.3 - 0.9
Cr <0.10 Zn <0.25 Ti <0.15.

Variants of this alloy are also used. These elements are asked structures a compromise between several properties such as: resistance mechanical (i.e. static mechanical characteristics), tolerance to damage (toughness and speed of fatigue cracking), fatigue resistance (especially oligocyclic), resistance to different forms of corrosion, fitness ability. In some cases, especially for airplanes supersonic, creep resistance can be critical.

In order to improve the compromise between the different properties required, in particular the mechanical strength and toughness, various alternative solutions have been proposed. Boeing developed the alloy 2034 of composition:
If <0.10 Fe <0.12 Cu: 4.2 - 4.8 Mg: 1.3 - 1.9
Mn: 0.8 - 1.3 Cr <0.05 Zn <0.20 Ti <0.15 Zr: 0.08 - 0.15

This alloy was the subject of patent EP 0 031 605 (= US 4,336,075). He introduces, compared to 2024 in the T351 state, a better specific yield strength due to the increase in the manganese content and the addition of another antirecrystallizer (Zr), as well as improved toughness and fatigue resistance.

US Patent 5,652,063 (Alcoa) relates to an aircraft structural element made from a composition alloy (% by weight):
Cu: 4.85-5.3 Mg: 0.51-1.0 Mn: 0.4-0.8 Ag: 0.2 - 0.8
If <0.1 Fe <0.1 Zr <0.25 with Cu / Mg between 5 and 9.
The sheet of this alloy in the T8 state has a yield strength> 77 ksi (531 MPa). The alloy is particularly intended for supersonic aircraft.

EP 0 473 122 (= US Pat. No. 5,213,639) to Alcoa discloses an alloy, registered at the Aluminum Association as 2524, of composition: Si <0.10 Fe <0.12 Cu: 3.8 - 4.5 Mg: 1.2 - 1.8 Mn: 0.3 - 0.9 may contain possibly another antirecrystallizer (Zr, V, Hf, Cr, Ag or Sc). This alloy is intended more particularly for thin sheets for fuselage and has a tenacity and improved crack propagation resistance over 2024.

The patent application EP 0 731 185 of the Applicant relates to an alloy, subsequently recorded under No. 2024A, of composition: Si <0.25 Fe <0.25 Cu: 3.5 - 5 Mg: 1 - 2 Mn < 0.55 with the relation: 0 <Mn-2Fe <0.2
The thick plates in this alloy have both improved toughness and a reduced level of residual stresses, without loss on the other properties.

US Patent 5,593,516 (Reynolds) relates to an alloy for aeronautical applications containing from 2.5 to 5.5% Cu and 0.1 to 2.3% Mg, in which the contents of Cu and Mg are maintained below their limit of solubility in aluminum, and are linked by the equations: Cu max = 5.59-0.91Mg and Cu min = 4.59 - 0.91Mg The alloy may also contain: Zr <0.20% V <0.20% Mn <0.80%
Ti <0.05% Fe <0.15% If <0.10%.

The patents US 5,376,192 and US 5,512,112, issued from the same initial application, relate to alloys of this type containing 0.1 to 1% silver. We can notice that the use of silver in this type of alloy leads to a increase in the cost of development and difficulties in the recycling of waterfalls. manufacturing.

The patent application EP 1 170 394 A2 (Alcoa) describes four AlCu-type alloys which have, respectively, the composition
Cu 4.08, Mn 0.29, Mg 1.36, Zr 0.12, Fe 0.02, Si 0.01;
Cu 4.33, Mn 0.30, Mg 1.38, Zr 0.10, Fe 0.01, Si 0.00;
Cu 4.09, Mn 0.58, Mg 1.35, Zr 0.11, Fe 0.02, Si 0.01; and
Cu 4.22, Mn 0.66, Mg 1.32, Zr 0.10, Fe 0.01, Si 0.01.
The patent teaches how to convert these products into sheets having an elongated grain structure, wherein the grains have a length to thickness ratio of greater than 4. While respecting both a specific microstructure and texture, this product has good characteristics. mechanical resistance and damage tolerance. One of the disadvantages of these alloys is to be based on a high purity aluminum (very low silicon and iron content), which is expensive. Another patent of the same applicant, US Pat. No. 5,630,889, discloses a sheet in the T6 or T8 state of AlCuMg alloy containing:
Cu 4.66, Mg 0.81, Mn 0.62, Fe 0.06, Si 0.04, Zn 0.36%.
A silver addition enhances the properties of this alloy. However, money is an expensive element, and it limits the possibilities of recycling of the products thus obtained as well as their production falls, which contributes to increase even more the cost price of these products.

The present invention aims to obtain aircraft structural elements, and in particular fuselage elements made of AlCuMg alloy having, compared to the prior art, improved damage tolerance, mechanical resistance to less equal, an improved resistance to corrosion, and this without resorting to expensive and troublesome addition elements for recycling.

Object of the invention

The subject of the invention is a wrought product, in particular a laminated, spun or forged product, made of an alloy of composition (% by weight):
Cu 3.80-4.30, Mg 1.25-1.45, Mn 0.20-0.50, Zn 0.40-1.30, Zr ≤ 0.05, Fe <0.15, Si < 0.15, Ag <0.01
other elements <0.05 each and <0.15 in total, remain Al,
said product can be treated by dissolving, quenching, and cold working, with a permanent deformation of between 0.5% and 15%, preferably between 1% and 5%, and even more preferably between 1.5% and 3.5%. Cold working can be obtained by controlled pulling and / or cold processing, for example rolling or drawing.

The invention also relates to a structural element for construction aeronautical equipment, in particular an aircraft fuselage element, manufactured from such an aircraft wrought product, and especially from such a rolled product.

Description of the invention

Unless stated otherwise, all the information relating to the chemical composition of the alloys is expressed in percent by weight. Therefore, in a mathematical expression, "0.4 Zn" means: 0.4 times the zinc content, expressed in mass percent; this applies mutatis mutandis to other chemical elements. The designation of the alloys follows the rules of The Aluminum Association. The metallurgical states are defined in the European standard EN 515. Unless stated otherwise, the static mechanical characteristics, that is to say the breaking strength R _m , the yield strength R _p0,2 , and the elongation at rupture A, are determined by a tensile test according to EN 10002-1. The term "spun product" includes so-called "stretched" products, i.e., products that are made by spinning followed by stretching.

In the AlCuMg alloys of the prior art that are the most efficient for the manufacture of aircraft fuselage structural elements, a good level of toughness is obtained by specifying very low levels of iron and silicon, and limiting the copper and magnesium contents to facilitate the dissolution of the particles coarse intermetallic. To obtain a sufficient level of mechanical resistance, the person skilled in the art is inclined to maintain a significant content of manganese, since it contributes to the hardening of the alloy. Almost all alloys of the 2xxx series contain not more than 0.25% zinc.

The copper content of the alloy according to the invention is between 3.80 and 4.30 %, and preferably between 4.05 and 4.30%; so it's in the low half of the content range of alloy 2024, so as to limit the volume fraction residual coarse copper particles. For the same reason, the interval of magnesium content, which must be between 1.25 and 1.45% and preferably between 1.28 and 1.42%, is shifted downwards compared to that of 2024. The manganese is maintained between 0.20 and 0.50%, preferably between 0.30 and 0.50, and more preferably between 0.35 and 0.48%. The implementation of the invention does not require significant addition of zirconium at a content greater than 0.05%.

The present invention requires careful control of the zinc content, the alloy being discharged into copper, magnesium and manganese. The zinc content must be between 0.40 and 1.30%, preferably between 0.50 and 1.10%, and still between more preferably between 0.50 and 0.70%. In one embodiment advantageous when the contents of copper, magnesium and manganese are lower than at 4.20%, 1.38% and 0.42% respectively, it is preferable that the zinc content at least equal to (1.2Cu - 0.3Mg + 0.3Mn - 3.75).

According to the findings of the plaintiff, this unloading of the copper, magnesium and manganese and adding an exactly controlled amount of Zinc leads, using appropriate methods of implementation, to sheets which have approximately the same mechanical resistance, but a better tolerance damage to sheet metal that does not contain this zinc addition, to a at least as good formability, and better corrosion resistance.

The silicon and iron contents are each kept below 0.15%, and preferably below 0.10%, to have good toughness. The man of know that the reduction of the iron and silicon content improves tolerance to damage of AlCuMg and AlZnMgCu alloys used in aircraft construction (See the article by J.T. Staley, Microstructure and Toughness of High Strength Aluminum Alloys ", published in" Properties Related to Fracture Toughness ", ASTM STP605, ASTM, 1976, pp. 71-103). However, this is only in very cases (depending on the type of alloy and the intended application) that the gain in damage tolerance related to the use of aluminum containing less than 0.06 % iron and silicon each is large enough to be valued. The implementation of the present invention does not require that the content of iron and silicon less than 0.06% each, because in the selected composition range, the Damage tolerance is very good.

Finally, unlike the alloys described in US Pat. No. 5,376,192, US 5 512,112 and US 5,593,516, the alloy contains no silver addition, nor any other element likely to increase the production cost of the alloy and to pollute the other alloys produced on the same site by recycling manufacturing scrap.

The preferred method of manufacture comprises casting of plates, in the case where the product to be made is a rolled sheet, or billets in the case where it is a spun section or a forged part. The plate or the billet is scalped, then homogenized between 450 and 500 ° C. The hot transformation is then carried out by rolling, spinning or forging, optionally completed by a cold transformation step. The laminated, spun or forged half-product is then dissolved between 480 and 505 ° C., so that this dissolution is as complete as possible, that is to say that the maximum of potentially soluble phases , in particular the precipitates Al ₂ Cu and Al ₂ CuMg, are effectively put back in solution. The quality of the dissolution can be assessed by differential enthalpy analysis (AED) by measuring the specific energy using the area of the peak on the thermogram. This specific energy must preferably be less than 2 J / g.

Then we quench with cold water, and then cold hardening leading to a permanent elongation of between 0.5% and 15%. This hardening cold can be a controlled pull with a permanent elongation between 1 and 5% bringing the product to a T351 state. Controlled traction with a permanent elongation of between 1.5% and 3.5%. It can also be a cold rolling by rolling in the case of sheets or by drawing in the case profiles, with a permanent elongation of up to 15%, bringing the in the T39 state, or in the T3951 state if the rolling or drawing is combined with the traction. The product finally undergoes natural aging at room temperature. The final microstructure is in general largely recrystallized, with grains relatively thin and fairly equiaxed.

The product according to the present invention is well suited for use as an aircraft structural element, for example as a fuselage skin element, and especially as an element for the fuselage skin panel (skin). . These sheets, preferably plated, have a thickness of between 1 and 16 mm, and have good resistance to intergranular corrosion and corrosion on riveted assembly. They have a breaking strength in the L direction and / or TL direction greater than 430 MPa, and preferably greater than 440 MPa, and a yield strength in the L and / or TL direction greater than 300 MPa, and preferably greater than 320. MPa. They have a good formability (elongation at break L and / or TL greater than 19% and preferably greater than 20%), and a damage tolerance Kr, calculated from a curve R obtained according to ASTM E 561 for a Δa _eff value of 60 mm, greater than 165 MPa√m in the TL and LT directions, greater than 180 MPa√m in the LT direction, and a crack propagation rate da / dN, determined according to the ASTM E standard. 647 in the TL or LT direction for a ΔK value of 50 MPa√m, less than 2.5 × 10 ^-2 mm / cycle (and preferably less than 2.0 × 10 ^-2 mm / cycle) and a charge ratio R = 0 1. This type of property compromise is particularly suitable for fuselage lining. The sheet according to the invention may be a sheet plated with at least one face with an alloy of the 1xxx series, and preferably with an alloy selected from the group consisting of alloys 1050, 1070, 1300 and 1145.

In view of the fact that riveting is the most frequently used assembly method for fuselage skins, it is preferred for the application as fuselage coating of plated sheets according to the invention which are particularly resistant to corrosion by galvanic coupling. in a riveted assembly. More particularly, plated sheets which exhibit a galvanic corrosion current of less than 4 μA / cm ² , and preferably less than 2.5 μA / cm ² , for exposure of up to 200 hours, during corrosion tests in a riveted assembly, placing the core alloy in a non-deaerated solution containing 0.06 M NaCl and the plating alloy in a 0.02 M AlCl _{3 solution} deaerated by nitrogen sparging.

In the examples which follow, an illustration is given of modes of advantageous embodiment of the invention. These examples are not limiting in nature.

Examples Example 1

Four N0, N1, N2 and N3 alloys were developed with chemical composition was in accordance with the invention. The liquid metal was first treated in the furnace maintaining by injection of gas using a rotor of a type known under the brand IRMA, and then in a pocket of type known as Alpur. The refining has been done online, that is to say between the holding oven and the Alpur pocket, with wire AT5B (0.7 kg / t for N0, N1 and N3, 0.3 kg / t for N2). 3.0 plates were cast m length and section 1450 mm x 377 mm (except for N3: section 1450 x 446 mm). They were relaxed for 10h at 350 ° C.

Alloy plates 2024 according to the prior art (references E and F) have also been developed by the same method.

The chemical compositions of the alloys N0, N1, N2, N3, E and F, measured on a spectrometric counter taken from the casting channel, are collated in Table 1: Chemical composition Alloy Yes Fe Cu mn mg Zn Cr N0 0.03 0.08 4.16 0.41 1.35 0.59 * 0,001 N1 0.03 0.08 4.00 0.40 1.22 0.63 N2 0.03 0.07 3.98 0.39 1.32 0.59 N3 0.06 0.07 4.14 0.43 1.26 1.28 E 0.06 0.19 4.14 0.51 1.36 0.11 0,007 F 0.06 0.16 4.15 0.51 1.38 0.12 0.014 Plating 1050 0.14 0.25 0,003 0,029 0,001 0,017

In all cases, alloy plating 1050 corresponds to about 2% of thickness.

For the alloys according to the prior art (alloys E and F), the plates have been heated to around 450 ° C, then hot-rolled at the reversing mill up to a thickness of about 20 mm. The resulting strips were rolled on a tandem rolling mill with three cages up to a final thickness close to 5 mm, then wound (at temperatures of 320 ° C and 260 ° C, respectively for alloys F and E). In the case of alloy F, the coil thus obtained was laminated to cold to a thickness of 3.2 mm. Sheet metal was cut, put in solution in a salt bath oven at a temperature of 498.5 ° C for a duration of 30 minutes (sheet metal E thickness 5 mm) or 25 min (sheet F thickness 3.2 mm), then completed (wrinkling followed by controlled traction with permanent elongation included between 1.5 and 3%).

As regards the alloys according to the invention, the N0 plate has undergone the following homogenization cycle: 8h at 495 ° C + 12h at 500 ° C (nominal values) while alloys N1, N2 and N3 have been homogenized for 12 hours at 500 ° C. After reheating (about 18 hours at 425-445 ° C), the plates were hot-rolled (inlet temperature: 413 ° C) to a thickness of about 90 mm. The N0 band thus obtained was cut in half in the direction perpendicular to the rolling direction. There were thus obtained two bands, labeled N01 and N02. These strips were rolled on a tandem hot rolling mill 3 cages to a final thickness of 6 mm (winding temperature about 320 - 325 ° C).

One N1 and N3 alloy plate and one N3 alloy plate were hot rolled at 5.5 mm before being cold rolled to the final thickness of 3.2 mm, and another N1 alloy plate. was hot rolled at 4.5 mm before being cold rolled to the final thickness of 1.6 mm.
An N2 alloy plate was hot rolled to a final thickness of 6 mm (tandem winding temperature 270 ° C).

The coil N01 has not undergone any other rolling pass, while the coil N02 was cold rolled to a final thickness of 3.2 mm.

The plates once cut were dissolved in a salt bath oven (thickness 6 mm: 60 minutes at 500 ° C., thickness 3.2 mm: 40 minutes at 500 ° C.; thickness 1, 6 mm: 30 minutes at 500 ° C) followed by quenching with water to about 23 ° C. After quenching, the sheets have undergone wrinkling and pulling with elongation cumulative standing between 1.5 and 3.5%. The waiting time between quenching and wrinkling did not exceed 6 hours.

The tensile strength R _m (in MPa), the conventional yield stress at 0.2% elongation R _p0.2 (in MPa) and the elongation at break A (in%) were measured by a tensile test according to EN 10002-1.

The results of measurements of static mechanical properties in the T351 state are shown in Table 2: Static mechanical characteristics sheet metal Ep [mm] Meaning L TL direction R _m [MPa] R _p0.2 [MPa] AT [%] R _m [MPa] R _p0.2 [MPa] AT [%] N01 6.0 442 336 22.8 442 323 23.5 N02 3.2 456 353 20.3 449 318 24.7 N1 1.6 455 359 20.2 434 298 21.8 N1 3.2 460 360 19.3 438 308 22.3 N2 6 471 384 19.8 462 343 19.9 N3 3.2 453 360 21.3 443 317 24.2 E 5.0 Not measured 456 341 17.7 F 3.2 454 318 19.2

The formability characterized by the ductility in tension (value of the elongation A) seems better for the alloy according to the invention, and this, for the two thicknesses considered. The formability of sheets with a thickness greater than 4 mm was also characterized using the LDH (Limit Dome Height) test on formats of 500 mm x 500 mm in the T351 state. The following results were obtained: Sheet N01 (thickness 6 mm) LDH = 81 mm Sheet E (thickness 5 mm) LDH = 75 mm This confirms the best formability of the alloy according to the invention.

The damage tolerance has been characterized in several ways. Curve R was measured according to ASTM standard E 561 on specimens of CCT type, W = 760 mm wide, 2a0 = 253 mm, e = sheet thickness, with piston displacement control and a pulling speed. of 1 mm / min, using anti-warp steel mounting. The specimens were taken in the TL and LT directions. The value of K _r [MPa√m] has been calculated for different values of Δ a _eff [mm].

The results are shown in Table 3: Results of the R curve test sheet metal Ep [mm] meaning K _r [MPa√m] for a value Δ a _eff of 10 mm 20 mm 30 mm 40 mm 50 mm 60 mm N02 3.2 TL 81 108 129 148 164 180 N01 6.0 TL 77 105 127 144 159 173 N1 1.6 TL 102 123 138 152 164 175 N1 3.2 TL 85 110 130 147 161 175 N2 6 TL 89 117 137 153 167 179 N3 3.2 TL 91 119 139 155 168 181 F 3.2 TL 82 107 125 139 151 162 E 5.0 TL 83 105 120 132 142 151 N2 3.2 LT 84 119 145 166 184 199 N1 6.0 LT 90 122 145 163 179 193 N1 1.6 LT 92 118 138 157 174 191 N1 3.2 LT 88 119 142 162 179 196 N2 6 LT 87 121 145 164 180 194 N3 3.3 LT 93 125 148 168 184 199 E 5.0 LT 104 126 141 154 165 174

It is found that for high values of Δ a _eff [mm], the product according to the invention exceeds the standard product alloy 2024.

The product according to the invention therefore has a better breaking strength in the case of a cracked panel.

The cracking rate da / dN (in mm / cycle) for different levels of ΔK (expressed in MPa√m) was determined according to the ASTM E 647 standard on CCT specimens taken in the TL direction and in the LT direction. , width W = 400 mm, 2ao = 4 mm, e = thickness of the sheet, under conditions of R = 0.1 and with a maximum stress of 120 MPa and an anti-fogging device for the lower thickness test pieces at 3.2 mm. The results are shown in Table 4. Results of the propagation velocity test sheet metal Ep [mm] meaning da / dN [mm / cycle] for Δ K [MPa√m] of 10 20 30 40 50 N02 3.2 TL 1.5 10 ^-4 6.5 10 ^-4 1.5 10 ^-3 0.4 10 ^-2 1.0 10 ^-2 N01 6.0 TL 1.5 10 ^-4 9.3 10 ^-4 1.8 10 ^-3 0.6 10 ^-2 1.4 10 ^-2 N1 1.6 TL 1.6 10 ^-4 4.6 10 ^-4 1.4 10 ^-3 0.4 10 ^-2 1.0 10 ^-2 N1 3.2 TL 1.8 10 ^-4 7.2 10 ^-4 1.6 10 ^-3 0.4 10 ^-2 1.0 10 ^-2 N2 6 TL 2.1 10 ^-4 8.7 10 ^-4 2.3 10 ^-3 0.6 10 ^-2 1.6 10 ^-2 N3 3.2 TL 1.6 10 ^-4 7.0 10 ^-4 1.4 10 ^-3 0.4 10 ^-2 0.8 10 ^-2 F 3.2 TL 1.4 10 ^-4 8.2 10 ^-4 3.2 10 ^-3 1.0 10 ^-2 2.9 10 ^-2 E 5.0 TL 1.9 10 ^-4 14.0 10 ^-4 6.1 10 ^-3 1.9 10 ^-2 4.4 10 ^-2 N02 3.2 LT 1.5 10 ^-4 5.4 10 ^-4 1.8 10 ^-3 0.5 10 ^-2 1.4 10 ^-2 N01 6.0 LT 1.8 10 ^-4 8.8 10 ^-4 1.4 10 ^-3 0.5 10 ^-2 1.1 10 ^-2 N1 1.6 LT 1.2 10 ^-4 4.42 10 ^-4 1.2 10 ^-3 0.3 10 ^-2 0.8 10 ^-2 N1 3.2 LT 1.7 10 ^-4 4.9 10 ^-4 1.8 10 ^-3 0.6 10 ^-2 1.6 10 ^-2 N2 6 LT 1.9 10 ^-4 10.4 10 ^-4 2.5 10 ^-3 0.7 10 ^-2 1.3 10 ^-2 N3 3.2 LT 1.66 10 ^-4 5.1 10 ^-4 1.6 10 ^-3 0.4 10 ^-2 1.0 10 ^-2 E 5.0 LT 1.5 10 ^-4 7.6 10 ^-4 2.4 10 ^-3 0.8 10 ^-2 2.2 10 ^-2

It can be seen that the plates of 2024, in particular for ΔK ≥ 20 MPa√m, exhibit a cracking rate two to three times higher than for the product according to the invention. The latter therefore allows longer inspection intervals (to mass given structure) or reductions in the inspection interval structure fixed.

As regards the R curves and the ΔK values, it should be noted that the most significant values with respect to the behavior of the real structure of an aircraft are in the range of 15 to 60 MPa√m.

Indeed, the fatigue stresses in a fuselage skin are generally of the order of 50 to 100 MPa, for detectable defects of the order of 20 to 50 mm, knowing that K = σ √ (πa), where σ is the constraint and the parameter a means the size of the default.

For a stiffener spacing greater than 100 mm, the breaking K values for a limiting load greater than 200 MPa are greater than about 120 MPa√m for the described R curves, with apparent K (K _r ) greater than about 110 MPa m. This means that the dimensioning portion of the curve R consists of points corresponding to a static crack advance Δ a _eff of more than 20 mm.

The corrosion resistance of the sheets has also been characterized. We observe that the alloy according to the invention shows intrinsically, that is to say after displacement by machining, resistance to intergranular corrosion, measured according to ASTM standard G 110, substantially comparable to that of the reference 2024.

On plated plates, measurement of the corrosion potential in the core and in the plating according to ASTM G 69 gave the results given in Table 5 below. These results show no significant difference in the potential difference between core and plating (characteristic of the cathodic protection power of a plating). This is surprising inasmuch as, in accordance with published data (see in particular "ASM Handbook", 9 ^th Edition, Volume 13, "Corrosion", page 584, figure 5), the addition of zinc in an aluminum alloy decreases significantly the corrosion potential, which should have had the effect of limiting the potential difference between core and plating of the alloy according to the invention. Potential [mV / ECS] and potential deviations [mV] sheet metal Ep [mm] Potential of the soul [mV / ECS] Potential of plating [mV / ECS] Potential gap [mV] N02 3.2 -620 -768 148 N01 6.0 -611 -801 190 N1 1.6 -634 -772 138 N1 3.2 -632 -775 143 N2 6 -636 -770 134 N3 3.2 -636 -755 119 E 5.0 -609 -775 166

On the other hand, and surprisingly, it is found that during a corrosion test by galvanic coupling in a riveted assembly, the product according to the invention behaves significantly better. According to the findings of the applicant, this test, which has been described for example in patent EP 0 623 462 B1, is particularly relevant for evaluating the ability of plated sheets for use in aircraft construction. The test consists in measuring the current which is established naturally between the anode (alloy of plating placed in a cell containing a solution of AlCl ₃ (0,02 M, deaerated by sparging of nitrogen)) and the cathode (alloy core placed in a cell containing a solution of NaCl (0.06 M, aerated)), a salt bridge ensuring the electrolytic contact between the two cells. Both elements (veneer and core) have the same surface (2.54 cm ² ). The coupling current densities are recorded throughout the duration of the test. It is observed that the current reaches a plateau after about 55 hours and hardly changes during the tests (200 hours or 15 days, depending on the sample). The results are summarized in Table 6. Electrochemical simulation of the assembly Sheet N2 Sheet N1 Sheet F Sheet E Current plateau after 55 hours [μA / cm ² ] 1.6 1.2 2.8 2.4 Measured mass loss [mg / cm ² ] after 5 days of testing 1.06 0.79 1.57 Not measured

By way of comparison, the examples described in patent specification EP 0 623 462 B1 give, for the standard alloy 2024 plated with an alloy 1070, a plateau current of 3.1 μA / cm ² .

It is found that the product according to the invention (N1 and N2) has a current of corrosion and mass loss much lower than the standard product according to the prior art. For some applications, for example structural elements for aircraft, this provides a very significant benefit in terms of service life.

Example 2

From hot-rolled and possibly cold-rolled sheets (state F) of the alloy according to the invention (see example 1), several other metallurgical states were developed in the form of a size of 600 mm (L-direction) x 160 mm (TL direction) x thickness. Raw lamination sheets of 3.2 mm (cold rolled) or 6.0 mm (hot rolled) were subjected to dissolution followed by quenching, ripening and pulling controlled, as shown in Table 7: Conditions for preparing the sheets of Example 2 landmark Thickness [mm] Solution time at 500 ° C [min] Maturation time Controlled traction N0A 3.2 30 <2 h 2% N0B 3.2 30 <2 h 4% N0C 3.2 30 <2 h 6% N0D 3.2 30 24 h 2% N0E 3.2 30 24 h 6% N0F 6.0 40 <2 h 2% N0G 6.0 40 <2 h 4% N0H 6.0 40 <2 h 6% N0I 6.0 40 24 h 2% N0J 6.0 40 24 h 6%

The marks ending in A, D, F and I correspond to T351 states. The different samples were characterized by tensile tests (L and TL directions) as well as toughness tests.

The tenacity was first evaluated in the TL and LT directions using the maximum stress R _e (in MPa) and the flow energy E _ec according to the Kahn test. The stress Kahn is equal to the ratio of the maximum load F _max that the specimen can withstand on the section of the specimen (product of the thickness B by the width W). The flow energy is determined as the area under the force-displacement curve up to the maximum force F _max supported by the specimen. The test is described in the article "Kahn-Type Tear Test and Crack Toughness of Aluminum Alloy Sheet", published in the journal Materials Research & Standards, April 1964, p. 151- 155. The sample used for the test of tenacity Kahn is described, for example, in the "Metals Handbook", 8 ^th Edition, vol. 1, American Society for Metals, pp. 241-242.

The tenacity was also approached for sheets of thickness 6 mm, using a curve-type test R, in the TL direction, but on smaller specimens than that described in Example 1. used CT specimens, width W = 127 mm, at ₀ = 38.5 mm, e = sheet thickness, with piston displacement control and a tensile speed of 1 mm / min.

The different results are given in Tables 8 and 9 below. Static mechanical characteristics landmark Maturation Traction Static characteristics meaning L TL sense static characteristics R _m [MPa] R _p0.2 [MPa] AT [%] R _m [MPa] R _p0.2 [MPa] AT [%] N0A <2 h 2% 450 345 21.6 444 307 23.7 N0B <2 h 4% 456 369 21.4 448 322 21.1 N0C <2 h 6% 464 394 17.6 453 339 18.2 N0D 24 h 2% 457 351 22.1 449 313 23.2 N0E 24 h 6% 473 413 18.7 464 352 18.6 N0F <2 h 2% 433 334 22.5 432 297 21.5 N0G <2 h 4% 437 353 22.3 436 308 21.1 N0H <2 h 6% 443 375 19.5 443 324 20.9 N0I 24 h 2% 440 338 24.1 443 308 23.1 N0J 24 h 6% 459 399 20.2 460 347 18.6 √ Toughness characteristics landmark Maturation Traction Kahn test tube test R curve test on specimen CT127 R _e [MPa] / E _ec [J] TL direction TL direction Meaning of LT K _app [MPa√m] K _eff [MPa√m] N0A <2 h 2% 163 / 15.0 166 / 15.4 Not measured N0B <2 h 4% 164 / 13.3 169 / 13.7 Not measured N0C <2 h 6% 167 / 12.3 172 / 12.9 Not measured N0D 24 h 2% 164 / 14.3 168 / 15.5 Not measured N0E 24 h 6% 172 / 12.0 176 / 12.4 Not measured N0F <2 h 2% 160 / 29.0 163 / 30.7 99.3 149.2 N0G <2 h 4% 165 / 28.4 166 / 27.8 99.9 137.6 NOH <2 h 6% 167 / 25.5 167 / 25.1 93.8 125.5 NOI 24 h 2% 165 / 30.0 165 / 28.9 99.6 149.3 NOJ 24 h 6% 172 / 24.0 172 / 24.2 101.1 137.1

Example 3

Sheets made according to Example 2 were subjected to a 5% cold work (by controlled pulling) after quenching. Tables 10 and 11 show the results of the characterizations. √ Static mechanical characteristics sheet metal Ep [mm] Meaning L TL direction Rm [MPa] R _p0.2 [MPa] AT [%] Rm [MPa] R _p0.2 [MPa] AT [%] N1 1.6 468 404 20.1 456 341 20.6 N1 3.2 472 408 18.2 464 348 19.3 N2 6 488 422 19.1 475 368 20.2 Results of the R curve test on 5% retracted sheet metal sheet metal Ep [mm] Meaning K _r [MPa√m] for a value Δ a _eff of 10 mm 20 mm 30 mm 40 mm 50 mm 60 mm N1 1.6 TL 66 91 112 130 148 164 N1 3.2 TL 96 124 144 160 173 186 N2 6 TL 84 111 131 147 161 173 N1 1.6 LT 86 111 132 152 171 189 N1 3.2 LT 101 133 157 178 195 212 N2 6 LT 82 112 136 157 175 192

Claims (23)

Wrought product, in particular rolled, spun or forged, of AlCuMg type alloy, characterized in that it comprises (% by weight):
Cu 3.80-4.30, Mg 1.25-1.45, Mn 0.20-0.50, Zn 0.40-1.30, Fe <0.15, Si <0.15, Zr ≤ 0.05, Ag <0.01
other elements <0.05 each and <0.15 in total, remain Al. The product of claim 1 wherein Cu 4.05 - 4.30. The product of claim 1 or 2, wherein Mg 1.28 - 1.42. Product according to any one of claims 1 to 3, wherein
Mn 0.30 - 0.50 and preferably Mn 0.35 - 0.48. Product according to any one of claims 1 to 4, wherein Zn 0.50-1.10 and preferentially Zn 0.50 - 0.70. A product according to any one of claims 1 to 5 wherein Fe <0.10. Product according to any one of claims 1 to 6, wherein Si <0.10. Product according to claim 1, wherein
Cu <4.20, Mg <1.38, Mn <0.42, Zn ≥ (1.2 Cu - 0.3 Mg + 0.3 Mn-3.75). Product according to any one of Claims 1 to 8, characterized in that it has been dissolved, quenched and cold-worked with a permanent deformation of between 0.5% and 15%, preferably between 1% and 5%. %, and even more preferably between 1.5% and 3.5%. Product according to any one of claims 1 to 9, characterized in that said product is a sheet with a thickness of between 1 and 16 mm. Product according to any one of claims 1 to 10, characterized in that said sheet is a plate plated with at least one face with an alloy of the series 1xxx, and preferably with an alloy selected from the group consisting of alloys 1050 , 1070, 1300 and 1145. Product according to any one of claims 1 to 11, characterized in that its resistance to rupture in the direction L and / or TL direction is greater than 430 MPa, and preferably greater than 440 MPa. Product according to any one of claims 1 to 12, characterized in that its yield strength L and / or TL direction is greater than 300 MPa, and preferably greater than 320 MPa. Product according to any one of Claims 1 to 13, characterized in that its elongation at break in direction L and / or direction TL is greater than 19% and preferably greater than 20%. Product according to any one of claims 1 to 14, characterized in that its damage tolerance Kr, calculated from a curve R obtained according to ASTM E 561 for a value Δ a _eff of 60 mm, is greater than 165 MPa √m in the TL and LT directions. Product according to any one of claims 1 to 15, characterized in that its damage tolerance Kr, calculated from a curve R obtained according to ASTM E 561 for a value Δ a _eff of 60 mm, is greater than 180 MPa √m in the LT direction. Product according to any one of claims 1 to 16, characterized in that its crack propagation rate da / dn, determined according to ASTM E 647 in the TL or LT direction for a load ratio R = 0.1 and a ΔK value of 50 MPa√m is less than 2.5 × 10 ^-2 mm / cycle, and preferably less than 2.0 ^× 10 ^-2 mm / cycle. Plated sheet according to one of claims 1 to 17, characterized in that the galvanic corrosion current is less than 4μA / cm ² for exposure of up to 200 hours during corrosion tests in a riveted assembly, placing the core alloy in an aerated solution containing 0.06 M NaCl and the plating alloy in a 0.02 M solution of AlCl ₃ deaerated by nitrogen sparging. Plated sheet according to claim 18, characterized in that said galvanic corrosion current is less than 2.5μA / cm ² . Aircraft structure element made from at least one product according to one of any of claims 1 to 19. Structural element according to claim 20, characterized in that said structural element is a fuselage skin element. A method of manufacturing a wrought product according to one of claims 1 to 19, comprising the following steps: (a) pouring a plate or billet, (b) homogenization between 450 ° C and 500 ° C, (c) hot transformation by spinning, rolling or forging, (d) optionally a cold transformation, (e) dissolving between 480 ° C and 505 ° C, (f) quenching, (g) cold working resulting in a permanent deformation of between 0.5% and 15%. The process according to claim 22, wherein the work hardening is carried out to lead to a permanent deformation of between 1 and 5%, and preferably between 1.5 and 3.5%.