Al-si-mg aluminum alloy aircraft structural component production method

Abstract

An aircraft structural component, particularly a fuselage component, production method using rolled, extruded or forged products made of aluminum alloy with the following composition (% by weight):

Description

FIELD OF THE INVENTION

[0001] The invention relates to the field of rolled, extruded or forged products made of series 6000 Al—Si—Mg aluminum alloy according to Aluminum Association references, intended to produce aircraft structural components, particularly fuselage components.

[0002] 2. Description of Related Art

[0003] Commercial aircraft fuselages are mostly produced using 2024 T3 or T351 alloy sheets, clad on both sides with a low content aluminum alloy, for example a 1050 or 1070 alloy, in order to improve corrosion resistance. The thickness of the cladding may represent, depending on the thickness of the core sheet, between 2 and 12% of the total thickness.

[0004] Several years ago, it was proposed to use series 6000 Al—Si—Mg alloys for fuselage panels instead of 2024 alloys or similar alloys. These alloys, also heat-treated, offer good mechanical characteristics when treated, a high modulus of elasticity and a lower density than 2024. The alloys are easier to weld, enabling a reduction in the number of riveted assemblies, which are a source of additional cost, and also sites in which stress is concentrated and corrosion initiated.

[0005] The patent U.S. Pat. No. 4,589,932 (Alcoa) discloses the use, for aircraft structural components, of an alloy subsequently registered under the reference 6013, with the following composition (% by weight):

[0006] Si: 0.4-1.2 Mg: 0.5-1.3 Cu: 0.6-1.1 Mn: 0.1-1 Fe<0.6

[0007] The patent EP 0 173 632, filed by the applicant, discloses an alloy, subsequently registered under the reference 6056, with the following composition:

[0008] Si: 0.9-1.2 Mg: 0.7-1.1 Cu: 0.3-1.1 Mn: 0.25-0.75 Zn: 0.1-0.7 Zr: 0.07-0.2 Fe<0.3

[0009] The patent EP 0 787 217, also filed by the applicant, relates to a specific ageing treatment, resulting in a T78 temper, for a 6056 type alloy, so as to desensitize it to intercrystalline corrosion, and thus enable its use without cladding for aircraft fuselages. This ageing is defined by a total duration, measured in equivalent time at 175° C., between 30 and 300 hours, and preferentially between 70 and 120 hours. This development was the subject of a presentation by R. Dif, D. Béchet, T. Warner and H. Ribes: “6056 T78: A corrosion resistant copper-rich 6xxx alloy for aerospace applications” at the ICAA-6 congress (July 1998) in Toyohashi (Japan) and published in the Congress Proceedings, pages 1991-1996.

[0010] The parts are preferentially shaped in the T4 temper, wherein the 6056 alloy shows excellent forming properties. The ageing is performed on shaped and possibly welded parts. The use of 6056-T78 results in complete desensitisation to intercrystalline corrosion of the welded join or the base product and in static mechanical characteristics equivalent to those of clad T3 or T351 2024. However, it seemed desirable to improve the results obtained in terms of tolerance to damage, while retaining the static mechanical properties and desensitization to intercrystalline corrosion.

SUMMARY OF THE INVENTION

[0011] The invention relates to an aircraft structural component production method using rolled, extruded or forged products made of aluminum alloy comprising:

[0012] the casting of a blank with the following composition (% by weight):

[0013] 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, the remainder being aluminum,

[0014] hot, and possibly cold, transformation of said blank to produce a product,

[0015] solution heat treatment of the product between 540 and 570° C.,

[0016] quenching of said product,

[0017] production of the structural component by forming of the product, and welding if required,

[0018] ageing of the structural component, in one or more stages, for which the total equivalent time at 175° C. expressed in hours is between (−160+57γ) and (−184+69γ), γ being the sum of the Si+2Mg+2Cu contents in % by weight.

[0019] The invention also relates to an aircraft structural component production method, wherein the composition of the products belongs to a preferential composition range (% by weight):

[0020] Si: 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+2 Mg: 2-2.6 other elements <0.05 each and <0.15 in total, the remainder being aluminum,

[0021] and the ageing time is between 40 and 65 hours of total equivalent time at 175° C.

[0022] It also relates to an aircraft fuselage component produced using products with the preferential composition given above.

BRIEF DESCRIPTION OF THE DRAWINGS

[0023] FIG. 1 represents, in the form of Wöhler curves, the fatigue service life of T6 and T78 temper samples according to Example 1, before and after prolonged exposure in a marine environment.

[0024] FIG. 2 represents the results of intercrystalline corrosion tests as a function of the yield strength in the TL direction in the T4 temper for the samples in Examples 6 and 7.

DETAILED DESCRIPTION OF THE INVENTION

[0025] The invention is based on the observation that within the composition and ageing range disclosed in the patent EP 0 787 217, there is a restricted range linking the major elements of the composition (Si, Mg and Cu) and the total equivalent ageing time at 175° C., as this parameter is defined in EP 0 787 217; with this range, in relation to the results given in the examples of this European patent, an improvement in the static mechanical characteristics and tolerance to damage is obtained, with no adverse effect on sensitivity to intercrystalline corrosion. It is thus possible to associate with each alloy composition a factor γ equal to the sum of the Si+2Mg+2Cu contents (in % by weight) and with said factor γ a period of equivalent ageing time at 175° C. between (in hours) (−160+57γ) and (−184+69γ) and preferentially between (−150+57γ) and (−184+69γ).

[0026] More specifically, the inventors revealed that by unloading the alloy in relation to compositions of the examples in the European patent, i.e. by positioning at the lower end of the content ranges for these 3 elements, while ensuring that these elements are put in solution as completely as possible, the alloy became less sensitive to intercrystalline corrosion at given over-ageing and that, as a result, it was possible to desensitize it with a lower level over-ageing.

[0027] In this way, in the preferential composition range mentioned above, with particularly Cu<0.7% and Si+2Mg<2.6%, the equivalent ageing time at 175° C. to reach the T78 temper with total desensitization is between 40 and 65 hours, i.e. below the preferential range (70 to 120 hours) indicated in the patent EP 0 787 217. However, to obtain a sufficient mechanical resistance, it is necessary to maintain Cu>0.5% and Si+2Mg>2.0 and preferentially >2.3%.

[0028] In this preferential composition range, associated with T78 ageing for an equivalent time at 175° C. between 40 and 65 hours, it is possible to obtain, in addition to complete desensitization to intercrystalline corrosion, the following level of properties in terms of static mechanical characteristics, toughness and crack growth:

[0029] a yield strength R0.2 (TL direction)>330 MPa, an ultimate tensile strength Rm (TL direction)>360 MPa and an elongation A (TL direction)>8%.

[0030] a plane strain fracture toughness, measured in the T-L direction, according to the ASTM E561 standard, such that at least one of the following properties is verified:

[0031] KR (Δa=20 mm)>90 MPa{square root}m

[0032] KR (Δa=40 mm)>115 MPa{square root}m

[0033] Kc0>80 MPa{square root}m

[0034] Kc>110 MPa{square root}m

[0035] The measurements are made on a CCT test specimen of width W=760 mm and initial cracking length 2a0=253 mm. The test makes it possible to define the R curve of the material, giving the tear strength KR as a function of the crack extension Δa. Using this curve, it is then possible to calculate, according to the procedures indicated by L. Schwarmann in Aluminium, 1991, vol. 67, No. 5, p. 479, the apparent fracture toughness Kc0 and effective fracture toughness Kc which correspond to the break of a virtual CCT type test specimen of width W=400 mm and initial crack length 2a0=133 mm.

[0036] a fracture toughness in the L-T direction, measured under the same conditions as that in the T-L direction, such that at least one of the following properties is verified:

[0037] Kc0>90 MPa{square root}m

[0038] Kc>130 MPa{square root}m

[0039] a crack growth rate da/dn, measured in the T-L direction according the ASTM E647 standard for R=0.1 on a CCT type test specimen of width W=160 mm, less than:

[0040] 2 10−3 mm/cycle for ΔK=20 MPa{square root}m

[0041] 4 10−3 mm/cycle for ΔK=25 MPa{square root}m

[0042] 8 10−3 mm/cycle for ΔK=30 MPa{square root}m

[0043] Finally, in this particular T78 temper, a lower drop in the fatigue resistance after prolonged exposure in a corrosive environment is observed in relation to the T6 temper.

[0044] This set of properties, associated with the fact that the alloy can be welded, makes it particularly suitable for the production of aircraft structural components, particularly fuselage components.

[0045] It is also possible to use the alloy, in the preferential composition according to the invention, in the T6 temper. The level of properties obtained in said T6 temper with the preferential composition according to the invention, in terms of static mechanical characteristics, fracture toughness and crack growth rate is as follows:

[0046] a yield strength R0.2 (TL direction)>350 MPa, an ultimate tensile strength Rm (TL direction)>380 MPa and an elongation A (TL direction)>6%.

[0047] a fracture toughness, measured under the same conditions as for the T78 temper mentioned above, such that at least one of the following properties is verified:

[0048] KR (Δa=20 mm)>95 MPa{square root}m

[0049] KR (Δa=40 mm)>120 MPa{square root}m

[0050] Kc0>85 MPa{square root}m

[0051] Kc>115 MPa{square root}m

[0052] a fracture toughness measured in the L-T direction under the same conditions, such that at least one of the following properties is verified:

[0053] Kc0>100 MPa{square root}m

[0054] Kc>150 MPa{square root}m

[0055] a crack growth rate da/dn, measured under the same conditions as for the T78 temper, less than:

[0056] 2 10−3 mm/cycle for ΔK=20 MPa{square root}m

[0057] 4 10−3 mm/cycle for ΔK=25 MPa{square root}m

[0058] 8 10−3 mm/cycle for ΔK=30 MPa{square root}m

[0059] This set of properties, associated with the fact that the alloy can be welded, makes the product particularly suitable for the production of aircraft fuselage components.

[0060] The production method according to the invention comprises the casting of a blank made of the composition mentioned, said blank being a plate for rolled products, a slug for extruded products or a forging block for forged products. The blank is scalped and then heated before hot transformation by rolling, extrusion or forging, and possibly undergoes cold transformation. After cutting in the required format, the semi-finished product obtained undergoes a heat treatment at a temperature between 540 and 570° C., quenched, generally in cold water, and finished, the purpose of said final step essentially being to absorb the deformations of the semi-finished product after quenching. The product is most frequently supplied in this T4 temper to shape the structural component and for welding if required. The formed, and if applicable welded, component then undergoes the ageing treatment according to the invention.

[0061] The applicant observed that it is advantageous to add, before scalping, a homogenization step at a temperature between 540 and 570°. Said homogenization may comprise a single stage, or two stages, the second stage being at a higher temperature than the first. It helps improve the forming properties of the product in the T4 temper and reduce the grain size, leading to a decrease in the roughness of the metal when it undergoes chemical machining. Excessive roughness may induce initial micro-cracking due to fatigue.

[0062] In addition, the tests demonstrated that the desensitization to intercrystalline corrosion increases in effectiveness with the level of cold-working in the T4 temper. This cold-working may be a result of finishing operations such as straightening or planing with rollers or traction for sheets and traction or drawing for profiles. It may also be the result of part forming operations by rolling, drawing-forming, embossing, flow turning or folding. Said cold-working, of at least 1%, and preferably of at least 2% permanent elongation, may be relatively significant, for example up to 10%, or even up to 15% permanent elongation; indeed, it is observed, surprisingly, that significant cold-working, although it accelerates the ageing kinetics, does not reduce the yield strength in the T78 temper with reference to the same non-cold-worked product.

[0063] This possibility to use significant cold-working improving the resistance to intercrystalline corrosion is particularly useful if, as is frequently the case for aircraft fuselage production, thin sheets and profiles must be assembled. Indeed, the applicant observed that desensitization to intercrystalline corrosion is more difficult to carry out on profiles than on sheets, probably due to the difference in their metallurgic structure. If the sheets and profiles are formed separately, and then welded before ageing, this is liable to induce a difference in corrosion resistance between the sections produced from profiles and those produced from sheets. To remedy this disadvantage, rather than choose a very high level of ageing to desensitize the profiles, which would induce a significant loss of mechanical resistance, it is preferable to retain the T78 ageing adapted to the desensitization of the sheets and to subject the profiles to an additional cold-working step to bring their resistance to intercrystalline corrosion to the same level as those of thin sheets.

EXAMPLES

Example 1

[0064] A plate of the composition (% by weight) corresponding to Example 3 of the patent EP 0 787 217 was cast, i.e.: Si: 0.92 Mg: 0.86 Cu: 0.87 Mn: 0.55 Fe: 0.19 Zn: 0.15 Zr: 0.10 i.e. Mg/Si=0.93 and Si+2Mg=2.64.

[0065] The plate was heated at 530° C., scalped, hot and then cold rolled to a thickness of 3.2 mm. Samples of the sheet obtained were subjected to a solution heat treatment at 550° C., quenched in water, finished and subjected to ageing. In some cases, the ageing lasted 8 hours at 175° C. to obtain the T6 temper, i.e. the temper corresponding to maximum mechanical resistance; in other cases, it lasted 6 hours at 175° C. and then 2 hours at 220° C., or an equivalent time at 175° C. of 95 hours, to obtain the T78 temper, as described in Example 3 of the patent EP 0 787 217.

[0066] The mechanical characteristics were measured in the TL direction, i.e. the tensile strength Rm (in MPa), the conventional yield strength at 0.2% elongation R0.2 (in MPa) and the fracture elongation A (in %), along with the sensitivity to intercrystalline corrosion IC according to the US Army standard MIL-H-6088. Complete desensitization is defined as the absence of corrosion ramifications over 5 μm long. The results are given in Table 1.

	TABLE 1
					IC
	Temper	R0.2 (TL)	Rm (TL)	A (TL)	sensitivity
	T6	364	408	7	Yes
	T78	304	343	8	No

[0067] For the T78 temper, the fracture toughness was also measured using the R curve method, according to the ASTM E 561 standard. The test, performed on a CCT type test specimen of width W=760 mm and central cracking length 2a0=253 mm, is used to deduce the curve linking the tear strength KR to the increase in cracking Δa. For the T-L direction, the value of KR for increases in cracking Δa=20 mm and Δa=40 mm is given in Table 2.

[0068] The R curve is also used, for example using L. Schwarmann's method mentioned above, to determine by calculation the plane strain toughnesses Kc0 (apparent toughness) and Kc (effective toughness), in MPa{square root}m, which correspond to the critical stress intensity factors for a CCT test specimen, with a width W=400 mm and initial cracking length 2a0=133 mm. The results in the T-L and L-T directions are also given in Table 2:

TABLE 2
	KR (T-L)	KR (T-L)	Kc0	Kc	Kc0	Kc
Temper	Δa = 20 mm	Δa = 40 mm	(T-L)	(T-L)	(L-T)	(L-T)
T78	89.5	107.5	75.2	105.9	88.8	137.8

[0069] The fatigue crack growth was also measured in the T78 temper in the T-L direction (in mm/cycle) for R=0.1 (ratio between minimal and maximal stress) and for different values of ΔK (in MPa{square root}m) according to the ASTM E 647 standard. The results, obtained on CCT type test specimens of width W=160 mm, are given in Table 3:

TABLE 3
Temper	ΔK = 20 MPa{square root}m	ΔK = 25 MPa{square root}m	ΔK = 30 MPa{square root}m
T78	10−3	3 10−3	6.3 10−3

Example 2

[0070] A plate of a composition included in the preferential 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.

[0071] The plate was transformed under the same conditions as in Example 1, except in terms of the ageing in the T78 temper. Part of the samples underwent ageing for 6 hours at 175° C. followed by 5 hours at 210° C., or a total equivalent time at 175° C. of 105 hours, according to the preferential disclosure in the patent EP 0787217. Another part underwent ageing for 6 hours at 175° C. followed by 13 hours at 190° C., or a total equivalent time at 175° C. of 55 hours, according to the present invention. The same measurements as in Example 1 were made for the T6 and T78 tempers at 105 hours and 55 hours. The results are given in Tables 4, 5 and 6.

	TABLE 4
					IC
	Temper	R0.2 (TL)	Rm (TL)	A (TL)	sensitivity
	T6	360	397	7.5	Yes
	T78	305	337	10.5	No
	(105 hrs)
	T78	339	367	9.2	No
	(55 hrs)

[0072] It is observed that ageing for 55 hours equivalent time improves mechanical resistance significantly in relation to that for 105 hours equivalent time, while showing the same desensitization to intercrystalline corrosion.

TABLE 5
	KR (T-L)	KR (T-L)	Kc0	Kc	Kc0	Kc
Temper	Δa = 20 mm	Δa = 40 mm	(T-L)	(T-L)	(L-T)	(L-T)
T6	101.1	126.2	87.9	121.7	104.4	155.1
T78	94.4	119.6	83.1	117.5	91.6	137.9
105 hrs
T78	96.5	125	86.9	125.7
55 hrs

[0073] It is observed firstly that, with the same ageing, the variation in composition between Example 1 and Example 2 results in an improvement in fracture toughness, irrespective of the measurement parameter used and secondly that, with the same composition, the ageing for 55 hours equivalent time also improves the toughness.

TABLE 6
Temper	ΔK = 20 MPa{square root}m	ΔK = 25 MPa{square root}m	ΔK = 30 MPa{square root}m
T6	1.2 10−3	3 10−3	5 10−3
T78 (105 hrs)	10−3	2 10−3	4 10−3
T78 (55 hrs)	1.2 10−3	3 10−3	5 10−3

[0074] It is observed that with the ageing and preferential composition according to the invention, there is no degradation of da/dn between T6 and T78.

[0075] On the same sheets in the T6 and T78 temper, fatigue specimen blanks were removed and exposed for one year to a marine environment on the Mediterranean coast. After machining, the test specimens, showing a strain concentration factor of almost 1, underwent fatigue-endurance tests, to determine the number of fracture cycles, at different levels of strain and a frequency of 30 Hz, for a load ratio R=0.1. The results are represented in FIG. 1 in the form of Wöhler curves, both on the non-corroded material (solid lines) and on the corroded test specimens (individual dots).

[0076] These results demonstrate the advantage of the T78 treatment with reference to the T6 treatment in terms of the drop in fatigue resistance after exposure to corrosion.

Example 3

[0077] Three plates made of alloys A, B and C were cast, for which the compositions (by weight %) within the preferential composition range according to the invention and the final rolling thicknesses e, are given in Table 7:

TABLE 7
Alloy	E (mm)	Si	Mg	Cu	Mn	Fe	Zn	Si + 2 Mg
A	1.4-3.2	0.93	0.75	0.60	0.63	0.10	0.16	2.43
B	4-8	0.91	0.76	0.64	0.59	0.13	0.17	2.43
C	4.5-6	0.94	0.80	0.64	0.56	0.10	0.13	2.54

[0078] The plates were transformed in the same way as those in the above examples up to the ageing step, apart from the fact that, for thicknesses greater than or equal to 4.5 mm, given in Table 7, no cold rolling was carried out. The same ageing for 6 hours at 175° C.+13 hours at 190° C., or a total equivalent time at 175° C. of 55 hours, was performed for all the samples. The same measurements as in the above examples were made: static mechanical characteristics (TL direction) R0.2 (in MPa), Rm (in MPa) and A (in %), sensitivity to intercrystalline corrosion, fracture toughness (T-L direction) and crack growth rate (T-L direction). The results are given in Tables 8, 9 and 10.

	TABLE 8
					IC
	Alloy - th	R0.2 (TL)	Rm (TL)	A (TL)	sensitivity
	A 1.4 mm	337	363	8.3	No
	A 3.2 mm	339	367	9.2	No
	B 4 mm	340	369	9.1	No
	B 8 mm	345	371	8.9	No
	C 4.5 mm	337	367	9.4	No
	C 6 mm	351	379	9.4	No

[0079]

TABLE 9
	KR (T-L)	KR (T-L)
Alloy - th	Δa = 20 mm	Δa = 40 mm	Kc0 (T-L)	Kc (T-L)
A 1.4 mm	90	122.5	85.5	129.7
A 3.2 mm	95.5	125	86.9	125.7
B 8 mm	110	134	93.8	126.1
C 4.5 mm	98.5	121.5	84.9	114.7

[0080]

TABLE 10
Alloy - th	ΔK = 20 MPa{square root}m	ΔK = 25 MPa{square root}m	ΔK = 30 MPa{square root}m
A 1.4 mm	1.3 10−3	2.5 10−3	5.2 10−3
A 3.2 mm	1.1 10−3	3 10−3	4.8 10−3
B 8 mm	8 10−4	2.3 10−3	4.1 10−3
C 4.5 mm	1.1 10−3	2.8 10−3	4.3 10−3

[0081] It is observed that, for all the thicknesses, irrespective of whether cold rolling was performed or not, the values measured for the static mechanical characteristics and toughnesses are greater than the minimum values given above for the T78 temper, and the crack growths da/dn are less than the maximum values given above for the same temper.

Example 4

[0082] An alloy with the following composition (% by weight) was cast: Si=1.01 Mg=0.71 Cu=0.67 Mn=0.55 Fe=0.14 Zn=0.15 remainder aluminum.

[0083] A first plate of this alloy was subjected to the production procedure A comprising the following steps: homogenization for 4 hours at 540° C.+24 hours at 565° C., scalping, heating at 530° C., hot rolling of a strip up to 4.5 mm, conversion of strip into sheets, solution heat treatment in a furnace for 40 min at 550° C. in air, water quenching, finishing, T6 ageing for 8 hours at 175° C.

[0084] A second plate underwent production procedure B comprising the same steps except for the preliminary homogenization. The grain size (thickness e and length l) on the surface and mid-thickness of the sheet was measured in the T4 temper (before ageing) by optical microscopy on a ground section, along with the distribution of the Al—Mn—Si dispersoids in electron microscopy in transmission. This distribution is evaluated with the parameter ECD (Equivalent Circle Diameter)={square root}4A/π wherein A in the area of the phases observed on the microscopic section. To characterize the formability, the parameter LDH (Limit Dome Height) is used. This parameter is defined in the patent application EP 1045043 filed by the applicant. The results are given in Table 11:

TABLE 11
	e grain	1 grain	e grain	1 grain
	surface	surface	mid-th.	mid-th.	ECD	LDH
Procedure	(μm)	(μm)	(μm)	(μm)	(nm)	(mm)
A	27	143	23	140	271	92
B	40	316	30	320	108	73

[0085] It is observed that, in the T4 temper, i.e. in the temper in which sheets are most frequently delivered to the aeronautic manufacturer for forming, ageing, procedure A with homogenization results in a smaller grain size and therefore lower roughness after chemical machining and improved formability.

[0086] The static mechanical characteristics R0.2 (in MPa), Rm (in MPa) and A (in %) in the L and TL directions in the T6 temper were also compared for both procedures. The results are given in Table 12:

TABLE 12
	R0.2
Procedure	(TL)	Rm (TL)	A (TL)	R0.2 (L)	Rm (L)	A (L)
A	361	390	11.3	374	386	12.0
B	359	389	10.5	367	386	12.7

[0087] It is possible to conclude that homogenization does not have a significant effect on the mechanical characteristics in the T6 temper.

Example 5

[0088] An alloy with the following composition (% by weight) was cast: Si=0.82 Mg=0.68 Cu=0.55 Mn=0.57 Fe=0.13 Zn=0.14 remainder aluminum.

[0089] This plate alloy was subjected to the following production procedure: homogenization for 4 hours at 540° C.+24 hours at 565° C., scalping, heating at 530° C., hot rolling of a strip up to 5 mm, conversion of strip into sheets, solution heat treatment in a furnace for 40 min at 550° C. in air, water quenching, finishing, T78 ageing for 6 hours at 175° C.+13 hours at 190° C. (or 55 hours of equivalent time at 175° C.).

[0090] The static mechanical characteristics R0.2, Rm (in MPa) and A (in %) in the TL direction were measured in said T78 temper, along with the fracture toughness in the T-L direction (in MPa{square root}m), the crack growth rate in the T-L direction and the sensitivity to intercrystalline corrosion in the same way as for Examples 1, 2 and 3. The results are given in Tables 13, 14 and 15:

	TABLE 13
					IC
	Procedure	R0.2 (TL)	Rm (TL)	A (TL)	sensitivity
	Homog. +	337	359	11	No
	T78

[0091]

TABLE 14
	KR (T-L)	KR (T-L)
Temper	Δa = 20 mm	Δa = 40 mm	Kc0 (T-L)	Kc (T-L)
Homog. +	115	142	98.8	136
T78

[0092]

TABLE 15
Temper	ΔK = 20 MPa{square root}m	ΔK = 25 MPa{square root}m	ΔK = 30 MPa{square root}m
Homog. + T78	1.1 10−3	2.1 10−3	4.0 10−3

[0093] If these results are compared to those in Table 4 of Example 2, it is noted that, also in the T78 temper, homogenization does not have a significant effect on mechanical characteristics, the crack growth rate or the sensitivity to intercrystalline corrosion, but seems to increase the fracture toughness measured by the R curve.

Example 6

[0094] Samples were taken from the sheets of Example 3 and Example 5 at different thicknesses and with different types of finishing, comprising at least one of the straightening D, roller planning P or traction planing T operations. In each case, the yield strength R0.2 in the TL direction (in MPa) in the T4 and T78 temper was measured, along with the sensitivity to intercrystalline corrosion in the T78 temper. This corrosion was qualified as “slight” when it induces pitting with short intergranular ramifications. The results are given in Table 16:

TABLE 16
Alloy			R0.2 (TL)
ex.	e (mm)	Finish	T4	R0.2 T78	IC sens.
3 A	1.4	D + P + T	218	337	No
3 A	3.2	D + P + T	215	339	No
3 B	4	D + P + T	218	340	No
3 B	8	T	181	345	Slight
3 C	4.5	D + P + T	203	337	No
3 C	6	T	198	351	Slight
5	2.2	D + P + T	179	340	Yes
5	2.2	D + P + T	211	336	Slight
5	2.5	D + P + T	224	332	No
5	2.5	D + P + T	200	330	Slight
5	3.2	D + P + T	245	326	No
5	5	P + T	218	337	No

[0095] The results given in FIG. 2 show, for a given composition, a clear correlation between the resistance to intercrystalline corrosion in the T78 temper and the yield strength in the T4 temper.

Example 7

[0096] Using the sheet corresponding to the seventh row of Table 14 (composition according to Example 5, thickness 2.2 mm), different finishing operations were carried out in the laboratory in the T4 temper consisting of controlled traction to 3.2% permanent elongation and cold rolling to different levels of permanent elongation between 2.6 and 8.7%. The samples obtained in this way were subjected firstly to ageing A for 6 hours at 175° C.+13 hours at 190° C., corresponding to a T78 temper with 55 hours of equivalent time at 175° C. and secondly ageing B for 6 hours at 175° C.+6 hours at 190°, in a slightly over-aged temper with an equivalent time at 175° C. of 31 hours, making it possible exacerbate the sensitivity to intercrystalline corrosion. The yield strength (in MPa) and the elongation (in %) in the TL direction in the T4 temper and the yield strength in the TL direction in the T78 temper after ageing A (Table 17) were measured, along with the sensitivity to intercrystalline corrosion for the ageing steps A and B, indicating the depth of the corrosion (in μm) and extent of the corrosion as a % of the surface affected (Table 18).

	TABLE 17
	Finishing	R0.2 (T4-TL)	R0.2 (T78-TL)	A (T4-TL)
	None	191	337	24.6
	3.2% traction	234	330	21.7
	2.6% rolling	235	333	21.4
	3.5% rolling	236	332	21.1
	5.3% rolling	261	336	18.7
	8.7% rolling	285	340	16.4

[0097]

TABLE 18
	Ageing A		Ageing B
	IC	Depth	IC	Depth
Finishing	sensitivity	Extent	sensitivity	Extent
None	Yes	190 μm-10%	Yes	150 μm-20%
3.2%	Slight	10 μm	Yes	140 μm-1%
traction
2.6%	Slight	10 μm	Yes	190 μm-5%
rolling
3.5%	No	—	Yes	125 μm-10%
rolling
5.3%	No	—	Yes	25 μm-1%
rolling
8.7%	No	—	No	—
rolling

[0098] The correlation between the yield strength in the T4 temper and the desensitization to intercrystalline corrosion in the T78 temper is again observed. It is also observed that a high cold-working rate does not induce a degradation of the yield strength after ageing, as could be expected, since the ageing kinetics is accelerated.

Claims

1. A method for production of aircraft structural components using rolled, extruded or forged products made of an aluminum alloy, comprising the steps of: casting a blank with a composition consisting essentially of, in % 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 aluminum; hot, and possibly cold, transforming of said blank to obtain a product; solution heat treating the product between 540 and 570° C.; quenching the solution heat treated product; forming and optionally welding the quenched product to produce the structural component; ageing the structural component, in one or more stages, for a total equivalent time at 175° C. expressed in hours between (−160+57γ) and (−184+69γ), γ being Si+2Mg+2Cu in % by weight.

2. Method according to claim 1, wherein the blank is homogenized at a temperature between 540 and 570° C.

3. Method according to claim 1, wherein the quenched product is subjected, before ageing, to cold-working resulting in a permanent elongation between 1 and 15%.

4. Method according to claim 3, wherein the cold working results in a permanent elongation between 2 and 10%.

5. Method according to claim 1, wherein total equivalent time at 175° C., in hours is between (−150+57γ) and (−184+69γ).

6. Method according to claim 1, wherein the product has a composition, in % by weight: Si: 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+2 Mg: 2.0-2.6; other elements <0.05 each and <0.15 in total; remainder alumnium.

7. Method according to claim 6, wherein Si+2Mg is between 2.3 and 2.6.

8. Method according to claim 6, wherein total equivalent ageing time at 175° C. is between 40 and 65 hours.

9. Method according to claim 3, additionally comprising assembly of sheets and profiles produced by the method, wherein the profiles undergo, before said assembly and said ageing, an additional cold-working step in relation to the sheets, such that resistance to intercrystalline corrosion of the profiles is at the same level as the sheets.

10. Aircraft fuselage component, produced using a rolled, extruded or forged product made of an aluminum alloy consisting essentially of, in % by weight: Si: 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+2 Mg: 2-2.6; other elements <0.05 each and <0.15 in total; remainder aluminum; subjected to a solution heat treatment, quenching, shaping and ageing in a T78 temper with a total equivalent time at 175° C. between 40 and 65 hours.

11. Fuselage component according to claim 10, wherein Si+2Mg is between 2.3 and 2.6.

12. Fuselage component according to claim 10, having, in TL direction, a yield strength R0.2>330 MPa, an ultimate tensile strength Rm>360 MPa and an elongation A>8%.

13. Fuselage component according to claim 10, having a plane strain fracture toughness in T-L direction, with at least one of the properties: KR (Δa=20 mm)>90 MPa{square root}m; KR (Δa=40 mm)>115 MPa{square root}m; Kc0>80 MPa{square root}m; Kc>110 MPa{square root}m.

14. Fuselage component according to claim 10, having a plane strain fracture toughness in L-T direction such that: Kc0>90 MPa{square root}m or Kc>130 MPa{square root}m.

15. Fuselage component according to claim 10, having a crack growth rate da/dn, measured in T-L direction for R=0.1, less than: 2 10−3 mm/cycle for ΔK=20 MPa{square root}m; 4 10−3 mm/cycle for ΔK=25 MPa{square root}m; 8 10−3 mm/cycle for ΔK=30 MPa{square root}m.

16. Fuselage component produced using a rolled, extruded or forged product made of alloy consisting essentially of, in % by weight: Si: 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+2 Mg: 2-2.6; other elements <0.05 each and <0.15 in total; the remainder aluminum; subjected to a solution heat treatment, quenching, shaping and ageing in T6 temper.

17. Fuselage component according to claim 16, having in TL direction, a yield strength R0.2>350 MPa, an ultimate tensile strength Rm>380 MPa and an elongation A>6%.

18. Fuselage component according to claim 16, having a plane strain fracture toughness in T-L direction, with at least one of the properties: KR (Δa=20 mm)>95 MPa{square root}m; KR (Δa=40 mm)>120 MPa{square root}m; Kc0>85 MPa{square root}m; Kc>115 MPa{square root}m.

19. Fuselage component according to claim 16, having a plane strain fracture toughness in L-T direction such that: Kc0>100 MPa{square root}m or Kc>150 MPa{square root}m.

20. Fuselage component according to claim 16, having a crack growth rate da/dn, measured in T-L direction for R=0.1, less than: 2 10−3 mm/cycle for ΔK=20 MPa{square root}m; 4 10−3 mm/cycle for ΔK=25 MPa{square root}m; 8 10−3 mm/cycle for ΔK=30 MPa{square root}m.