EP1114877B1 - Al-Cu-Mg alloy aircraft structural element - Google Patents

Al-Cu-Mg alloy aircraft structural element Download PDF

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Publication number
EP1114877B1
EP1114877B1 EP00420263A EP00420263A EP1114877B1 EP 1114877 B1 EP1114877 B1 EP 1114877B1 EP 00420263 A EP00420263 A EP 00420263A EP 00420263 A EP00420263 A EP 00420263A EP 1114877 B1 EP1114877 B1 EP 1114877B1
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Prior art keywords
structure element
element according
temperature
product
alloy
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German (de)
French (fr)
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EP1114877A1 (en
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Timothy Warner
Philippe Lassince
Philippe Leqeu
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Constellium Issoire SAS
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Pechiney Rhenalu SAS
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Classifications

    • 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
    • C22C21/16—Alloys based on aluminium with copper as the next major constituent with magnesium
    • 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
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10S—TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10S428/00—Stock material or miscellaneous articles
    • Y10S428/922—Static electricity metal bleed-off metallic stock
    • Y10S428/923—Physical dimension
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12229—Intermediate article [e.g., blank, etc.]
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T428/00—Stock material or miscellaneous articles
    • Y10T428/12—All metal or with adjacent metals
    • Y10T428/12493—Composite; i.e., plural, adjacent, spatially distinct metal components [e.g., layers, joint, etc.]
    • Y10T428/12736—Al-base component
    • Y10T428/12764—Next to Al-base component

Definitions

  • the invention relates to aircraft structural elements, in particular skin and underside sail stiffeners for large commercial aircraft capacity, made from rolled, extruded or forged AlCuMg alloy the state treated by dissolution, quenching and tempering, and presenting, with respect to prior art products used for the same application, an improved compromise between the different job properties required.
  • Large commercial aircraft wings have an upper section (or extrados) consisting of a skin made from thick alloy plates 7150 to the T651 state, or 7055 alloy to the T7751 state or 7449 to the T7951 state, and stiffeners made from profiles of the same alloy, and a lower part (or intrados) consisting of a skin made from thick 2024 alloy state T351 or 2324 in state T39, and stiffeners made from same alloy. Both parts are assembled by longitudinal members and ribs.
  • the alloy 2024 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
  • US Patent 5652063 (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, 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.
  • the alloy may also contain: Zr ⁇ 0.20% V ⁇ 0.20% Mn ⁇ 0.80% Ti ⁇ 0.05% Fe ⁇ 0.15% If ⁇ 0.10%
  • the current trend is to use growing of very thick products, in the bulk of which structural elements are machined.
  • the skins of wing are machined from relatively thick sheets to allow machining in the mass of wing stiffeners, whereas these are usually made from profiles or folded sheets, and are then mechanically fixed to the skin.
  • the integral machining in the mass of the skin-stiffener assembly leads to a reduction in manufacturing costs, since the number of parts is reduced and we avoid assembly.
  • the use of an unassembled structure allows a reduction of the weight of the whole.
  • the sheets have homogeneous mechanical properties over the entire thickness, that is, the properties do not vary so significant depending on the thickness, typically between 10 and 120 mm.
  • the more machining is used the greater the stability to machining is desirable, which is obtained by a low level of internal constraints.
  • the mechanical characteristics are all the more homogeneous, and the internal constraints all the more reduced, as the sheet has a low sensitivity to the quenching.
  • aircraft wings especially large aircraft, have a curved wing profile, with a curvature both in the longitudinal direction and in the transversal direction.
  • This complex shape can be obtained during the operation of returned to an autoclave, by forming on a mold, by depressing the face the side of the mold with respect to the opposite side, using a partial vacuum. he It is imperative that this operation be successful, to avoid the costly waste of a part high added value, especially for large parts.
  • the pledge of success lies in the lowest possible springback for a form of mold given, because the springback is usually difficult to control.
  • the object of the present invention is to provide aircraft structural elements having properties at least equivalent to those of the same elements made alloy 2024 in the T351 state with respect to mechanical characteristics static, toughness, speed of crack propagation and resistance to corrosion, using rolled, spun or forged products with a low level of residual stresses, low sensitivity to quenching and good ability to train for income.
  • the subject of the invention is an aircraft structural element, in particular an aircraft wing surface element, made from a rolled or forged product, made of an alloy of composition (% by weight): Cu: 4.6 - 5.3 Mg: 0.10 - 0.50 Mn: 0.25 - 0.45 Si ⁇ 0.10 Fe ⁇ 0.15 Zn ⁇ 0.20 Cr ⁇ 0.10 other elements ⁇ 0.05 each and ⁇ 0.15 in total, remains Al treated by dissolution, quenching, controlled tensile stress relieving to more than 1.5% of permanent deformation and income.
  • the inlet temperature to hot rolling is preferably at least 40.degree. C., and more preferably at least 40.degree. 50 ° C, at the dissolution temperature.
  • the invention is based on the finding that a 2001 type alloy, with certain changes in composition and an appropriate manufacturing range, could present a set of properties making it suitable for use in structures of aircraft, and more particularly in the lower surface of the wings of commercial aircraft large capacity, with more interesting properties in terms of low quenching sensitivity, low residual stresses and income shaping.
  • the copper content range is significantly shifted towards the low, while remaining higher than that of the 2024 or 2034 alloys for intrados, for compensate, in its influence on the mechanical strength, the low magnesium. It is better to choose a copper content higher than 4.8% or at 4.9% or even 5%.
  • the magnesium content is of the same order as in the alloy 2001, and preferably between 0.20 and 0.40%.
  • the ratio Cu / Mg is thus almost always above 10, contrary to the teaching of the US patent 5652063, which recommends a Cu / Mg ratio of between 5 and 9.
  • the manganese content is controlled in a relatively narrow range. Below 0.15%, you could have a grain too big; above 0.45%, we get a non recrystallized structure which is not favorable to the control of the constraints residual. A preferred range is from 0.25 to 0.40%. Note that, for the same reason, and contrary to the teaching of US Patent 5593516, the alloy contains no other anti-recrystallizing element such as vanadium or zirconium.
  • the iron and silicon contents are maintained below 0.15, respectively. and 0.10%, and preferably below 0.09 and 0.08%, to ensure good tenacity.
  • the alloy may comprise up to 0.2% zinc, this addition having an effect favorable on the mechanical strength, without risk for other properties, such as corrosion resistance.
  • the transformation range includes the casting of a veneer or billet, a reheating or homogenization at a temperature close to the temperature of beginning fusion of the alloy and a hot transformation by rolling, spinning or forging.
  • rolling it may include a pass, called enlargement, in the direction perpendicular to that of the other passes, and intended for improve the isotropy of the product.
  • the hot transformation temperature is, preferably at a slightly lower level than that which would be profession with reference to the solution temperature. So, as far as the rolling, the inlet temperature is preferably at least 40 ° C or 50 ° C, below the solution temperature, and the outlet temperature of 20 to 30 ° C below the inlet temperature.
  • the product is then submitted to a dissolution as complete as possible, to a temperature close to, for example less than 10 ° C below, the temperature of beginning fusion of the alloy, while avoiding the burn. This temperature is between 520 and 535 ° C.
  • the quality of the dissolution in solution can be controlled by analysis differential enthalpy.
  • the product is then quenched, for example by immersion in cold water, so as to ensure a cooling rate of between 10 and 50 ° C / s. After quenching, the product is triturated until deformation at least 1.5%, so as to relax it and improve its flatness.
  • this traction also has the effect of improving, by a hardening effect, the elasticity limit after income, so that one can qualify the state obtained from state T851, as if it were a specific pass hardening after quenching.
  • the income itself can at the same time as the shaping of the curve of the intrados.
  • This income is preferably at a temperature greater than 160 ° C (and higher preferentially> 170 ° C.), of a duration enabling the limit peak to be reached elasticity, as for a T6 state.
  • a time income equivalent to that corresponding to 12 to 24 h at a temperature of 173 ° C is carried out; all time-temperature combination allowing to reach the peak of income of the alloy is usable.
  • the metallurgical structure obtained is, unlike that of alloys 2024 and 2034, strongly recrystallized, with a recrystallization rate still exceeding 70%, and the more often 90%, over the entire thickness.
  • the structural elements according to the invention have a compromise of properties (static mechanical characteristics, toughness, speed of crack propagation, corrosion resistance) that make them suitable for use in construction aeronautics, and in particular to the manufacture of wing bottoms.
  • properties static mechanical characteristics, toughness, speed of crack propagation, corrosion resistance
  • these elements can be easily made by machining and trained to income.
  • the alloy used is easily soldered by the usual techniques, which can reduce the number of riveted assemblies.
  • alloys have been prepared, the composition of which is indicated in Table 1.
  • the alloy A is a 2024-T3 alloy of the usual composition for the intrados application of the wing.
  • Alloy B is an alloy falling within the composition range described in US Patent No. 5652063, but without the addition of silver.
  • Alloy C is in accordance with the invention.
  • Alloys D and E differ from alloy C only by higher silicon for D, higher manganese and copper for E and F, and zirconium addition for F.
  • Cast plates with a cross section of 380 ⁇ 120 mm were homogenized, hot-rolled to a thickness of 22 mm, put into solution, quenched with cold water, triturated at 2.3% permanent deformation and returned.
  • the parameters of homogenization, hot rolling (inlet temperatures), dissolution and tempering are given in Table 2.
  • the toughness was also measured by the critical stress intensity factor K 1c (in MPa ⁇ m) measured, according to ASTM E 399, on CT20 specimens taken at quarter-thickness in the LT and TL directions (2 test pieces by case).
  • the alloy C according to the invention leads to a limit of elasticity significantly higher than that of 2024, and slightly lower than that of alloys B, E and F. Elongation is lower than for 2024, but better than that of alloys B, D, E and F. Toughness is the best of all the alloys tested. So we have a favorable compromise of these various properties. In particular, the results show the adverse effect, both on toughness and elongation, of increase in the silicon and manganese content, as well as an addition of zirconium.
  • the alloy according to the invention has the second best resistance to inter-crystalline corrosion on the surface, and the best at heart.
  • the difference between results at heart and on the surface is low, which is a favorable property when the structural element is manufactured by machining.
  • those according to the invention have an arrow such that the product fe is less than 0.10 l 2 , which is, as can be seen in the patent EP 0731185 mentioned above, the indication of a low rate of internal stresses.
  • the recrystallization rate (in%) at the surface, at quarter-thickness and at the core was measured by image analysis on micrographs of the 4 preceding samples. The results are shown in Table 8: Alloy e (mm) Area Recrist rate (Quarter-th.) Recrist rate (to heart) 2024 40 80 60 30 2034 40 12 0 0 Inv. 40 100 100 100 Inv. 80 100 100 100 100
  • the alloy according to the invention has, in the treated state, a structure completely recrystallized throughout the thickness of the product.
  • Samples according to the invention with a thickness of 15, 40 and 80 mm, treated in the T851 state, were measured with a hot-rolling entry temperature of 475 ° C. and dissolved for 2 hours. at 528 ° C, and a 24 hour yield at 173 ° C, the static mechanical characteristics (yield strength R 0.2 and tensile strength R m in MPa and elongation A in%)) at quarter-thickness and at half -thickness, in the L and TL directions.
  • the overall results are reproduced in Table 9. They show the low evolution of the properties as a function of the thickness, resulting from a low sensitivity to quenching. e (mm) CUPS.
  • These sheets are particularly suitable for the manufacture of wingtip elements. of aircraft by a manufacturing range involving machining and one or more formatting operations.

Description

Domaine de l'inventionField of the invention

L'invention concerne des éléments de structure d'avion, notamment des panneaux de peau et des raidisseurs d'intrados de voilure pour avions commerciaux de grande capacité, réalisés à partir de produits laminés, filés ou forgés en alliage AlCuMg à l'état traité par mise en solution, trempe et revenu, et présentant, par rapport aux produits de l'art antérieur utilisés pour la même application, un compromis amélioré entre les différentes propriétés d'emploi requises.The invention relates to aircraft structural elements, in particular skin and underside sail stiffeners for large commercial aircraft capacity, made from rolled, extruded or forged AlCuMg alloy the state treated by dissolution, quenching and tempering, and presenting, with respect to prior art products used for the same application, an improved compromise between the different job properties required.

La désignation des alliages et des états métallurgiques utilisée ci-après correspond à la nomenclature de l'Aluminum Association, reprise par les normes européennes EN 515 et EN 573 partie 3.The designation of the alloys and metallurgical states used below corresponds to the nomenclature of the Aluminum Association, adopted by European standards EN 515 and EN 573 part 3.

Etat de la techniqueState of the art

Les ailes d'avions commerciaux de grande capacité comportent une partie supérieure (ou extrados) constituée d'une peau fabriquée à partir de tôles épaisses en alliage 7150 à l'état T651, ou en alliage 7055 à l'état T7751 ou 7449 à l'état T7951, et de raidisseurs fabriqués à partir de profilés du même alliage, et une partie inférieure (ou intrados) constituée d'une peau fabriquée à partir de tôles épaisses en alliage 2024 à l'état T351 ou 2324 à l'état T39, et de raidisseurs fabriqués à partir de profilés du même alliage. Les deux parties sont assemblées par des longerons et des nervures.Large commercial aircraft wings have an upper section (or extrados) consisting of a skin made from thick alloy plates 7150 to the T651 state, or 7055 alloy to the T7751 state or 7449 to the T7951 state, and stiffeners made from profiles of the same alloy, and a lower part (or intrados) consisting of a skin made from thick 2024 alloy state T351 or 2324 in state T39, and stiffeners made from same alloy. Both parts are assembled by longitudinal members and ribs.

L'alliage 2024 selon la désignation de l'Aluminum Association ou la norme EN 573-3 a la composition chimique suivante (% en poids) :
Si < 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 < 0,15The alloy 2024 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

Dans le but d'améliorer le compromis entre les différentes propriétés requises, notamment la résistance mécanique et la ténacité, diverses solutions alternatives ont été proposées. Boeing a développé l'alliage 2034 de composition :
Si < 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 - 0,15In 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

Cet alliage a fait l'objet du brevet EP 0031605 (= US 4336075). Il présente, par rapport au 2024 à l'état T351, une meilleure limite d'élasticité spécifique due à l'augmentation de la teneur en manganèse et à l'ajout d'un autre antirecristallisant (Zr), ainsi qu'une ténacité et une résistance à la fatigue améliorées.This alloy was the subject of patent EP 0031605 (= US 4336075). He presents, by 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.

Le brevet US 5652063 (Alcoa) concerne un élément de structure d'avion réalisé à partir d'un alliage de composition (% en poids) :
Cu: 4,85 - 5,3   Mg: 0,51 - 1,0   Mn : 0,4 - 0,8   Ag: 0,2 - 0,8    Si < 0,1   Fe < 0,1   Zr < 0,25 avec Cu/Mg compris entre 5 et 9.US Patent 5652063 (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, With Cu / Mg between 5 and 9.

La tôle de cet alliage à l'état T8 présente une limite d'élasticité > 77 ksi (531 MPa). L'alliage est particulièrement destiné aux avions supersoniques.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.

Le brevet US 5593516 (Reynolds) concerne un alliage pour applications aéronautiques contenant de 2,5 à 5,5% Cu et 0,1 à 2,3% Mg, dans lequel les teneurs en Cu et Mg sont maintenues en dessous de leur limite de solubilité dans l'aluminium, et sont liées par les équations :
Cumax = 5,59 - 0,91 Mg et Cumm = 4,59 - 0,91MgUS Pat. No. 5,593,516 (Reynolds) relates to an alloy for aeronautical applications containing from 2.5 to 5.5% Cu and from 0.1 to 2.3% Mg, in which the contents of Cu and Mg are kept below their limit. of solubility in aluminum, and are linked by the equations:
Cu max = 5.59 - 0.91 Mg and Cu mm = 4.59 - 0.91Mg

L'alliage peut contenir également : Zr < 0,20% V < 0,20% Mn < 0,80% Ti < 0,05% Fe < 0,15% Si < 0,10%The alloy may also contain: Zr <0.20% V <0.20% Mn <0.80% Ti <0.05% Fe <0.15% If <0.10%

Les brevets US 5376192 et US 5512112, issus de la même demande initiale, concernent des alliages de ce type contenant de 0,1 à 1% d'argent. On peut remarquer que l'utilisation d'argent dans ce type d'alliage conduit à une augmentation du coût d'élaboration et des difficultés pour le recyclage des chutes de fabrication.US Pat. Nos. 5,376,192 and 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 an increase in the cost development and difficulties in the recycling of manufacturing scrap.

Par ailleurs, on connaít depuis de nombreuses années des alliages du type « AU6MGT » selon l'ancienne désignation des alliages en France. Le brevet FR 1379764, déposé en 1963 par Pechiney, concerne l'utilisation d'un alliage de ce type de composition : Cu : 5 - 7 Mg : 0,10 - 0,50 Mn 0,05 - 0,50 Si < 0,30 Fe < 0,50 Ti : 0,05 - 0,25 pour la fabrication de bouteilles pour gaz comprimés.Furthermore, for many years, alloys of the type have been known. "AU6MGT" according to the former designation of alloys in France. The FR patent 1379764, filed in 1963 by Pechiney, relates to the use of an alloy of this type of composition: Cu: 5 - 7 Mg: 0,10 - 0,50 Mn 0,05 - 0,50 If <0,30 Fe <0.50 Ti: 0.05 - 0.25 for the manufacture of compressed gas bottles.

L'Aluminum Association a enregistré en 1976 l'alliage 2001 de composition :
Cu : 5,2 - 6   Mg : 0,20 - 0,45   Mn: 0,15 - 0,50   Si < 0,20   Fe < 0,20    Cr < 0,10   Zn < 0,10   Ni < 0,05   Ti < 0,20   Zr < 0,05 The Aluminum Association recorded in 1976 the 2001 alloy of composition:
Cu: 5.2 - 6 Mg: 0.20 - 0.45 Mn: 0.15 - 0.50 If <0.20 Fe <0.20 Cr <0.10 Zn <0.10 Ni <0.05 Ti <0.20 Zr <0.05

A la connaissance des inventeurs, il n'existe pas d'autre utilisation industrielle de cet alliage que les bouteilles de gaz comprimés fabriquées par filage inverse.To the knowledge of the inventors, there is no other industrial use of this alloy than compressed gas cylinders made by reverse spinning.

Problème poséProblem

Dans la construction d'avions commerciaux, la tendance actuelle est à l'utilisation croissante de produits très épais, dans la masse desquels les éléments de structure sont usinés. Par exemple, pour certains avions de petite dimension, les peaux de voilure sont usinées à partir de tôles relativement épaisses afin de permettre l'usinage dans la masse des raidisseurs de voilure, alors que ceux-ci sont habituellement réalisés à partir de profilés ou de tôles pliées, et sont ensuite fixées mécaniquement à la peau. L'usinage intégral dans la masse de l'ensemble peau-raidisseurs conduit à une réduction des coûts de fabrication, puisque le nombre de pièces est réduit et qu'on évite l'assemblage. Par ailleurs, l'utilisation d'une structure non assemblée permet une réduction du poids de l'ensemble.In the construction of commercial aircraft, the current trend is to use growing of very thick products, in the bulk of which structural elements are machined. For example, for some small aircraft, the skins of wing are machined from relatively thick sheets to allow machining in the mass of wing stiffeners, whereas these are usually made from profiles or folded sheets, and are then mechanically fixed to the skin. The integral machining in the mass of the skin-stiffener assembly leads to a reduction in manufacturing costs, since the number of parts is reduced and we avoid assembly. Moreover, the use of an unassembled structure allows a reduction of the weight of the whole.

Il est donc souhaitable qu'en plus des propriétés habituellement recherchées pour les éléments de structure d'avions, à savoir une résistance mécanique élevée, une bonne tolérance aux dommages et une bonne résistance à la fatigue et aux différentes formes de corrosion, les tôles présentent des caractéristiques mécaniques homogènes sur toute l'épaisseur, c'est-à-dire que les propriétés ne varient pas de manière significative en fonction de l'épaisseur, typiquement entre 10 et 120 mm. D'autre part, plus on recourt à l'usinage, plus la stabilité à l'usinage est souhaitable, ce qui s'obtient par un faible niveau de contraintes internes. Or, il est connu que, pour une tôle épaisse, les caractéristiques mécaniques sont d'autant plus homogènes, et les contraintes internes d'autant plus réduites, que la tôle présente une faible sensibilité à la trempe.It is therefore desirable that in addition to the properties usually sought for aircraft structural elements, namely high mechanical strength, good tolerance to damage and good resistance to fatigue and different forms of corrosion, the sheets have homogeneous mechanical properties over the entire thickness, that is, the properties do not vary so significant depending on the thickness, typically between 10 and 120 mm. Else On the one hand, the more machining is used, the greater the stability to machining is desirable, which is obtained by a low level of internal constraints. However, it is known that for a sheet metal, the mechanical characteristics are all the more homogeneous, and the internal constraints all the more reduced, as the sheet has a low sensitivity to the quenching.

Enfin, les ailes d'avions, notamment les avions de grande capacité, présentent un profil d'aile galbé, avec une courbure à la fois dans le sens longitudinal et dans le sens transversal. Cette forme complexe peut être obtenue pendant l'opération de revenu dans un autoclave, par formage sur un moule, en mettant en dépression la face de la tôle du côté du moule par rapport à la face opposée, à l'aide d'un vide partiel. Il est impératif que cette opération soit réussie, pour éviter le rebut coûteux d'une pièce à forte valeur ajoutée, notamment pour les pièces de grande dimension. Le gage du succès réside dans un retour élastique le plus faible possible pour une forme de moule donnée, car le retour élastique est le plus souvent difficile à contrôler.Finally, aircraft wings, especially large aircraft, have a curved wing profile, with a curvature both in the longitudinal direction and in the transversal direction. This complex shape can be obtained during the operation of returned to an autoclave, by forming on a mold, by depressing the face the side of the mold with respect to the opposite side, using a partial vacuum. he It is imperative that this operation be successful, to avoid the costly waste of a part high added value, especially for large parts. The pledge of success lies in the lowest possible springback for a form of mold given, because the springback is usually difficult to control.

Le but de la présente invention est de fournir des éléments de structure d'avions présentant des propriétés au moins équivalentes à celles des mêmes éléments réalisés en alliage 2024 à l'état T351 en ce qui concerne les caractéristiques mécaniques statiques, la ténacité, la vitesse de propagation de fissures et la résistance à la corrosion, en utilisant des produits laminés, filés ou forgés présentant un faible niveau de contraintes résiduelles, une faible sensibilité à la trempe et une bonne aptitude au formage au revenu.The object of the present invention is to provide aircraft structural elements having properties at least equivalent to those of the same elements made alloy 2024 in the T351 state with respect to mechanical characteristics static, toughness, speed of crack propagation and resistance to corrosion, using rolled, spun or forged products with a low level of residual stresses, low sensitivity to quenching and good ability to train for income.

Objet de l'inventionObject of the invention

L'invention a pour objet un élément de structure d'avion, notamment un élément d'intrados d'aile d'avion, réalisé à partir d'un produit laminé, filé ou forgé, en alliage de composition (% en poids) :
Cu : 4,6 - 5,3   Mg : 0,10 - 0,50   Mn : 0,25 - 0,45   Si < 0,10   Fe < 0,15 Zn < 0,20 Cr < 0,10 autres éléments < 0,05 chacun et < 0,15 au total, reste Al traité par mise en solution, trempe, détensionnement par traction contrôlée à plus de 1,5% de déformation permanente et revenu.The subject of the invention is an aircraft structural element, in particular an aircraft wing surface element, made from a rolled or forged product, made of an alloy of composition (% by weight):
Cu: 4.6 - 5.3 Mg: 0.10 - 0.50 Mn: 0.25 - 0.45 Si <0.10 Fe <0.15 Zn <0.20 Cr <0.10 other elements < 0.05 each and <0.15 in total, remains Al treated by dissolution, quenching, controlled tensile stress relieving to more than 1.5% of permanent deformation and income.

Cet élément présente l'une au moins des propriétés suivantes :

  • limite d'élasticité R0.2 (sens TL) > 350 MPa, de préférence > 370 MPa,
  • ténacité K1c (sens L-T) > 42 MPa√m
  • résistance à la corrosion intercristalline de type P selon la norme ASTM G110.
This element has at least one of the following properties:
  • yield strength R 0.2 (TL direction)> 350 MPa, preferably> 370 MPa,
  • toughness K 1c (LT direction)> 42 MPa√m
  • P type intercrystalline corrosion resistance according to ASTM G110.

L'invention a également pour objet un procédé de fabrication d'un élément de structure comportant :

  • a) la coulée d'une plaque ou d'une billette de la composition mentionnée ci-dessus,
  • b) l'homogénéisation de cette plaque ou billette,
  • c) la transformation à chaud de cette plaque par laminage ou de cette billette par filage ou forgeage pour obtenir un produit d'épaisseur supérieure à 10 mm,
  • d) la trempe du produit transformé à chaud,
  • e) la mise en solution de ce produit, de préférence à une température inférieure de moins de 10°C à la température de fusion commençante de l'alliage,
  • f) le détensionnement par traction contrôlée du produit jusqu'à une déformation permanente de plus de 1,5%,
  • g) le revenu du produit à une température supérieure à 160°C, éventuellement associé à un formage,
  • h) l'usinage du produit éventuellement formé jusqu'à la forme finale de l'élément de structure.
    • The invention also relates to a method of manufacturing a structural element comprising:
  • a) pouring a plate or a billet of the composition mentioned above,
  • b) the homogenization of this plate or billet,
  • c) hot processing of this plate by rolling or billet by spinning or forging to obtain a product with a thickness greater than 10 mm,
  • d) quenching the hot-processed product,
  • e) the dissolution of this product, preferably at a temperature less than 10 ° C at the starting melting temperature of the alloy,
  • f) controlled tensile stress relieving of the product to a permanent deformation of more than 1.5%,
  • (g) the product's income at a temperature greater than 160 ° C, possibly associated with forming,
  • h) machining the product possibly formed to the final shape of the structural element.

    • Dans le cas où le produit est une tôle, la température d'entrée au laminage à chaud est de préférence inférieure d'au moins 40°C, et plus préférentiellement d'au moins 50°C, à la température de mise en solution.In the case where the product is a sheet, the inlet temperature to hot rolling is preferably at least 40.degree. C., and more preferably at least 40.degree. 50 ° C, at the dissolution temperature.

    Description de l'inventionDescription of the invention

    L'invention repose sur la constatation qu'un alliage de type 2001, avec certaines modifications de composition et une gamme de fabrication appropriée, pouvait présenter un ensemble de propriétés le rendant apte à l'utilisation dans des structures d'avions, et plus particulièrement dans les intrados d'ailes d'avions commerciaux de grande capacité, avec en plus des propriétés intéressantes en matière de faible sensibilité à le trempe, de faibles contraintes résiduelles et de formage au revenu.The invention is based on the finding that a 2001 type alloy, with certain changes in composition and an appropriate manufacturing range, could present a set of properties making it suitable for use in structures of aircraft, and more particularly in the lower surface of the wings of commercial aircraft large capacity, with more interesting properties in terms of low quenching sensitivity, low residual stresses and income shaping.

    Par rapport à l'alliage 2001, la plage de teneur en cuivre est nettement décalée vers le bas, tout en restant supérieure à celle des alliages 2024 ou 2034 pour intrados, pour compenser, dans son influence sur la résistance mécanique, la faible teneur en magnésium. Il est préférable de choisir une teneur en cuivre supérieure à 4,8%, voire à 4,9% ou même 5%. La teneur en magnésium est du même ordre que dans l'alliage 2001, et de préférence située entre 0,20 et 0,40%. Le rapport Cu/Mg est ainsi pratiquement toujours supérieur à 10, contrairement à l'enseignement du brevet US 5652063, qui préconise un rapport Cu/Mg compris entre 5 et 9.Compared to the 2001 alloy, the copper content range is significantly shifted towards the low, while remaining higher than that of the 2024 or 2034 alloys for intrados, for compensate, in its influence on the mechanical strength, the low magnesium. It is better to choose a copper content higher than 4.8% or at 4.9% or even 5%. The magnesium content is of the same order as in the alloy 2001, and preferably between 0.20 and 0.40%. The ratio Cu / Mg is thus almost always above 10, contrary to the teaching of the US patent 5652063, which recommends a Cu / Mg ratio of between 5 and 9.

    La teneur en manganèse est contrôlée dans une plage relativement étroite. En dessous de 0,15%, on risquerait d'avoir un grain trop gros ; au-dessus de 0,45%, on obtient une structure non recristallisée qui n'est pas favorable à la maítrise des contraintes résiduelles. Un domaine préférentiel est compris entre 0,25 et 0,40%. Il est à noter que, pour la même raison, et contrairement à l'enseignement du brevet US 5593516, l'alliage ne comporte aucun autre élément anti-recristallisant tel que le vanadium ou le zirconium.The manganese content is controlled in a relatively narrow range. Below 0.15%, you could have a grain too big; above 0.45%, we get a non recrystallized structure which is not favorable to the control of the constraints residual. A preferred range is from 0.25 to 0.40%. Note that, for the same reason, and contrary to the teaching of US Patent 5593516, the alloy contains no other anti-recrystallizing element such as vanadium or zirconium.

    Les teneurs en fer et en silicium sont maintenues respectivement en dessous de 0,15 et 0,10%, et de préférence en dessous de 0,09 et 0,08%, pour garantir une bonne ténacité. L'alliage peut comporter jusqu'à 0,2% de zinc, cette addition ayant un effet favorable sur la résistance mécanique, sans risque pour d'autres propriétés, comme la résistance à la corrosion.The iron and silicon contents are maintained below 0.15, respectively. and 0.10%, and preferably below 0.09 and 0.08%, to ensure good tenacity. The alloy may comprise up to 0.2% zinc, this addition having an effect favorable on the mechanical strength, without risk for other properties, such as corrosion resistance.

    La gamme de transformation comporte la coulée d'une plaqué ou d'une billette, un réchauffage ou une homogénéisation à une température proche de la température de fusion commençante de l'alliage et une transformation à chaud par laminage, filage ou forgeage. Dans le cas du laminage, celui-ci peut comporter une passe, dite d'élargissement, dans le sens perpendiculaire à celui des autres passes, et destiné à améliorer l'isotropie du produit. La température de transformation à chaud se situe, de préférence, à un niveau légèrement plus bas que celle qu'adopterait l'homme de métier en référence à la température de mise en solution. Ainsi, en ce qui concerne le laminage, la température d'entrée se situe, de préférence, à au moins 40°C, voire 50°C, en dessous de la température de mise en solution, et la température de sortie de 20 à 30°C en dessous de la température d'entrée.The transformation range includes the casting of a veneer or billet, a reheating or homogenization at a temperature close to the temperature of beginning fusion of the alloy and a hot transformation by rolling, spinning or forging. In the case of rolling, it may include a pass, called enlargement, in the direction perpendicular to that of the other passes, and intended for improve the isotropy of the product. The hot transformation temperature is, preferably at a slightly lower level than that which would be profession with reference to the solution temperature. So, as far as the rolling, the inlet temperature is preferably at least 40 ° C or 50 ° C, below the solution temperature, and the outlet temperature of 20 to 30 ° C below the inlet temperature.

    Le produit est soumis ensuite à une mise en solution aussi complète que possible, à une température proche, par exemple moins de 10°C en dessous, de la température de fusion commençante de l'alliage, tout en évitant la brûlure. Cette température se situe entre 520 et 535°C. La qualité de la mise en solution peut être contrôlée par analyse enthalpique différentielle. Le produit est ensuite trempé, par exemple par immersion dans l'eau froide, de manière à assurer une vitesse de refroidissement comprise entre 10 et 50°C/s. Après la trempe, le produit est tractionné jusqu'à une déformation permanente d'au moins 1,5%, de manière à le détensionner et à améliorer sa planéité.The product is then submitted to a dissolution as complete as possible, to a temperature close to, for example less than 10 ° C below, the temperature of beginning fusion of the alloy, while avoiding the burn. This temperature is between 520 and 535 ° C. The quality of the dissolution in solution can be controlled by analysis differential enthalpy. The product is then quenched, for example by immersion in cold water, so as to ensure a cooling rate of between 10 and 50 ° C / s. After quenching, the product is triturated until deformation at least 1.5%, so as to relax it and improve its flatness.

    Pour l'alliage selon l'invention, cette traction a également pour effet d'améliorer, par un effet d'écrouissage, la limite d'élasticité après revenu, de sorte qu'on peut qualifier l'état obtenu d'état T851, comme s'il s'agissait d'une passe spécifique d'écrouissage après trempe. Comme indiqué plus haut, le revenu proprement dit peut s'effectuer en même temps que la mise en forme du galbe de l'intrados. Ce revenu est effectué de préférence à une température supérieure à 160°C (et plus préférentiellement > 170°C), d'une durée permettant d'atteindre le pic de limite d'élasticité, comme pour un état T6. Typiquement, un revenu de temps équivalent à celui correspondant à 12 à 24 h à une température de 173°C est effectué ; toute combinaison temps-température permettant d'atteindre le pic de revenu de l'alliage est utilisable.For the alloy according to the invention, this traction also has the effect of improving, by a hardening effect, the elasticity limit after income, so that one can qualify the state obtained from state T851, as if it were a specific pass hardening after quenching. As noted above, the income itself can at the same time as the shaping of the curve of the intrados. This income is preferably at a temperature greater than 160 ° C (and higher preferentially> 170 ° C.), of a duration enabling the limit peak to be reached elasticity, as for a T6 state. Typically, a time income equivalent to that corresponding to 12 to 24 h at a temperature of 173 ° C is carried out; all time-temperature combination allowing to reach the peak of income of the alloy is usable.

    La structure métallurgique obtenue est, à l'inverse de celle des alliages 2024 et 2034, fortement recristallisée, avec un taux de recristallisation dépassant toujours 70%, et le plus souvent 90%, sur toute l'épaisseur.The metallurgical structure obtained is, unlike that of alloys 2024 and 2034, strongly recrystallized, with a recrystallization rate still exceeding 70%, and the more often 90%, over the entire thickness.

    Les éléments de structure selon l'invention présentent un compromis de propriétés (caractéristiques mécaniques statiques, ténacité, vitesse de propagation de fissures, résistance à la corrosion) qui les rendent aptes à être utilisés dans la construction aéronautique, et notamment à la fabrication d'intrados d'ailes. De plus, ces éléments peuvent être aisément réalisés par usinage et formés au revenu. Enfin, l'alliage utilisé se révèle facilement soudable par les techniques habituelles, ce qui peut permettre de réduire le nombre des assemblages rivetés.The structural elements according to the invention have a compromise of properties (static mechanical characteristics, toughness, speed of crack propagation, corrosion resistance) that make them suitable for use in construction aeronautics, and in particular to the manufacture of wing bottoms. In addition, these elements can be easily made by machining and trained to income. Finally, the alloy used is easily soldered by the usual techniques, which can reduce the number of riveted assemblies.

    ExemplesExamples Exemple 1Example 1

    On a préparé 6 alliages dont la composition est indiquée au tableau 1. L'alliage A est un alliage 2024-T3 de composition habituelle pour l'application intrados de voilure. L'alliage B est un alliage entrant dans le domaine-de composition décrit dans le brevet US 5652063, mais sans addition d'argent. L'alliage C est conforme à l'invention. Les alliages D et E ne diffèrent de l'alliage C que par un silicium plus élevé pour D, un manganèse et un cuivre plus élevés pour E et F, et une addition de zirconium pour F. Alliage Si Fe Cu Mn Mg Ti Zr A 0,07 0,07 4,11 0,53 1,28 0,008 B 0,06 0,08 4,73 0,30 0,67 0,065 C 0,05 0,08 5,26 0,30 0,28 0,062 D 0,15 0,08 5,28 0,30 0,31 0,065 E 0,07 0,10 5,64 0,99 0,29 0,012 F 0,06 0,08 5,47 0,67 0,29 0,014 0,11 Six alloys have been prepared, the composition of which is indicated in Table 1. The alloy A is a 2024-T3 alloy of the usual composition for the intrados application of the wing. Alloy B is an alloy falling within the composition range described in US Patent No. 5652063, but without the addition of silver. Alloy C is in accordance with the invention. Alloys D and E differ from alloy C only by higher silicon for D, higher manganese and copper for E and F, and zirconium addition for F. Alloy Yes Fe Cu mn mg Ti Zr AT 0.07 0.07 4.11 0.53 1.28 0,008 B 0.06 0.08 4.73 0.30 0.67 0,065 VS 0.05 0.08 5.26 0.30 0.28 0.062 D 0.15 0.08 5.28 0.30 0.31 0,065 E 0.07 0.10 5.64 0.99 0.29 0.012 F 0.06 0.08 5.47 0.67 0.29 0.014 0.11

    Des plaques coulées de section 380 x 120 mm ont été homogénéisées, laminées à chaud à l'épaisseur 22 mm, mises en solution, trempées à l'eau froide, tractionnées à 2,3% de déformation permanente et revenues. Les paramètres de l'homogénéisation, du laminage à chaud (températures d'entrée), de mise en solution et de revenu sont indiqués au tableau 2. Alliage Homogénéisation Laminage à Chaud (entrée) Mise en Solution Revenu A 4h 490°C 467°C 3h à 497°C - B 4h 490°C 467°C 3h à 518°C 16h à 173°C C 4h 490°C 467°C 6h à 527°C 16h à 173°C D 4h 490°C 472°C 6h à 527°C 16h à 173°C E 1h 520°C 479°C 6h à 527°C 16h à 173°C F 1h 520°C 474°C 6h à 527°C 16h à 173°C Cast plates with a cross section of 380 × 120 mm were homogenized, hot-rolled to a thickness of 22 mm, put into solution, quenched with cold water, triturated at 2.3% permanent deformation and returned. The parameters of homogenization, hot rolling (inlet temperatures), dissolution and tempering are given in Table 2. Alloy homogenization Hot Rolling (entrance) Dissolution Returned AT 4h 490 ° C 467 ° C 3h at 497 ° C - B 4h 490 ° C 467 ° C 3h to 518 ° C 16h at 173 ° C VS 4h 490 ° C 467 ° C 6h at 527 ° C 16h at 173 ° C D 4h 490 ° C 472 ° C 6h at 527 ° C 16h at 173 ° C E 1h 520 ° C 479 ° C 6h at 527 ° C 16h at 173 ° C F 1h 520 ° C 474 ° C 6h at 527 ° C 16h at 173 ° C

    On a mesuré sur les tôles traitées les caractéristiques mécaniques : résistance à la rupture Rm (en MPa), limite d'élasticité conventionnelle à 0,2% R0,2 (en MPa) et allongement à la rupture A (en %), sur des éprouvettes de traction de section circulaire selon la norme ASTM B 557, prélevées à mi-épaisseur dans les sens L et TL (3 éprouvettes par cas).The mechanical characteristics were measured on the treated sheets: tensile strength R m (in MPa), conventional yield strength at 0.2% R 0.2 (in MPa) and elongation at break A (in%) , on test specimens of circular section according to ASTM B 557, taken at mid-thickness in the L and TL directions (3 specimens per case).

    On a mesuré également la ténacité par le facteur d'intensité critique de contrainte K1c (en MPa√m) mesuré, selon la norme ASTM E 399, sur des éprouvettes CT20 prélevées à quart-épaisseur dans les sens L-T et T-L (2 éprouvettes par cas).The toughness was also measured by the critical stress intensity factor K 1c (in MPa√m) measured, according to ASTM E 399, on CT20 specimens taken at quarter-thickness in the LT and TL directions (2 test pieces by case).

    L'ensemble des résultats est regroupé au tableau 3. Alliage Rm (L) R0,2 (L) A (L) Rm (TL) R0,2 (TL) A (TL) K1c (L-T) K1c (T-L) A 472 362 21,3 467 321 21,4 41,8 40,5 B 476 439 12,5 475 427 11,2 41,3 34,6 C 458 396 13,9 463 384 12,6 45,4 42,9 D 460 397 13,6 465 387 12,2 40,5 36,4 E 488 423 10,5 480 403 9,4 36,8 29,3 F 480 418 11,6 481 402 10,1 40,2 33,6 The overall results are summarized in Table 3. Alloy R m (L) R 0.2 (L) A (L) R m (TL) R 0.2 (TL) A (TL) K 1c (LT) K 1c (TL) AT 472 362 21.3 467 321 21.4 41.8 40.5 B 476 439 12.5 475 427 11.2 41.3 34.6 VS 458 396 13.9 463 384 12.6 45.4 42.9 D 460 397 13.6 465 387 12.2 40.5 36.4 E 488 423 10.5 480 403 9.4 36.8 29.3 F 480 418 11.6 481 402 10.1 40.2 33.6

    On constate que l'alliage C selon l'invention conduit à une limite d'élasticité nettement supérieure à celle du 2024, et un peu plus faible que celle des alliages B, E et F. L'allongement est plus faible que pour le 2024, mais meilleur que celui des alliages B, D, E et F. La ténacité est la meilleure de tous les alliages testés. On a donc un compromis favorable de ces diverses propriétés. En particulier, les résultats montrent l'effet défavorable, à la fois sur la ténacité et l'allongement, d'une augmentation de la teneur en silicium et en manganèse, ainsi que d'une addition de zirconium.It is found that the alloy C according to the invention leads to a limit of elasticity significantly higher than that of 2024, and slightly lower than that of alloys B, E and F. Elongation is lower than for 2024, but better than that of alloys B, D, E and F. Toughness is the best of all the alloys tested. So we have a favorable compromise of these various properties. In particular, the results show the adverse effect, both on toughness and elongation, of increase in the silicon and manganese content, as well as an addition of zirconium.

    On a procédé par ailleurs à des essais accélérés de corrosion intercristalline sur des échantillons des 6 alliages, à l'état T351 pour l'alliage 2024 (A) et T851 pour les autres, en surface et à coeur, selon la norme ASTM G110. On note le type de corrosion observé : P pour piqûres, I pour corrosion intercristalline et P + I pour les deux. On mesure la profondeur maximum (P max en µm), la profondeur de corrosion intercristalline (P CI en µm) et le pourcentage de corrosion intercristalline sur l'échantillon. Les résultats sont indiqués au tableau 4 : All. Surf. Surf. Surf. Surf. Coeur Coeur Coeur Coeur Type P max P CI % CI Type P max P CI % CI A I+P 160 70 10 I + P 260 260 60 B P+I 130 30 10 P+I 160 50 10 C P 150 - - P 120 - - D P 150 - - P 120 - - E P 200 - - P 140 - - F P 220 - - P 170 - - Accelerated tests for intercrystalline corrosion were also carried out on samples of the 6 alloys, in the T351 state for alloy 2024 (A) and T851 for the others, at the surface and at the core, according to the ASTM G110 standard. Note the type of corrosion observed: P for pitting, I for intercrystalline corrosion and P + I for both. The maximum depth (P max in μm), the intercrystalline corrosion depth (P CI in μm) and the percentage of intercrystalline corrosion on the sample are measured. The results are shown in Table 4: All. Surf. Surf. Surf. Surf. Heart Heart Heart Heart Type P max P CI % THIS Type P max P CI % THIS AT I + P 160 70 10 I + P 260 260 60 B P + I 130 30 10 P + I 160 50 10 VS P 150 - - P 120 - - D P 150 - - P 120 - - E P 200 - - P 140 - - F P 220 - - P 170 - -

    On observe que l'alliage selon l'invention présente la seconde meilleure résistance à la corrosion inter cristalline en surface, et la meilleure à coeur. La différence entre les résultats à coeur et en surface est faible, ce qui est une propriété favorable lorsque l'élément de structure est fabriqué par usinage.It is observed that the alloy according to the invention has the second best resistance to inter-crystalline corrosion on the surface, and the best at heart. The difference between results at heart and on the surface is low, which is a favorable property when the structural element is manufactured by machining.

    On a comparé enfin, pour les alliages A et C, les vitesses de propagation de fissures de fatigue da/dn dans la direction T-L, en mm/cycle, pour des valeurs de ΔK comprises entre 15 et 30 MPa√m, selon la norme ASTM E647. Les résultats (2 essais par alliage) sont indiqués au tableau 5. Alliage 10 MPa√m 15 MPa√m 20 MPa√m 25 MPa√m 30 MPa√m A 6,2 10-5 3,8 10-4 8,3 10-4 1,8 10-3 3,8 10-3 A 6,3 10-5 3,8 10-4 8,7 10-4 1,9 10-3 3,6 10-3 C 1,2 10-4 4,0 10-4 8,6 10-4 1,5 10-3 2,6 10-3 C 1,2 10-4 4,2 10-4 9,5 10-4 1,8 10-3 3,1 10-3 Finally, for alloys A and C, the fatigue crack growth rates da / dn in the TL direction, in mm / cycle, were compared for ΔK values between 15 and 30 MPa√m, according to the standard ASTM E647. The results (2 tests per alloy) are shown in Table 5. Alloy 10 MPa√m 15 MPa√m 20 MPa√m 25 MPa√m 30 MPa√m AT 6.2 10 -5 3.8 10 -4 8.3 10 -4 1.8 10 -3 3.8 10 -3 AT 6.3 10 -5 3.8 10 -4 8.7 10 -4 1.9 10 -3 3.6 10 -3 VS 1.2 10 -4 4.0 10 -4 8.6 10 -4 1.5 10 -3 2.6 10 -3 VS 1.2 10 -4 4.2 10 -4 9.5 10 -4 1.8 10 -3 3.1 10 -3

    On observe que les résultats sont à peu près comparables pour les deux alliages. It is observed that the results are roughly comparable for both alloys.

    Exemple 2Example 2

    On a mesuré le niveau de contraintes résiduelles sur des tôles d'épaisseur 40 mm en alliage 2024, 2034 et selon l'invention, traitées toutes trois au même état T351. Les compositions (% en poids) sont données au tableau 6 : Alliage Si Fe Cu Mn Mg Ti Zr 2024 0,12 0,20 4,06 0,54 1,36 0,02 2034 0,05 0,07 4,30 0,98 1,34 0,02 0,10 Invent. 0,05 0,07 5,12 0,35 0,29 0,02 The level of residual stresses was measured on sheets of thickness 40 mm in alloy 2024, 2034 and according to the invention, all three treated in the same state T351. The compositions (% by weight) are given in Table 6: Alloy Yes Fe Cu mn mg Ti Zr 2024 0.12 0.20 4.06 0.54 1.36 0.02 2034 0.05 0.07 4.30 0.98 1.34 0.02 0.10 Invent. 0.05 0.07 5.12 0.35 0.29 0.02

    La méthode de mesure des contraintes résiduelles est la méthode du barreau décrite dans le brevet EP 0731185 de la demanderesse. On a mesuré les flèches fL et fTL dans les sens L et TL (en microns) et calculé dans les deux cas le quotient fe/l2, l'épaisseur e et la longueur l du barreau étant exprimés en mm. Les résultats sont donnés au tableau 7 : alliage e (mm) 1 (mm) fL (µm) fLe/l2 fTL (µm) fTLe/l2 2024 40 180 210 0,26 120 015 2034 40 180 147 0,18 129 0,16 invention 40 180 46 0,06 4 0,005 invention 80 385 84 0,05 136 0,07 The method of measurement of residual stresses is the bar method described in patent EP 0731185 of the applicant. The arrows f L and f TL were measured in the L and TL directions (in microns) and in both cases the quotient f e / 1 2 was calculated, the thickness e and the length l of the bar being expressed in mm. The results are given in Table 7: alloy e (mm) 1 (mm) f L (μm) f L e / l 2 f TL (μm) f TL e / l 2 2024 40 180 210 0.26 120 015 2034 40 180 147 0.18 129 0.16 invention 40 180 46 0.06 4 0.005 invention 80 385 84 0.05 136 0.07

    On constate que, contrairement aux éprouvettes en alliage 2024 ou 2034, celles selon l'invention présentent une flèche telle que le produit fe est inférieur à 0,10 l2, ce qui est, comme on peut le voir dans le brevet EP 0731185 mentionné ci-dessus, l'indication d'un faible taux de contraintes internes.It is noted that, unlike alloy test pieces 2024 or 2034, those according to the invention have an arrow such that the product fe is less than 0.10 l 2 , which is, as can be seen in the patent EP 0731185 mentioned above, the indication of a low rate of internal stresses.

    On a mesuré, par analyse d'image sur des micrographies des 4 échantillons précédents, le taux de recristallisation (en %) en surface, à quart-épaisseur et à coeur. Les résultats sont indiqués au tableau 8 : Alliage e (mm) Surface Taux recrist. (quart-ép.) Taux recrist. (à coeur) 2024 40 80 60 30 2034 40 12 0 0 Inv. 40 100 100 100 Inv. 80 100 100 100 The recrystallization rate (in%) at the surface, at quarter-thickness and at the core was measured by image analysis on micrographs of the 4 preceding samples. The results are shown in Table 8: Alloy e (mm) Area Recrist rate (Quarter-th.) Recrist rate (to heart) 2024 40 80 60 30 2034 40 12 0 0 Inv. 40 100 100 100 Inv. 80 100 100 100

    On constate que l'alliage selon l'invention présente, à l'état traité, une structure complètement recristallisée dans toute l'épaisseur du produit..It is found that the alloy according to the invention has, in the treated state, a structure completely recrystallized throughout the thickness of the product.

    Exemple 3Example 3

    On a mesuré sur des échantillons selon l'invention, d'épaisseur 15, 40 et 80 mm, traités à l'état T851, avec une température d'entrée au laminage à chaud de 475°C, une mise en solution de 2 h à 528°C, et un revenu de 24 h à 173°C, les caractéristiques mécaniques statiques (limite d'élasticité R0.2 et résistance à la rupture Rm en MPa et allongement A en %)) à quart-épaisseur et à mi-épaisseur, dans les sens L et TL. L'ensemble des résultats est reproduit au tableau 9. Ils montrent la faible évolution des propriétés en fonction de l'épaisseur, résultant d'une faible sensibilité à la trempe. e (mm) Prélév. R0,2(L) Rm(L) A(L) R0,2(TL) Rm(TL) A(TL) 15 ½ ép. 400 451 13,6 392 458 12,1 40 ½ ép. 387 439 13,7 376 448 11,2 80 ½ ép. 388 436 11,4 376 443 9,8 80 ¼ ép. 410 466 11,9 467 400 9,7 Samples according to the invention, with a thickness of 15, 40 and 80 mm, treated in the T851 state, were measured with a hot-rolling entry temperature of 475 ° C. and dissolved for 2 hours. at 528 ° C, and a 24 hour yield at 173 ° C, the static mechanical characteristics (yield strength R 0.2 and tensile strength R m in MPa and elongation A in%)) at quarter-thickness and at half -thickness, in the L and TL directions. The overall results are reproduced in Table 9. They show the low evolution of the properties as a function of the thickness, resulting from a low sensitivity to quenching. e (mm) CUPS. R 0.2 (L) R m (L) A (L) R 0.2 (TL) R m (TL) A (TL) 15 ½ ep. 400 451 13.6 392 458 12.1 40 ½ ep. 387 439 13.7 376 448 11.2 80 ½ ep. 388 436 11.4 376 443 9.8 80 ¼ thick 410 466 11.9 467 400 9.7

    Ces tôles sont particulièrement adaptées à la fabrication d'éléments d'intrados d'ailes d'avions par une gamme de fabrication comportant un usinage et une ou plusieurs opérations de mise en forme.These sheets are particularly suitable for the manufacture of wingtip elements. of aircraft by a manufacturing range involving machining and one or more formatting operations.

    Claims (22)

    1. Aircraft structure element, particularly an aircraft wing intrados element, made from a rolled, extruded or forged product, made of an alloy with the following composition (% by weight):
         Cu 4.6 - 5.3; Mg 0.10 - 0.50; Mn 0.25 - 0.45; Si < 0.10; Fe < 0.15; Zn < 0.20; Cr < 0.10
         other elements < 0.05 each and < 0.15 total, remainder Al, solution heat treated, quenched, stress relieved by controlled stretching with a permanent set of more than 1,5% and artificial ageing.
    2. Structure element according to claim 1, characterised in that Si < 0.08%.
    3. Structure element according to claim 1 or 2, characterised in that Fe < 0.09%.
    4. Structure element according to one of claims 1 to 3, characterised in that Cu > 4.8% and preferably > 4.9%.
    5. Structure element according to one of claims 1 to 4, characterised in that Cu > 5%.
    6. Structure element according to one of claims 1 to 5, characterised in that Mg is between 0.20 and 0.40%.
    7. Structure element according to one of claims 1 to 6, characterised in that it has a yield point R0.2 (TL direction) > 350 MPa, preferably > 370 MPa.
    8. Structure element according to one of claims 1 to 7, characterised in that its toughness Klc (L-T direction) > 42 MPa√m.
    9. Structure element according to one of claims 1 to 8, characterised in that it has a type P resistance to intercrystalline corrosion according to standard ASTM G110.
    10. Structure element according to one of claims 1 to 9, characterised in that solution heat treatment is performed at a temperature of less than 10°C less than the temperature at which the alloy begins to melt.
    11. Structure element according to one of claims 1 to 10, characterised in that artificial ageing is carried out at a temperature > 160°C (preferably > 170°C).
    12. Structure element according to one of claims 1 to 11, characterised in that artificial ageing is carried out at the same time as a forming operation.
    13. Structure element according to one of claims 1 to 12, characterised in that the recrystallisation rate throughout its thickness is more than 70%, and preferably more than 90%.
    14. Structure element according to one of claims 1 to 13, characterised in that it forms part of an aircraft wing intrados.
    15. Structure element as claimed in one of claims 1 to 13, characterised in that it is obtained by machining.
    16. Aircraft wing intrados element according to claim 14, characterised in that the skin and stiffeners are obtained by machining of the same initial product.
    17. Structure element according to either claim 15 or 16, characterised in that after machining, it has a deflection f in the L and LT directions such that: fe < 0.10 l2 where f is expressed in µm, e being the thickness of the element and l the length of the bar-shaped test piece in mm.
    18. Process for manufacturing an aircraft structure element comprising:
      a) casting an ingot or billet with the following composition:
         Cu 4.6 - 5.3; Mg 0.10 - 0.50; Mn 0.25 - 0.45; Si < 0.10; Fe < 0.15; Zn < 0.20; Cr < 0.10; other elements < 0.05 each and < 0.15 total, remainder aluminium,
      b) homogenising this ingot or billet,
      c) hot transformation of this ingot by rolling or of this billet by extrusion or forging to obtain a product more than 10 mm thick,
      d) quenching of the hot transformed product,
      e) solution heat treatment of this product, preferably at a temperature of less than 10°C less than the temperature at which the alloy begins to melt,
      f) stress relieving of the product, by controlled tension, until permanent deformation of more than 1.5%,
      g) artificial ageing of the product at a temperature greater than 160°C, possibly combined with forming,
      h) machining of the product, possibly formed, until obtaining the final shape of the structure element.
    19. Process according to claim 18, characterised in that the cast ingot or billet has a Cu content > 4.8%, and preferably > 4.9%.
    20. Process according to either claim 18 or 19, characterised in that the cast ingot or billet has an Mg content between 0.20 and 0.40%.
    21. Process according to one of claims 18 to 20, characterised in that artificial ageing is carried out at a temperature > 170°C.
    22. Process according to one of claims 18 to 21, characterised in that the product is a sheet metal obtained by hot rolling with an inlet temperature at least 40°C (and preferably at least 50°C) less than the solution heat treatment temperature.
    EP00420263A 1999-12-28 2000-12-20 Al-Cu-Mg alloy aircraft structural element Revoked EP1114877B1 (en)

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    FR9916610A FR2802946B1 (en) 1999-12-28 1999-12-28 AL-CU-MG ALLOY AIRCRAFT STRUCTURAL ELEMENT
    FR9916610 1999-12-28

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    US20010006082A1 (en) 2001-07-05
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    US6692589B2 (en) 2004-02-17

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