EP1114877B1 - Strukturelement eines Flugzeugs aus Al-Cu-Mg Legierung - Google Patents

Strukturelement eines Flugzeugs aus Al-Cu-Mg Legierung 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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European Patent Office
Prior art keywords
structure element
element according
temperature
product
alloy
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Revoked
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EP00420263A
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English (en)
French (fr)
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EP1114877A1 (de
Inventor
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.

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Materials Engineering (AREA)
  • Mechanical Engineering (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • Physics & Mathematics (AREA)
  • Thermal Sciences (AREA)
  • Crystallography & Structural Chemistry (AREA)
  • Heat Treatment Of Steel (AREA)

Claims (22)

  1. Strukturelement eines Flugzeugs, insbesondere ein Element der Flügelunterseite eines Flugzeugs, hergestellt aus einem Walz-, Strangpressoder Schmiedeprodukt, aus einer Legierung mit der Zusammensetzung (Gew.-%):
    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
    weitere Elemente jeweils < 0,05 und insgesamt < 0,15, Rest Al,
    behandelt durch Lösungsglühen, Abschrecken, Entspannen durch kontrolliertes Recken mit einer bleibenden Verformung von mehr als 1,5 % und Warmauslagern.
  2. Strukturelement nach Anspruch 1, dadurch gekennzeichnet, dass Si < 0,08%.
  3. Strukturelement nach einem der Ansprüche 1 oder 2, dadurch gekennzeichnet, dass Fe < 0,09 %.
  4. Strukturelement nach einem der Ansprüche 1 bis 3, dadurch gekennzeichnet, dass Cu > 4,8 % und vorzugsweise > 4,9 %.
  5. Strukturelement nach einem der Ansprüche 1 bis 4, dadurch gekennzeichnet, dass Cu > 5 %.
  6. Strukturelement nach einem der Ansprüche 1 bis 5, dadurch gekennzeichnet, dass Mg zwischen 0,20 und 0,40 % liegt.
  7. Strukturelement nach einem der Ansprüche 1 bis 6, dadurch gekennzeichnet, dass es eine Dehngrenze R0,2 (Quer-Längsrichtung) > 350 MPa und vorzugsweise > 370 MPa aufweist.
  8. Strukturelement nach einem der Ansprüche 1 bis 7, dadurch gekennzeichnet, dass es eine Bruchzähigkeit Klc (Längs-Querrichtung) > 42 MPa√m aufweist.
  9. Strukturelement nach einem der Ansprüche 1 bis 8, dadurch gekennzeichnet, dass es eine Beständigkeit gegen interkristalline Korrosion vom Typ P nach der Norm ASTM G110 aufweist.
  10. Strukturelement nach einem der Ansprüche 1 bis 9, dadurch gekennzeichnet, dass das Lösungsglühen bei einer Temperatur erfolgt, die weniger als 10°C unter der Temperatur liegt, bei der die Legierung zu schmelzen beginnt.
  11. Strukturelement nach einem der Ansprüche 1 bis 10, dadurch gekennzeichnet, dass die Warmauslagerung bei einer Temperatur > 160°C (vorzugsweise > 170°C) durchgeführt wird.
  12. Strukturelement nach einem der Ansprüche 1 bis 11, dadurch gekennzeichnet, dass die Warmauslagerung zur gleichen Zeit wie ein Umformvorgang durchgeführt wird.
  13. Strukturelement nach einem der Ansprüche 1 bis 12, dadurch gekennzeichnet, dass es in der gesamten Dicke einen Rekristallisationsgrad größer 70 % und vorzugsweise größer 90 % aufweist.
  14. Strukturelement nach einem der Ansprüche 1 bis 13, dadurch gekennzeichnet, dass es Bestandteil der Flügelunterseite eines Flugzeugs ist.
  15. Strukturelement nach einem der Ansprüche 1 bis 13, dadurch gekennzeichnet, dass es durch Bearbeitung gewonnen wird.
  16. Element der Flügelunterseite eines Flugzeugs nach Anspruch 14, dadurch gekennzeichnet, dass Haut und Versteifungen durch Bearbeitung des gleichen Ausgangsproduktes gewonnen werden.
  17. Strukturelement nach einem der Ansprüche 15 oder 16, dadurch gekennzeichnet, dass es nach erfolgter Bearbeitung eine Durchbiegung f in Längs- und Quer-Längsrichtung fe < 0,10 l2 aufweist, wobei f in µm ausgedrückt ist, e die Dicke des Elementes und I die Länge der stabförmigen Probe in mm ist.
  18. Verfahren zur Herstellung eines Flugzeugstrukturelementes, umfassend:
    a) das Gießen einer Platte bzw. eines Pressbarrens mit der Zusammensetzung:
    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 weitere Elemente jeweils < 0,05 und insgesamt < 0,15, Rest Aluminium,
    b) die Homogenisierung dieser Platte bzw. dieses Pressbarrens,
    c) die Warmumformung dieser Platte durch Walzen bzw. dieses Pressbarrens durch Strangpressen oder Schmieden, um ein Produkt mit einer Dicke größer 10 mm zu erhalten,
    d) das Abschrecken des warmumgeformten Produktes,
    e) das Lösungsglühen dieses Produktes, vorzugsweise bei einer Temperatur, die weniger als 10°C unter der Temperatur liegt, bei der die Legierung zu schmelzen beginnt.,
    f) das Entspannen des Produktes durch kontrolliertes Recken bis zu einer bleibenden Verformung von mehr als 1,5 %,
    g) die Warmauslagerung des Produktes bei einer Temperatur oberhalb 160°C, eventuell in Verbindung mit einer Umformung,
    h) die Bearbeitung des eventuell umgeformten Produktes bis auf die Endform des Strukturelementes.
  19. Verfahren nach Anspruch 18, dadurch gekennzeichnet, dass die gegossene Platte bzw. der gegossene Pressbarren einen Cu-Gehalt > 4,8 % und vorzugsweise > 4,9 % hat.
  20. Verfahren nach Anspruch 18 oder 19, dadurch gekennzeichnet, dass die gegossene Platte bzw. der gegossene Pressbarren einen Mg-Gehalt von 0,20 bis 0,40 % hat.
  21. Verfahren nach einem der Ansprüche 18 bis 20, dadurch gekennzeichnet, dass die Warmauslagerung bei einer Temperatur > 170°C durchgeführt wird.
  22. Verfahren nach einem der Ansprüche 18 bis 21, dadurch gekennzeichnet, dass das Produkt ein Blech ist, das durch Warmwalzen mit einer Eintrittstemperatur gewonnen wird, die um mindestens 40°C (und vorzugsweise mindestens 50°C) unter der Lösungsglühtemperatur liegt.
EP00420263A 1999-12-28 2000-12-20 Strukturelement eines Flugzeugs aus Al-Cu-Mg Legierung Revoked EP1114877B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR9916610 1999-12-28
FR9916610A FR2802946B1 (fr) 1999-12-28 1999-12-28 Element de structure d'avion en alliage al-cu-mg

Publications (2)

Publication Number Publication Date
EP1114877A1 EP1114877A1 (de) 2001-07-11
EP1114877B1 true EP1114877B1 (de) 2005-02-02

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EP00420263A Revoked EP1114877B1 (de) 1999-12-28 2000-12-20 Strukturelement eines Flugzeugs aus Al-Cu-Mg Legierung

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US (2) US6569542B2 (de)
EP (1) EP1114877B1 (de)
DE (1) DE60017868T2 (de)
FR (1) FR2802946B1 (de)

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US7883591B2 (en) 2004-10-05 2011-02-08 Aleris Aluminum Koblenz Gmbh High-strength, high toughness Al-Zn alloy product and method for producing such product
CN102108476A (zh) * 2010-12-28 2011-06-29 重庆市宇一机械有限公司 一种高强高韧铝合金航空安全件改性制备方法
US8002913B2 (en) 2006-07-07 2011-08-23 Aleris Aluminum Koblenz Gmbh AA7000-series aluminum alloy products and a method of manufacturing thereof
US8608876B2 (en) 2006-07-07 2013-12-17 Aleris Aluminum Koblenz Gmbh AA7000-series aluminum alloy products and a method of manufacturing thereof

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US7604704B2 (en) * 2002-08-20 2009-10-20 Aleris Aluminum Koblenz Gmbh Balanced Al-Cu-Mg-Si alloy product
US7323068B2 (en) * 2002-08-20 2008-01-29 Aleris Aluminum Koblenz Gmbh High damage tolerant Al-Cu alloy
US7494552B2 (en) * 2002-08-20 2009-02-24 Aleris Aluminum Koblenz Gmbh Al-Cu alloy with high toughness
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FR2848480B1 (fr) 2002-12-17 2005-01-21 Pechiney Rhenalu Procede de fabrication d'elements structuraux par usinage de toles epaisses
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US20050034794A1 (en) * 2003-04-10 2005-02-17 Rinze Benedictus High strength Al-Zn alloy and method for producing such an alloy product
CA2519390C (en) 2003-04-10 2015-06-02 Corus Aluminium Walzprodukte Gmbh An al-zn-mg-cu alloy
US7666267B2 (en) * 2003-04-10 2010-02-23 Aleris Aluminum Koblenz Gmbh Al-Zn-Mg-Cu alloy with improved damage tolerance-strength combination properties
US8043445B2 (en) * 2003-06-06 2011-10-25 Aleris Aluminum Koblenz Gmbh High-damage tolerant alloy product in particular for aerospace applications
DE10332003B3 (de) * 2003-07-14 2004-12-16 Eads Deutschland Gmbh Geschweißtes Aluminium-Strukturbauteil mit Aluminium-Guss-Werkstoffelementen
FR2858984B1 (fr) * 2003-08-19 2007-01-19 Corus Aluminium Walzprod Gmbh Produit en alliage ai-cu a haute tenacite et son procede de production
US20060032560A1 (en) * 2003-10-29 2006-02-16 Corus Aluminium Walzprodukte Gmbh Method for producing a high damage tolerant aluminium alloy
US20050098245A1 (en) * 2003-11-12 2005-05-12 Venema Gregory B. Method of manufacturing near-net shape alloy product
US20070204937A1 (en) * 2005-07-21 2007-09-06 Aleris Koblenz Aluminum Gmbh Wrought aluminium aa7000-series alloy product and method of producing said product
US20070151636A1 (en) * 2005-07-21 2007-07-05 Corus Aluminium Walzprodukte Gmbh Wrought aluminium AA7000-series alloy product and method of producing said product
CN101297054A (zh) * 2005-10-25 2008-10-29 阿勒里斯铝业科布伦茨有限公司 适用于航空航天应用的Al-Cu-Mg合金
US20070151637A1 (en) * 2005-10-28 2007-07-05 Aleris Aluminum Koblenz Gmbh Al-Cu-Mg ALLOY SUITABLE FOR AEROSPACE APPLICATION
WO2010081889A1 (en) 2009-01-16 2010-07-22 Aleris Aluminum Koblenz Gmbh Method for the manufacture of an aluminium alloy plate product having low levels of residual stress
US9314826B2 (en) 2009-01-16 2016-04-19 Aleris Rolled Products Germany Gmbh Method for the manufacture of an aluminium alloy plate product having low levels of residual stress
US9123930B1 (en) 2011-04-29 2015-09-01 Greatbatch Ltd. Dual glass to metal seal cell
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FR3011252B1 (fr) * 2013-09-30 2015-10-09 Constellium France Tole d'intrados a proprietes de tolerance aux dommages ameliorees
FR3040711B1 (fr) * 2015-09-03 2017-08-11 Constellium Issoire Produit extrude en alliage al-cu-mg a compromis ameliore entre resistance mecanique et tenacite
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CN115786787B (zh) * 2022-07-18 2024-02-23 山东浩信机械有限公司 一种高强韧Al-Cu系铸造铝合金及其制备方法

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US6692589B2 (en) 2004-02-17
DE60017868D1 (de) 2005-03-10
FR2802946B1 (fr) 2002-02-15
US6569542B2 (en) 2003-05-27
DE60017868T2 (de) 2005-12-29
US20010006082A1 (en) 2001-07-05
EP1114877A1 (de) 2001-07-11
FR2802946A1 (fr) 2001-06-29

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