FR2807449A1 - METHOD FOR MANUFACTURING STRUCTURE ELEMENTS OF ALUMINUM AL-SI-MG ALLOY AIRPLANES - Google Patents

Abstract

The subject of the invention is a process for manufacturing an aircraft structural element, in particular a fuselage element, from rolled, extruded or forged aluminum alloy products of composition (% by weight): Si: 0, 7 - 1, 3 Mg: 0, 6 - 1, 1 Cu: 0, 5 - 1, 1 Mn: 0, 3 - 0, 8 Zn <1 Fe <0, 30 Zr <0, 20 Cr <0 , 25 other elements <0.05 each and <0.15 in total, remainder of aluminum, comprising: - dissolution of the product between 540 and 570degreC, - quenching - realization of the structural element by shaping of the product, and possibly welding, - the tempering of the structural element, in one or more stages, for which the total equivalent time at 175degreC expressed in hours is between (-160 + 57y) and (-184 + 69gamma) , gamma being the sum of the contents in% by weight Si + 2Mg + 2Cu. The invention leads to an improvement in tolerance to damage without loss of other use properties.

Description

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Process for manufacturing Al-Si-Mg aluminum alloy aircraft structural elements
Field of the invention
The invention relates to the field of rolled, extruded or forged aluminum alloy products Al-Si-Mg 6000 series according to the alloy designations of the Aluminum Association, for the manufacture of structural elements of aircraft, including fuselage elements.

 State of the art The fuselages of commercial aircraft are made for the most part from 2024 alloy sheets in the T3 or T351 state, plated on both sides with a lightly loaded aluminum alloy, for example a 1050 alloy or 1070, for the purpose of improving the corrosion resistance. According to the thickness of the core plate, the thickness of the plating may be between 2 and 12% of the total thickness.

For several years, it has been proposed to use, for the fuselage panels, in place of alloy 2024 or neighboring alloys, Al-Si-Mg alloys of the 6000 series. These alloys, also heat-treated, exhibit good mechanical properties in the treated state, a high modulus of elasticity and a lower density than the 2024. This is more easily weldable alloys, which would reduce the number of riveted assemblies, which are a source of additional cost, and also sites of stress concentration and initiation of corrosion.

Si : 0,4 - 1,2 Mg : O5 -- 1,3 Cu : 0,6 - LI Mn : 0. 1 - 1 re < 0,6 Le brevet LP 0173632. au nom de la demanderesse, décrit un alliage, enregistré ultérieurement sous la désignation 6056, de composition : US Patent 4589932 (Alcoa) discloses the use, for aircraft structural elements, of an alloy, subsequently registered under the designation 6013, of composition (% by weight):

If: 0.4 - 1.2 Mg: O5 - 1.3 Cu: 0.6 - LI Mn: 0. 1 - 1 re <0.6 LP 0173632. in the name of the applicant, describes an alloy , subsequently registered under the designation 6056, of composition:

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Si : 0,9-1.2 Mg : 0,7 - I .I Cu : 0,3 - 1,1 Mn : 0.25 -- 0,75 Zen:0.1 - 0.7 Zr : 0.07 - 0.2 Fe < 0,3
Le brevet EP 0787217, également au nom de la demanderesse, concerne un traitement de revenu particulier, conduisant à un état T78, pour un alliage du type
6056. de manière à le désensibiliser à la corrosion intercristalline, et à permettre ainsi son utilisation sans placage pour le fuselage des avions. Ce revenu se définit par une durée totale, mesurée en temps équivalent à 175 C, comprise entre 30 et 300 h, et de préférence entre 70 et 120 h. Ce développement a fait l'objet d'une communication de R. Dif, D. Béchet. T. Warner et H. Ribes : 6056 T78 : A corrosion résistant copper-rich 6xxx alloy for aerospace applications au congrès ICAA-6 (juillet 1998) à Toyohashi (Japon), et publié dans les Proceedings du congrès, pages 1991-1996.

If: 0.9-1.2 Mg: 0.7 - I .I Cu: 0.3 - 1.1 Mn: 0.25 - 0.75 Zen: 0.1 - 0.7 Zr: 0.07 - 0.2 Fe <0.3
EP 0787217, also in the name of the applicant, relates to a particular income treatment, leading to a T78 state, for an alloy of the type
6056. in order to desensitize it to intercrystalline corrosion, and thus to allow its use without plating for the fuselage of aircraft. This income is defined by a total duration, measured in time equivalent to 175 C, between 30 and 300 hours, and preferably between 70 and 120 hours. This development was the subject of a communication by R. Dif, D. Béchet. T. Warner and H. Ribes: 6056 T78: Copper-rich corrosion resistant 6xxx alloy for aerospace applications at the ICAA-6 congress (July 1998) in Toyohashi (Japan), and published in the Congress Proceedings, pages 1991-1996.

The shaping of the pieces is preferably in the T4 state. wherein alloy 6056 exhibits excellent formability. The income is made on the formed parts and possibly welded. The use of 6056 in the T78 state leads to complete desensitization to intercrystalline corrosion of the welded joint or the base product, and to static mechanical characteristics equivalent to those of the plated 2024 T3 or T351. However, it has been found desirable to improve the results obtained with respect to damage tolerance, while maintaining the static mechanical properties and the desensitization to intercrystalline corrosion.

Si : 0,7 - 1,3 Mg : 0,6 - 1,1 Cu : 0.5 - 1,1 Mn : 0,3 - 0,8 Zn < 1 Fe < 0,30 Zr < 0,20 Cr < 0,25 autres éléments < 0,05 chacun et < 0,15 au total, reste aluminium. comportant : - une mise en solution du produit entre 540 et 570 C. une trempe la réalisation de l'élément de structure par mise en forme du produit, et éventuellement soudage. OBJECT OF THE INVENTION The subject of the invention is a method for manufacturing an aircraft structural element from rolled, spun or forged products made of aluminum alloy of composition (% by weight):

If: 0.7 - 1.3 Mg: 0.6 - 1.1 Cu: 0.5 - 1.1 Mn: 0.3 - 0.8 Zn <1 Fe <0.30 Zr <0.20 Cr <0 , 25 other elements <0.05 each and <0.15 in total, remains aluminum. comprising: - dissolving the product between 540 and 570 C. quenching the realization of the structural element by shaping the product, and optionally welding.

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 the income of the structural element, in one or more stages, for which the total equivalent time at 175 C expressed in hours is between (-160 - 57y) and (-184 + 69y). where y is the sum of the contents in% by weight Si + 2Mg + 2Cu.

The invention also relates to a method of manufacturing an aircraft structural element. in which the composition of the products belongs to a preferred composition domain (% by weight):
If: 0.7 - 1.1 Mg: 0.6 - 0.9 Cu: 0.5 - 0.7 Mn: 0.3 - 0.8 Zr <0.2
Fe <0.2 Zn <0.5 Cr <0.25 Mg / Si <1, Si + 2Mg: 2 - 2.6 other elements <0.05 each and 0.15 in total, remains aluminum, and the income has a duration between 40 and 65 hours of total equivalent time at 175 C.

 It also relates to an aircraft fuselage element made from products of the preferred composition indicated above.

DESCRIPTION OF THE INVENTION The invention is based on the observation that within the range of composition and income described in EP 0787217, there is a restricted domain connecting the major elements of the composition (Si, Mg and Cu ) and the total equivalent time at 175 C of the income, as this parameter is defined in EP 0787217, a domain for which, compared with the results disclosed in the examples of this European patent, an improvement in the static mechanical characteristics and in the damage tolerance, without adverse influence on sensitivity to intercrystalline corrosion. It is thus possible to connect to each alloy composition a factor y equal to the sum of the contents (in% by weight) Si + 2Mg + 2Cu. and at this factor y a time range equivalent to 175 C for the income included (in hours) between (-160 + 57 [gamma]) and (-184 + 69y), and preferably between (-150 + 57y) and (-184 + 69y).

More particularly, the inventors have demonstrated that by discharging the alloy with respect to the compositions of the examples of the European patent, that is to say by placing themselves rather in the lower part of the ranges of contents for these 3 elements. while striving for these elements to be dissolved as completely as possible, the alloy became less sensitive to the intercrystalline corrosion at given yield, and consequently, it could be desensitized with a less intense over-income.

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So. in the field of preferred composition mentioned above, with in particular Cu <0. 7% and Si + 2Mg <2.6%. the time equivalent to 175 C of income to reach the T78 state with total desensitization is between 40 and 65 h. either below the preferred range (70 to 120 h) indicated in EP 0787217.

However, to obtain sufficient mechanical strength, it is necessary to maintain Cu> 0.5% and Si + 2Mg> 2.0 and preferably> 2. 3%.

KR (as = 40 mm) > 1 1 mua-vu Kc0 > 80 MPam Kc > 1 10 0 MPaVm Les mesures sont effectuées sur une éprouvette CCT de largeur W = 760 mm et de longueur de fissure initiale 2a0 = 253 mm. L'essai permet de définir la courbe R du matériau, donnant la résistance à la déchirure KR en fonction de l'extension de la fissure Aa. On peut ensuite calculer à partir de cette courbe, et selon la procédure indiquée par L. Schwarmann dans Aluminium, 1991, voi.67, n 5. p. 479, les ténacités apparente Kc0 et effective Kc qui correspondent à la rupture d'une éprouvette virtuelle de type CCT de largeur W = 400 mm et de longueur de fissure initiale 2ao = 133 mm. une ténacité dans le sens L-T, mesurée dans les mêmes conditions que celle dans le sens telle que l'une au moins des propriétés suivantes soit vérifiée :
Kc0 > 90 MPa#m In this area of preferential composition, associated with a revenue T78 equivalent time 175 C between 40 and 65 h. we can get. in addition to complete desensitization to intercrystalline corrosion, the following level of properties in terms of static mechanical characteristics, fracture toughness and speed of propagation: - a yield strength R0.2 (TL direction)> 330 MPa. a breaking strength Rm (TL direction)> 360 MPa and elongation A (TL direction)> 8%. a toughness in plane stress, measured in the direction according to the ASTM standard
E561, such that at least one of the following properties is verified:
KR (Aa = 20 mm)> 90 MPa # m

KR (as = 40 mm)> 11 mua-vu Kc0> 80 MPam Kc> 1 10 0 MPaVm The measurements are made on a CCT test specimen with width W = 760 mm and initial crack length 2a0 = 253 mm. The test makes it possible to define the curve R of the material, giving the tear strength KR as a function of the extension of the crack Aa. We can then calculate from this curve, and according to the procedure indicated by L. Schwarmann in Aluminum, 1991, voi.67, n 5. p. 479, the apparent tenacity Kc0 and effective Kc which correspond to the rupture of a CCT type virtual specimen of width W = 400 mm and initial crack length 2ao = 133 mm. a toughness in the direction LT, measured under the same conditions as in the direction such that at least one of the following properties is verified:
Kc0> 90 MPa # m

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Kc > 130MPam. - une vitesse de propagation de fissures da/dn, mesurée dans le sens T-L selon la norme ASTM E647 pour R = 0,1 sur une éprouvette de type CCT de largeur W = 160 mm, inférieure à :
2 10-3 mm/cycle pour AK = 20 MPam
4 10-3 mm/cycle pour #K = 25 MPam

8 l 0-' mm/cycle pour 1K = 30 MPam
Cet ensemble de propriétés, associé au fait que l'alliage est soudable, le rend particulièrement apteà la fabrication d'éléments de structure d'avions, notamment de fuselage. Il est également possible d'utiliser l'alliage, dans la composition préférentielle de l'invention, à l'état T6.

Kc> 130MPam. a crack propagation rate da / dn, measured in the TL direction according to ASTM E647 standard for R = 0.1 on a CCT type specimen of width W = 160 mm, less than:
2 10-3 mm / cycle for AK = 20 MPam
4 10-3 mm / cycle for #K = 25 MPam

8 l 0 - mm / cycle for 1K = 30 MPam
This set of properties, associated with the fact that the alloy is weldable, makes it particularly suitable for the manufacture of structural elements of aircraft, including fuselage. It is also possible to use the alloy, in the preferred composition of the invention, in the T6 state.

The level of properties obtained at this state T6 with the preferred composition of the invention, in terms of static mechanical characteristics, toughness and crack propagation speed is as follows: a yield strength R0.2 (TL direction) )> 350 MPa, a breaking strength Rm (TL direction)> 380 MPa and an elongation A (TL direction)> 6%.

a toughness in the TL direction, measured under the same conditions as for the T78 state mentioned above, such that at least one of the following properties is verified: KR (#a = 20 mm)> 95 MPam KR (# a = 40 mm)> 120 MPam
KcO> 85 MPaVm
Kc> 115 MPam - a tenacity measured in the LT direction under the same conditions. such that at least one of the following properties is verified:
Kc0> 100 MPam
Kc> 150 MPam.

8 10-' mm/cycle pour 1K = 30 MPaY'm a crack propagation speed da / dn, measured under the same conditions as at the T78 state, less than:
2 10-3 mm / cycle for #K = 20 MPam
4 10-3 mm / cycle for #K = 25 MPaVm

8 10- 'mm / cycle for 1K = 30 MPaY'm

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This set of properties, associated with the weldability of the alloy. makes the product particularly suitable for the manufacture of aircraft fuselage elements.

Examples
Example 1
A composition plate (% by weight) corresponding to Example 3 of Patent EP 0787217 was cast, namely: Si: 0.92 Mg: 0.86 Cu: 0.87 Mn: 0.55 Fe: 0.19 Zn : 0.15 Zr: 0.10, that is to say Mg / Si = 0.93 and Si + 2Mg = 2.64 The plate was homogenized at 530 C, scalped, hot rolled and then cold to 3.2 mm. Samples of the sheet obtained were dissolved at 550 ° C., quenched with water and subjected to an income. For some. the income was from 8 am to
175 C to obtain the state T6. that is, the state corresponding to the maximum mechanical strength; for the others, it was 6 hours at 175 ° C. and then 2 hours at 220 ° C., ie a time equivalent to 175 ° C. for 95 hours, in order to obtain the 1'78 ° state, as indicated in example 3 of the patent. EP 0787217.

Etat 1 Ru, (TL) j Rm (TL) l (fi) - rselisibilité CI ! 1 T6 1364 1408 17 pÔui I 78'304343"8 Non Pour l'état T78, on a mesure également la ténacité par la méthode de la courbe R.

selon la norme ASTM K 561. L'essai, effectué sur une éprouvette de type CCT de The mechanical characteristics in the TL direction were measured, namely the breaking strength Rm (in MPa). the conventional yield strength at 0.2% elongation R0.2 (in MPa), and elongation at break A (in%). as well as sensitivity to intercrystalline corrosion (IC) according to the US military standard MIL-H-6088. Complete desensitization is defined by the absence of corrosion branches more than 5 m long. The results are given in Table 1:
Table 1

State 1 Ru, (TL) 1 Rm (TL) 1 (f) -selectability CI! For the T78 state, the toughness was also measured by the R curve method.

according to ASTM K 561. The test, carried out on a CCT test specimen of

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 width W = 760 mm and central crack length 2a0 = 253 mm. allows to deduce the curve which connects the resistance to the tear KR to the increase of the crack Aa. For the direction 1 '-!,. the value of KR has been reported in Table 2 for crack growth # a = 20 mm and Aa = 40 mm.

Etat KR(T-L) K,,(T- 1.)-1 K,o (1' -L) r K, (T-L) Il Kea (L-T) 1 Kc (L- T) -1 1a--20nun 1a-40nun T78 89.5 tIQ7':S:= 75,2 1059 188.8 1'7,8 On a mesuré également à l'état T78 la vitesse de propagation de fissure de fatigue da/dn dans le sens T-L (en mm/cycle) pour R = 0,1 (rapport entre contrainte minimale et maximale) et pour différentes valeurs de #K (en MPaVm) selon la norme ASTM E 647. Les résultats, obtenus sur éprouvettes de type CCT de largeur W = 160 mm, sont indiqués au tableau 3 :
Tableau 3

[État |AK = 20MPaVm K = 2 IIPa/m AK-30MPa\m T78 10- 3 10- I b,3 10-' Exemple 2 The curve R also allows, for example by the method of L. Schwarmann mentioned above, to determine by calculation the toughness in plane stress Kco (apparent toughness) and Kc (effective toughness), in MPa # m, which correspond to the factors of critical stress intensity for a CCT test specimen, which would have width W = 400 mm and initial crack length 2 ao = 133 mm. The results in the TL and LT directions are also given in Table 2:
Table 2

KR (TL) K ,, (T-1.) - 1 K, o (1 '-L) r K, (TL) II Kea (LT) 1 Kc (L-T) -1 1a - 20nun 1a The fatigue crack propagation rate da / dn in the TL direction (in mm / cycle) was also measured in the T78 state. for R = 0.1 (ratio between minimum and maximum stress) and for different values of #K (in MPaVm) according to ASTM E 647. The results, obtained on CCT type specimens of width W = 160 mm, are indicated in table 3:
Table 3

[Condition | AK = 20MPaVm K = 2 IIPa / m AK-30MPa \ m T78 10- 3 10- I b, 3 10- Example 2

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 A plate of composition included in the preferred composition of the present invention was cast: Si = 0.93 Mg-0.75 Cu = 0.60 Mn = 0.63 Fe 0.10 Zn = 0.16, which corresponds to Mg / Si = 0.81 and Si + 2Mg = 2.43 The plate was transformed under the same conditions as in Example 1, except for the income at the T78 state. A portion of the samples received an income of 6 hours at 175 C and 5 hours at 210 C, a total equivalent time at 175 C of 105 hours, in accordance with the preferential teaching of EP 0787217. Another party suffered a returned from 6 h to 175 C and 13 h to 190 C. a total equivalent time at 175 C of 55 hours, according to the present invention. For the conditions T6 and T78 105 h and 55 h were carried out in the same manner as in Example 1. The results are collated in Tables 4,5 and 6.

Etat ' RO.2 CI'W--JRm (TL) 1 (I'I.) Sensibilité CI f6- 360 397 17,5 Oui T78 (10 h) 30 337 10, Non T78 (5 h) 339 i 3G7 9,2 Non "1 On constate que le revenu à 55 h de temps équivalent améliore nettement la résistance mécanique par rapportà celui à 105 h de temps équivalent, tout en présentant la même désensibilisation à la corrosion intercristalline. Table 4

RO.2 CI'W - JRm (TL) 1 (I'I.) CI sensitivity f6- 360 397 17.5 Yes T78 (10 h) 30 337 10, No T78 (5 h) 339 i 3G7 9 , 2 No "1 It is found that the income at 55 h of equivalent time clearly improves the mechanical resistance compared to that at 105 h of equivalent time, while presenting the same desensitization to intercrystalline corrosion.

'Etat KR(T-lJ ~~KR(T-L) Tkc0 (T-IP K, (1'- Kc0 (L-T) TkL-T) Etat Aa-20mm Aa=40mm (H) ! KoCH) Kco (L-T)I T6 IIOIT-f1262 187,9 -F2T:7--044 -lC551- IT8~1 ()5h J944=-12 6 - ~-Ü'l- -- D 175 =- -6 ==1 1379 =] iT7855h 1 9675 |T25 l69~ \ Ï25J 1 Table 5

KR (T-1) KR (TL) Tkc0 (T-IP K, (1-KcO (LT) TkL-T) State Aa-20mm Aa = 40mm (H)! KOCH) Kco (LT) I T6 IIOIT-f1262 187,9-F2T: 7--044-lC551-IT8 ~ 1 () 5h J944 = -126 - ~ -Ü'l- - D 175 = - -6 == 1 1379 =] iT7855h 1 9675 | T25 l69 ~ \ Ï25J 1

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 On the one hand, it can be seen that, at identical income, the variation in composition between Example 1 and Example 2 leads to an improvement in toughness, whatever the measurement parameter used, and that, on the other hand, to identical composition, income at 55 hours of equivalent time also improves toughness.

,'Htat ~TÂK = 20 MPaVm TaK = 25 MPaVm [AK = 30 MPaVm |T6 uTo"1 TK?" 5 10--' lrG bzz01 3IOJ~-~'j510: ~T78(105h) 10"3 ~ " 2 10"3 " TuT 78 (55 h) L12 1 cr' 13 1 0-3~~~~~~--J----J On constate qu'avec le revenu et la composition préférentielle selon l'invention, il n'y a pas de dégradation de da/dn entre l'état T6 et l'état T78. Table 6

Htat ~ TK = 20 MPaVm TaK = 25 MPaVm [AK = 30 MPaVm | T6 uTo "1 TK?" 5 10-- 'lrG bzz01 3IOJ ~ - ~' j510: ~ T78 (105h) 10 "3 ~" 2 10 "3" TuT 78 (55 hrs) L12 1 cr '13 1 0-3 ~~~~~~ It is noted that with the income and the preferred composition according to the invention, there is no degradation of da / dn between the T6 state and the T78 state.

alliage I e (mm) I Si Mg Cu Mn Fc FZn' |Si+2Mg| (mils) ML', CLI Mil fe 1-i- rf42'''0''"'i'OJ''5'"'0(r".63 0.10 0,16 ! ?,4 > X fïTÏI~~T0J5 I 1T60 | ["Ô.63 L f Ô". 10 f 01 6 1T,43 ! j Li3 4-8 0,91 0,7 6 --tà64 0,59 0.13 0.17 2.43 C ,5-6 0.94 T0,80 0,64 0,6 OJO'(U3! 2.54'"' Les plaques ont été transformées de manière identique à celles des exemptes précédents jusqu'au revenu, à ceci près que. pour les épaisseurs supérieures ou égales à 4,5 mm. indiquées au tableau 7. il n'y a pas eu de laminage à froid. On a effectué

pour tous les échantillons le même revenu 6 h à 175 C - L 13 h à 190 C, soit un temps Example 3 3 alloy plates A, B and C were cast, the compositions of which (by weight%). included in the preferred composition range of the invention, and the final rolling thickness e, are shown in Table 7:
Table 7

alloy I e (mm) I Si Mg Cu Mn Fc FZn '| Si + 2Mg | (mils) ML ', CLI. T0J5 I 1T60 | ["Ô.63 L f Ô". 10 f 01 6 1T, 43! J Li3 4-8 0.91 0.7 6 --tab 64 0.59 0.13 0.17 2.43 C, 5-6 0.94 T0 The plates were transformed identically to those of the previous ones until they reached the yield, except for thicknesses greater than or equal to 4.5. in Table 7. There was no cold rolling.

for all samples the same income 6 h at 175 C - L 13 h at 190 C, a time

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 total equivalent at 175 C of 55 h. The same measurements were made as in the previous examples: static mechanical characteristics (TL direction), sensitivity to intercrystalline corrosion. toughness (T-L direction). and crack propagation velocity (T-L direction). The results are shown in Tables 8, 9 and 10.

Allia(je -ep- Ro2(TL) Rm(TL) A(TIJ' Sensibilité CI Ali iageép. R02 (TL) (TL) CfL) --1 Sens; lA 1.4 mm 337 363 8,3 Non X 3,2 mm "339" "367 92 Non ~B~Tmm 340 369 9.t Non ::- - 1 - - - I :: ---l:: c 4,5mm 337 367 9.4 Non le 6mm ~J31 -- 379 14-~-- Non Tableau 9

Alliage-ép. Kp(T-L) Ki(T-L) ~Tk~(T-L) ' [KâïT) Aa20mm Aa-40mm A 1,4 i-nni 90 122,5 ""85'"""" "p29V il A 1 A mm 190 125 ~86~ 25J B 8 mm 110 134 ~J~:------ ~4.5 mm 98.5 - - ---- ~JI 21 ,5~~~~- J-'- --- - -~u--~J Tableau 10

Alliage-ép. 6.K = 20 MPa--Jm 6.K = 25 MPa--Jm JAK = 30MPaVm 'X 1.4mmÏ3 !0'''20" 5IT07-1 ,,) 2,5 15,- t!\ 3,2 mm 1.1 1 -:- 3T0" '"" 8'KT" ilJ- 8 ,;, ---8 6T - 12,)IOT ""'"10""" - -1 le 4,5 ;"m ---r+O' n - UIF |T3T(r Table 8

Allia (I -ep-Ro2 (TL) Rm (TL) A (TIJ 'Sensitivity IC Aliquot R02 (TL) (TL) CfL) - 1 Sense; IA 1.4 mm 337 363 8.3 Not X 3.2 mm "339""367 92 No ~ B ~ Tmm 340 369 9.t No :: - - 1 - - - I :: --- l :: c 4,5mm 337 367 9.4 No the 6mm ~ J31 - 379 14- ~ - No Table 9

-Alloy thickness. Kp (TL) Ki (TL) ~ Tk ~ (TL) '(Kât) Aa20mm Aa-40mm A 1.4 i-nni 90 122.5 ""85'"""""p29V it A 1 A mm 190 125 ~ 86 ~ 25J B 8 mm 110 134 ~ J ~: ------ ~ 4.5 mm 98.5 - - ---- ~ JI 21, 5 ~~~~ - J -'- --- - - ~ u - ~ J Table 10

-Alloy thickness. 6.K = 20 MPa · m · 6 · K = 25 MPa · Jm · JAK = 30 MPa · m · 1.4 mm · 3 · 20 · 5IT07 · 1 ·) 2.5 15, - 3.2 mm 1.1 1 -: - 3T0 "'""8'KT" ilJ- 8,;, 8 6T - 12,) IOT ""'"10""" - -1 the 4.5; "m - --r + O 'n - UIF | T3T (r

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 It can be seen that, for all the thicknesses, and whether or not there has been cold rolling, the values measured for the static mechanical characteristics and the tenacities are greater than the minimum values indicated above for the T78 state, and the crack propagation rates da / dn are lower than the maximum values indicated above for this same state.

Claims (11)

 If: 0.7 - 1.3 Mg: 0.6 -1.1 Cu: 0.5 - 1.1> 1111: 0.3 - 0.8 Zn <1 Fe <0.30 Zr <0.20 Cr < 0.25 other elements <0.05 each and <0.15 in total, remainder aluminum, comprising: a dissolving of the product between 540 and 570 C, quenching the realization of I element of structure by shaping of the product, and possibly welding, the income of the structural element, in one or more stages, for which the total equivalent time at 175 C expressed in hours is between (-160 + 57y) and (-184 + 69y), where y is the sum of the contents in% by weight Si + 2Mg + 2Cu.

 1. A method of manufacturing an aircraft structural member from rolled, extruded or forged aluminum alloy products of composition (% by weight): 2. Method according to claim 1, characterized in that the total equivalent time at 175 C (in h) is between (-150 + 57y) and (-184 + 69 [gamma]). 3. Method according to one of claims 1 or 2, characterized in that the composition of the products is the following (% by weight):

 If: 0.7 - 1.1 Mg: 0.6 - 0.9 Cu: 0.5-0.7 iVln: 0.3-0.8 Zr <0.2 Fe <0.2 Zn <0.5 Cr <0.25 Mg / If <1, Si + 2Mg: 2.0 - 2.6 other elements <0.05 each and 0.15 in total, remains aluminum. 4. Method according to claim 3, characterized in that Si + 2Mg is between 2.3 and 2.6. 5. Method according to one of claims 3 or 4, characterized in that the total equivalent time of income @ 75 C is between 40 and 65 h. <Desc / Clms Page number 13>  6. Aircraft fuselage element, characterized in that it is made from a rolled product, spun or forged alloy composition (% by weight): If: 0.7 - 1.1 Mg: 0.6 - 0.9 Cu: 0.5 - 0.7 Mn: 0.3 - 0.8 Zr <0.2 Fe <0,2 Zn <0,5 Cr <0,25 Mg / Si <1, Si + 2Mg: 2,0 - 2,6 other elements <0,05 each and 0,15 in total, remain aluminum, put in solution, tempered, formed and returned to the T78 state with a total equivalent time at 175 C between 40 and 65 h. 7. fuselage element according to claim 6, characterized in that Si + 2Mg is between 2.3 and 2.6. 8. fuselage element according to one of claims 6 or 7, characterized in that it has in the TL direction, a yield strength R0.2> 330 MPa, a breaking strength Rm> 360 MPa and a lengthening A> 8%. 9. fuselage element according to one of claims 6 to 8, characterized in that it has a plane stress toughness in the TL direction such that at least one of the following properties is verified: KR (Aa = 20 mm) )> 90 MPa # m KR (Aa = 40 mm)> 115 MPam Kc0> 80 MPam Kc> 110 MPa # m 10. fuselage element according to one of claims 6 to 9, characterized in that it has a plane stress toughness in the direction L-T such that: Kc0> 90 MPam or Kc> 130MPam. 11. fuselage element according to one of claims 6 to 10, characterized in that it has a crack propagation rate da / dn, measured in the direction T-L for R = 0.1, less than: 2 10-3 mm / cycle for AK = 20 MPam 4 10-3 mm / cycle for AK = 25 MPam <Desc / Clms Page number 14>  8 10- mm / cycle for AK = 30 MPam