EP2017361A1 - Aluminium-Kupfer-Lithium-Blech mit hoher Zähigkeit für Flugzeugrumpf - Google Patents

Aluminium-Kupfer-Lithium-Blech mit hoher Zähigkeit für Flugzeugrumpf Download PDF

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EP2017361A1
EP2017361A1 EP08018130A EP08018130A EP2017361A1 EP 2017361 A1 EP2017361 A1 EP 2017361A1 EP 08018130 A EP08018130 A EP 08018130A EP 08018130 A EP08018130 A EP 08018130A EP 2017361 A1 EP2017361 A1 EP 2017361A1
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Prior art keywords
weight
mpa
sheet
toughness
samples
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EP08018130A
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English (en)
French (fr)
Inventor
Bernard Bes
Hervé Ribes
Christophe Sigli
Timothy Warner
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Constellium Issoire SAS
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Alcan Rhenalu SAS
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Priority claimed from FR0508374A external-priority patent/FR2889542B1/fr
Application filed by Alcan Rhenalu SAS filed Critical Alcan Rhenalu SAS
Publication of EP2017361A1 publication Critical patent/EP2017361A1/de
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    • 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
    • C22C21/12—Alloys based on aluminium with copper as the next major constituent
    • 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
    • C22F1/057—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 copper as the next major constituent

Definitions

  • the present invention generally relates to aluminum alloy products and, in particular, such products useful in the aerospace industry and suitable for use in fuselage applications.
  • compressive and compressive shear strength is an extremely important design guideline, since the heavier fuselage panels suffer this type of constraint.
  • it In order for a new material to be able to reduce the weight of these compression-stressed panels, it must have a high modulus of elasticity, a high 0.2% yield strength (to withstand buckling) and a low mass volume.
  • the second major guideline is the residual resistance of panels longitudinally (in the axis of the fuselage) cracked.
  • Aeronautical certification regulations require the consideration of damage tolerance in design, so it is usual to consider large longitudinal or circumferential cracks in the fuselage panels, to demonstrate that a certain level of stress can be applied. without a catastrophic break.
  • a known property of the materials governing the design here is the toughness under plane stress. All known factors of critical stress intensity, however, only give a limited view of toughness.
  • the R curve test is a widely recognized means for characterizing toughness properties.
  • the curve R represents the evolution of the critical effective stress intensity factor for the crack propagation as a function of the effective crack extension under a monotonic stress. It allows the determination of the critical load for unstable failure for any configuration relevant to cracked aircraft structures.
  • the values of the effective stress intensity factor and the crack extension effective are values defined in ASTM E561.
  • the length of the curve R - namely the maximum crack extension of the curve - is a parameter in itself important for the fuselage design.
  • K app The classical analysis, generally used, of the tests carried out on panels with central crack gives a factor of intensity of stress apparent to the rupture (K app ). This value does not vary significantly with the length of the R curve, especially when the slope of the R curve is close to the slope of the stress intensity factor curve applied to the crack length (applied curve) .
  • the applied curve drops due to the bridging effect of the stiffener.
  • a lower density is clearly beneficial for the weight of a structural member.
  • a third major guideline is thus the density of the material.
  • large parts of the fuselage are not so heavily loaded and the weight of the design is limited by a certain limit generally called "minimum thickness".
  • the minimum thickness concept is the lowest usable thickness for manufacturing (especially panel handling) and repair (repair riveting). The only way to reduce the weight in this case is to use a lower density material.
  • the fuselages of civil aircraft are, for the most part, made of alloy sheet 2024, 2056, 2524, 6013, 6156 or 7475, plated on each side with an aluminum alloy lightly loaded with alloying elements. , an alloy 1050 or 1070 for example.
  • the purpose of the coating alloy is to impart sufficient corrosion resistance. Light, generalized or pitting corrosion is tolerable but must not be penetrating so as not to attack the core alloy. There is a tendency to try to use non-plated materials for fuselage design, so as to reduce the cost. Corrosion resistance, and in particular intergranular corrosion and corrosion under stress, the fuselage panel is thus an important aspect of its properties.
  • the patent US 5,032,359 discloses a family of alloys based on aluminum-copper-magnesium-silver alloys to which lithium has been added, in specific ranges and which have high resistance at ambient temperature and at high temperature, high ductility at ambient temperatures and at high temperature, extrusionability, forgeability, and good solderability and natural aging response properties.
  • the examples describe extruded products. No information is provided on toughness, fatigue behavior or corrosion resistance.
  • the alloy has a composition of 3.0 to 6.5% copper, 0.05 to 2.0% magnesium, 0.05 to 1.2% silver, from 0.2 to 3.1% lithium, from 0.05 to 0.5% of an element chosen from zirconium, chromium, manganese, titanium, boron, hafnium, vanadium, titanium diboride and mixtures thereof.
  • the document US 5,211,910 discloses aluminum-based alloys containing Cu, Li, Zn, Mg and Ag which have favorable properties, such as relatively low density, high modulus, high mechanical strength / ductility combinations , a strong response to natural aging with and without anterior work hardening, and a high modulus after income with or without prior work hardening.
  • the alloys have a composition of 1 to 7% Cu, 0.1 to 4% Li, 0.01 to 4% Zn, 0.05 to 3% Mg, 0.01 to 2% of Ag, from 0.01 to 2% of an element selected from Zr, Cr, Mn, Ti, Hf, V, Nb, B and TiB 2 , the remainder being Al together with its unavoidable impurities.
  • This invention describes how Zn additions can be used to reduce the Ag content present in the alloys taught in the document. US 5,032,359 in order to reduce the cost.
  • the document US 5,455,003 discloses a process for producing aluminum-copper-lithium alloys that exhibit improved strength and toughness at cryogenic temperatures.
  • the improved cryogenic properties are achieved by adjusting the composition of the alloy, along with the processing parameters such as the amount of work hardening and the income.
  • the product is used for cryogenic tanks in space launch vehicles.
  • the document US 5,389,165 discloses an aluminum alloy useful in aircraft and aerospace structures which has low density, high mechanical strength and high toughness and has the formula: Cu a Li b Mg c Ag d Zr e Al bal wherein a, b, c, d, e and bal indicate the amount in% by weight of alloying components, and wherein 2.8 ⁇ a ⁇ 3.8, 0.80 ⁇ b ⁇ 1.3, 0.20 ⁇ c ⁇ 1.00, 0.20 ⁇ d ⁇ 1.00 and 0.08 ⁇ e ⁇ 0.40.
  • the copper and lithium components are adjusted so that the combined copper and lithium content is kept below the solubility limit in order to avoid a loss of toughness during high temperature exposure.
  • the relationship between copper and lithium grades must also satisfy the following relationship: Cu ( % in weight ) + 1 , 5 Li ( % in weight ) ⁇ 5 , 4.
  • Al-Cu-Mg alloy comprising from 3 to 5% by weight of Cu, from 0.5 to 2% by weight of Mg and from 0.01 to 0.9% by weight of Li.
  • patent application the toughness of alloys for which an addition of Li between 0.2 and 0.7% by weight is significantly improved over similar alloys containing either no Li or a higher amount of Li.
  • the present inventors have come to the present invention concerning an aluminum-copper-lithium-magnesium-silver alloy, which exhibits high mechanical strength, high toughness and specifically high crack extension prior to fracture. unstable pre-cracked wide panels, and a high resistance to corrosion.
  • Another subject of the invention is a laminated, extruded and / or forged aluminum alloy product comprising 2.7 to 3.4% by weight of Cu, 0.8 to 1.4% by weight of Li, 0 , 1 to 0.8% by weight of Ag, 0.2 to 0.6% by weight of Mg and at least one element selected from Zr, Mn, Cr, Sc, Hf and Ti, the amount of said element, it is selected from 0.05 to 0.13% by weight for Zr, 0.05 to 0.8% by weight for Mn, 0.05 to 0.3% by weight for Cr and for Sc, 0 0.5 to 0.5 wt% for Hf and 0.05 to 0.15 wt% for Ti, the remainder being aluminum and unavoidable impurities, with the additional requirement that the amount of Cu and Li either Cu ( % in weight ) + 5 / 3 Li ( % in weight ) ⁇ 5 , 2.
  • Still other objects of the invention are elements of structures, stiffeners and fuselage panels obtained from said rolled, extruded and / or forged products.
  • the static mechanical characteristics in other words the ultimate tensile strength Rm, the conventional yield stress at 0.2% elongation R p0.2 and the elongation at break A, are determined by a tensile test according to the EN 10002-1, the sampling and the sense of the test being defined by EN-485-1.
  • the cracking rate (da / dN) is determined according to the ASTM E 647 standard.
  • the critical stress intensity factor K C in other words the intensity factor that makes the crack unstable, is calculated from the curve R.
  • the stress intensity factor K CO is also calculated by assigning the initial crack length at the critical load at the beginning of the monotonic load. These two values are calculated for a specimen of the required form.
  • K app represents the K CO factor corresponding to the specimen that was used to perform the R curve test.
  • K eff represents the K C factor corresponding to the specimen that was used to perform the R curve test.
  • ⁇ a eff (max) represents the crack extension of the last valid point of the R curve.
  • the crack size at the end of the pre-fatigue cracking stage is W / 3 for M-type specimens ( T), wherein W is the width of the specimen as defined in ASTM E561.
  • W is the width of the specimen as defined in ASTM E561.
  • the width of the specimen used in an R curve test can have a substantial influence on the stress intensity measured in the test.
  • the fuselage sheets being large panels, the results of curve R obtained on sufficiently large samples, such as samples having a width greater than or equal to 400 mm, are judged the most significant for the evaluation of toughness. For this reason, CCT760 test specimens, which had a width of 760 mm, were used preferentially for the evaluation of toughness.
  • the initial crack length 2ao 253 mm.
  • the toughness was also evaluated in the TL directions using the global energy at break E g according to the Kahn test.
  • the stress Kahn R e (in MPa) 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). R e does not make it possible to evaluate the relative toughness of samples whose static mechanical characteristics are different.
  • the overall energy at break E g is determined as the area under the Force-Displacement curve until the test piece breaks, E g is directly related to toughness.
  • the test is described in the article Kahn-Type Tear Test and Crack Toughness of Aluminum Alloy Sheet, published in Materials Research & Standards, April 1964, p. 151-155 .
  • the test specimen used for the Kahn toughness test is described, for example, in the "Metals Handbook", 8th Edition, Vol. 1, American Society for Metals, pp. 241-242 .
  • sheet is meant here a rolled product not exceeding 12 mm thick.
  • structural element refers to an element used in mechanical engineering for which the static and / or dynamic mechanical characteristics are of particular importance for the performance and integrity of the structure, and for which a calculation of the structure is usually prescribed or performed. It is typically a mechanical part whose failure is likely to endanger the safety of said construction, its users, its users or others.
  • these structural elements include the elements that make up the fuselage (such as fuselage skin (fuselage skin in English), stiffeners or stringers, bulkheads, fuselage (circumferential frames), wings (such as wing skin), stiffeners (stiffeners), ribs (ribs) and spars) and empennage including horizontal stabilizers and vertical stabilizers horizontal or vertical stabilizers, as well as floor beams, seat tracks and doors.
  • fuselage such as fuselage skin (fuselage skin in English
  • stiffeners or stringers such as fuselage skin
  • bulkheads fuselage (circumferential frames)
  • wings such as wing skin
  • stiffeners stiffeners (stiffeners), ribs (ribs) and spars
  • empennage including horizontal stabilizers and vertical stabilizers horizontal or vertical stabilizers, as well as floor beams, seat tracks and doors.
  • the aluminum-copper-lithium-silver-magnesium alloy according to one embodiment of the invention advantageously has the following composition: ⁇ u> Table 1 ⁇ / u>: Alloy composition ranges (% by weight, the balance being Al) Cu Li Ag mg Large 2.7 to 3.4 0.8 to 1.4 0.1 to 0.8 0.2 to 0.6 favorite 3.0 to 3.4 0.8 to 1.2 0.2 to 0.5 0.2 to 0.6 Most preferred 3.1 to 3.3 0.9 to 1.1 0.2 to 0.4 0.2 to 0.4
  • the relationship between copper and lithium is preferably: Cu % in weight + 5 / 3 ⁇ Li % in weight ⁇ 5.
  • At least one element such as Zr, Mn, Cr, Sc, Hf, Ti or a combination thereof is included to refine the grain.
  • the additions depend on the element: from 0.05 to 0.13% by weight (preferably from 0.09 to 0.13% by weight) for Zr, from 0.05 to 0.8% by weight for Mn from 0.05 to 0.3% by weight for Cr and Sc, from 0.05 to 0.5% by weight for Hf and from 0.05 to 0.15% by weight for Ti.
  • the sum can be limited by the appearance of primary phases.
  • grain refining is achieved by the addition of 0.05 to 0.13% by weight of Zr, from 0.02 to 0.3% by weight of Sc and optionally from 0.05 to 0.8% in Mn weight, 0.05 to 0.3% by weight of Cr, 0.05 to 0.5% by weight of Hf and 0.05 to 0.15% by weight of Ti.
  • Mn content In some cases, and particularly for hot-rolled sheets with a thickness of between 4 and 12 mm, it may be advantageous to limit the Mn content to 0.05% by weight and preferably to 0.03% by weight. The inventors have observed that for such thicknesses the presence of Mn makes it more difficult to control the granular structure and may affect both the mechanical properties and the toughness.
  • Fe and Si generally affect toughness properties.
  • the amount of Fe should preferably be limited to 0.1% by weight and the amount of Si should preferably be limited to 0.1% by weight (preferably to 0.05% by weight). All other elements should also preferably be limited to 0.1% by weight (preferably to 0.05% by weight).
  • the inventors have found that if the copper content is greater than 3.4% by weight, the toughness properties can in some cases fall rapidly. For certain embodiments of the invention, it is recommended not to exceed a copper content of 3.3% by weight. Preferably, the copper content is greater than 3.0% or even 3.1% by weight.
  • the present inventors have found that Zr contents greater than 0.13% by weight can, in some cases, lead to a lower toughness performance. Whatever the reason for this drop in toughness, the inventors found that the higher Zr content led to a formation of Al 3 Zr primary phases. In this case, a temperature of High casting may be used to avoid formation of the primary phases, but this may lead to lower liquid metal quality, in terms of inclusion and gas content. This is why the present inventors consider that the Zr should advantageously not exceed 0.13% by weight.
  • the inventors have found that if the Li content is less than 0.8% by weight or even 0.9% by weight, the improvement in mechanical strength is too low. In some cases, it may be advantageous if the Li content is> 0.9% by weight. Also, with these low Li content, the decrease in the density of the alloy is too low. For a Li content greater than 1.4% or more than 1.2% by weight or even greater than 1.1% by weight, the toughness is significantly reduced. Also, these high Li levels have several disadvantages related in particular to the thermal stability, flowability and cost of raw materials.
  • the addition of Ag is an essential feature of the invention.
  • the strength and toughness performances observed by the inventors are not usually achieved for alloys containing no silver.
  • the inventors believe that silver plays a role in the formation of copper-containing hardening phases, formed during natural or artificial aging and in particular allows the formation of finer phases and a more homogeneous distribution of these phases.
  • the advantageous effect of Ag is observed for a content of this element greater than 0.1% by weight and preferably greater than 0.2% by weight. In order to limit the cost associated with the addition of Ag, it may be advantageous not to exceed 0.5% by weight or even 0.4% by weight.
  • Mg improves the mechanical strength and decreases the density.
  • An excessive addition of Mg would have a detrimental effect on toughness.
  • the Mg content is limited to 0.4% by weight. The inventors believe that the addition of Mg could also have a role during the formation of copper-containing phases.
  • the bath of liquid metal having a composition according to the invention is then cast.
  • the present invention makes it possible to obtain a laminated, extruded and / or forged product whose thickness is, advantageously, between 0.8 and 12 mm and preferably between 2 and 12 mm.
  • an alloy having adjusted amounts of alloying elements is cast as a plate.
  • the plate is then homogenized at 490-530 ° C for 5-60 hours.
  • the inventors have observed that homogenization temperatures above 530 ° C can tend to reduce the tenacity performance in some cases.
  • the plates are heated at 490 to 530 ° C for 5 to 30 hours. Hot rolling is carried out to obtain a thickness of between 4 and 12 mm. For a thickness of approximately 4 mm or less, a cold rolling step may be added, if necessary.
  • the sheet obtained at a thickness preferably between 0.8 and 12 mm, and the invention is more advantageous for sheets of 2 to 12 mm thick and even 2 to 9 mm and even more advantageous for sheets of 3 to 7 mm thick.
  • the sheets are then put in solution, for example by heat treatment between 490 and 530 ° C for 15 min to 2 h, and then quenched with water at room temperature or preferably cold water.
  • the product then undergoes a controlled pull of 1 to 5% and preferably 2.5 to 4%.
  • Such cold hardening levels can also be achieved by cold rolling, planing, forging or a combination of these methods and controlled pulling.
  • the total cold working after quenching is between 2.5 and 4%.
  • the controlled tensile deformation may be between 1.7 and 3.5. %.
  • the inventors have observed that the tenacity tends to decrease when the controlled tensile deformation is greater than 5%.
  • the results of Kahn test, in particular E g tends to decrease for permanent deformations greater than 5%. It is therefore recommended not to exceed a permanent deformation of 5%.
  • the traction is greater than 5%, there may be industrial difficulties such as high implementation as well as that formatting difficulties, which would increase the cost of the product.
  • An income is achieved at a temperature between 140 and 170 ° C for 5 to 30 hours, which provides a T8 state.
  • the reaction is more preferably carried out between 140 and 155 ° C for 10 to 30 hours.
  • Low tempering temperatures generally favor high toughness.
  • the revenue step is divided into two steps: a pre-revenue step prior to a welding operation, and a final heat treatment of a welded structure member.
  • friction stir welding is a preferred welding technique.
  • the sheets according to the invention have advantageous properties for recrystallized, non-recrystallized or mixed microstructures (that is to say comprising recrystallized zones and non-recrystallized zones).
  • the inventors have observed that it could be advantageous to avoid mixed microstructures: for sheets whose thickness is between 4 and 12 mm, it may be advantageous for the microstructure to be completely uncrystallized.
  • the forming of the sheet of the invention may advantageously be carried out by deep drawing, stretching, spinning, rolling or folding, these techniques being known to those skilled in the art.
  • a structural element formed of at least one product according to the invention, in particular a sheet according to the invention and stiffeners or frames, these stiffeners or frames being preferably made of extruded profiles, can be used in particular for the manufacture of aircraft fuselage panels as well as any other use where the present properties could be advantageous.
  • structural members, stiffeners, and / or fuselage panels can be made from the rolled, extruded, and / or forged products obtained.
  • the inventors have found that the sheet of the invention has mechanical properties Particularly favorable statics and high tenacity.
  • the high-tenacity sheets generally have low yield strengths and breaking strength.
  • the high mechanical properties favor an industrial application for aircraft structural parts, the elastic limit and the breaking strength of said sheet being characteristics which are directly taken into account for the calculation. structural dimensioning.
  • Calculations of structural elements and in particular of fuselage panels comprising sheets and / or stiffeners according to the invention have shown a possibility of weight reduction with respect to structural elements of comparable properties comprising only metal sheets. prior art alloy 2024, 2056, 2098, 7475 or 6156. Such weight reductions are generally from 1 to 10% and in some cases even greater weight reductions can be achieved.
  • the simple substitution of the alloy 2024 with an alloy according to the invention may allow a weight reduction of the order of 3 to 3.5%.
  • the high mechanical properties of the alloys according to the invention make it possible to develop products of a lighter size and shape, which makes it possible to reach or even exceed a weight reduction of 10%.
  • the sheet of the invention does not generally induce any particular problem during subsequent surface treatment operations conventionally used in aircraft construction.
  • the resistance to intergranular corrosion of the sheet of the invention is generally high; for example, only pits are generally detected when the metal is subjected to a corrosion test.
  • the sheet of the invention can be used without plating.
  • the density of the various alloys tested is shown in Table 3.
  • the samples F to I have the lowest density of the various materials tested.
  • Table 3 ⁇ / u>: Density of tested alloys Reference Density (g / cm 3 ) A (2024) 2.78 B (2056) 2.78 C (7475) 2.81 D (6156) 2.72 E (2098) 2.70 F, G, H, I, J, K 2, 69
  • Table 5 provides the reference of the different samples and their dimensions. ⁇ u> Table 5 ⁇ / u>: Final dimensions of the samples Sample Thickness [mm] Width [mm] Length [mm] AT 6.0 2,000 3000 B 6.0 2,000 3000 VS 6.3 1,900 4000 D 4.6 2,500 4,500 E # 1 2.0 1,000 2,500 E # 2 3.2 1,000 2,500 E # 3 4.5 1,250 2,500 E # 31 4.5 1,250 2,500 E # 4 6.7 1,250 2,500 F # 1 3.0 1,000 2,500 F # 2 5.0 1,250 2,500 F # 3 6.7 1,250 2,500 G # 1 3.8 2,450 9,600 H # 1 5.0 2,450 9,600 I # 1 5.0 1,500 3000 K # 1 2.0 1,000 2,500
  • the static mechanical properties of the samples according to the invention are very high compared to the conventional alloy of the 2XXX range which is tolerant to damage, and of the same order of magnitude as the sample 7475 T76 referenced C.
  • the mechanical strength of the samples according to the invention considers that the lower copper content and the lower zirconium content of the samples according to the invention have a slight influence on their mechanical strength.
  • the curves R of certain samples according to the invention and reference samples E are provided on the figures 1 and 2 , for the TL and LT directions, respectively.
  • the figure 1 clearly shows that the crack extension of the last valid point of the curve R ( ⁇ a eff (max) ) is much greater for the samples of the invention than for the sample E # 1, E # 3, E # 31 and E # 4.
  • This parameter is at least as critical as the K app values because, as explained in the description of the prior art, the length of the curve R is an important parameter for the design of the fuselage.
  • the figure 2 shows the same trend, although the LT leadership inherently gives a better result.
  • the curve R of the sample F # 3 could not be measured in the direction LT because the maximum load of the machine was reached. Table 7 summarizes the results of the toughness tests.
  • the value of K app in the TL direction is greater than 110 MPa. m and even greater than 130 MPa m while for the reference alloy samples 2098, the value of K app in the TL direction is less than 110 MPa m except for sample E # 3 which has undergone a special annealing step before dissolution.
  • the results from the curve R are grouped together in Table 8.
  • the crack extension of the last valid point of the curve R is greater for the samples of the invention than for the reference samples.
  • all the samples according to the invention reach a crack extension of at least 30 mm and even at least 40 mm while the maximum crack extension is less than 40 mm for the samples of the invention. reference.
  • the inventors consider that several reasons can be proposed to explain this performance, such as the lowest Cu content, and / or the lowest Zr content.
  • the figures 3 and 4 show the evolution of the cracking rate da / dN (in mm / cycle) in the TL and LT orientation, respectively, for different levels of stress intensity factor ( ⁇ K).
  • the cracking rate of the sample F is in the same range as that typically obtained for alloy 2056 (Sample B) and lower than that obtained for alloy 6156 (Sample D).
  • the intergranular corrosion resistance was tested according to ASTM G110. For all the samples according to the invention, no intergranular corrosion was detected. No intergranular corrosion was either detected on the 2098 alloy reference samples (E # 1 to E # 4). For Sample B (for which plating had been removed), intergranular corrosion with an average depth of 120 ⁇ m was observed and for Sample D (for which plating had been removed) intergranular corrosion was observed. with an average depth of 180 ⁇ m. The resistance to intergranular corrosion was thus very high for the samples according to the invention.
  • the income was made either before or after assembly by friction stir welding.
  • the results are given in Table 13.
  • the performance of the welded joints obtained with the sheets according to the invention was particularly satisfactory for two aspects.
  • the joint efficiency coefficient which is the ratio between the breaking strength of the welded joint and that of the non-welded sheet, is greater than 70% and even greater than 75% for the sheets of the invention. This coefficient reaches 80% in some cases. This result is better than that obtained with sheets from casting E.
  • the results are little influenced by the position of the stage of income (before or after welding), which allows a flexible process. On the contrary, for the sheets obtained from the casting D (6156), a significant influence of the position of the income stage is observed.
  • Samples L and M reach the mechanical characteristics according to the invention in the T8 state. Furthermore, the static strength and toughness performances are lower for the sample L, which contains Mn and a low Zr content, than for the other examples according to the invention. The inventors believe that the lower performance of the sample L is related to a less favorable microstructure characterized in particular by the presence of recrystallized zones and non-recrystallized zones (mixed microstructure).

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EP08018130A 2005-06-06 2006-06-02 Aluminium-Kupfer-Lithium-Blech mit hoher Zähigkeit für Flugzeugrumpf Withdrawn EP2017361A1 (de)

Applications Claiming Priority (3)

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US68744405P 2005-06-06 2005-06-06
FR0508374A FR2889542B1 (fr) 2005-08-05 2005-08-05 Tole en aluminium-cuivre-lithium a haute tenacite pour fuselage d'avion
EP06764718A EP1891247B1 (de) 2005-06-06 2006-06-02 Hochfestes aluminium-kupfer-lithium-blech für flugzeugrümpfe

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AT (1) ATE414183T1 (de)
BR (1) BRPI0610937B1 (de)
CA (1) CA2608971C (de)
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FR3004464B1 (fr) * 2013-04-12 2015-03-27 Constellium France Procede de transformation de toles en alliage al-cu-li ameliorant la formabilite et la resistance a la corrosion
FR3014448B1 (fr) 2013-12-05 2016-04-15 Constellium France Produit en alliage aluminium-cuivre-lithium pour element d'intrados a proprietes ameliorees
FR3014905B1 (fr) * 2013-12-13 2015-12-11 Constellium France Produits en alliage d'aluminium-cuivre-lithium a proprietes en fatigue ameliorees
FR3026747B1 (fr) * 2014-10-03 2016-11-04 Constellium France Toles isotropes en alliage d'aluminium-cuivre-lithium pour la fabrication de fuselages d'avion
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FR3075078B1 (fr) * 2017-12-20 2020-11-13 Constellium Issoire Procede de fabrication ameliore de toles en alliage d'aluminium-cuivre-lithium pour la fabrication de fuselage d'avion
FR3082210B1 (fr) 2018-06-08 2020-06-05 Constellium Issoire Toles minces en alliage d’aluminium-cuivre-lithium pour la fabrication de fuselages d’avion
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DE602006003656D1 (de) 2008-12-24
BRPI0610937A2 (pt) 2010-08-03
ATE414183T1 (de) 2008-11-15
EP1891247A1 (de) 2008-02-27
RU2007145191A (ru) 2009-06-10
WO2006131627A1 (fr) 2006-12-14
CA2608971A1 (fr) 2006-12-14
EP1891247B1 (de) 2008-11-12
BRPI0610937B1 (pt) 2015-12-08
ES2314929T3 (es) 2009-03-16
RU2415960C2 (ru) 2011-04-10

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