Aircraft structure element made of an Al-Cu-Mg- alloy

Details Aircraft structure element made of an Al-Cu-Mg- alloy Aircraft structure element made of an Al-Cu-Mg- alloy

2003-04-24

2004-02-17

Koehler, Robert R. (Department: 1775)

Metal treatment

Process of modifying or maintaining internal physical...

With casting or solidifying from melt

C148S550000, C148S552000, C148S690000, C148S691000, C148S694000, C148S695000, C148S700000

Reexamination Certificate

active

06692589

ABSTRACT:

FIELD OF THE INVENTION
This invention relates to aircraft structure elements, particularly skin panels and lower wing stringers for high capacity commercial aircraft, made from rolled, extruded or forged products made of an AlCuMg alloy in the treated temper by solution heat treating, quenching and aging, and introducing a compromise between the different required usage properties that is better than is possible with products according to prior art.
The designations of the alloys and metallurgical tempers used below are according to the designations used by the Aluminum Association, and reused in European Standards EN 515 and EN 573 part 3.
STATE OF THE ART
Wings of high capacity commercial aircraft comprise an upper part consisting of a skin made of thick plates made from a 7150 alloy in T651 temper, or a 7055 alloy in T7751 temper or a 7449 alloy in T7951 temper, and stringers made from sections of the same alloy, and a lower part composed of a prefabricated skin made of thick plates of 2024 alloy in the T351 temper or 2324 alloy in the T39 temper, and stringers made from sections of the same alloy. The two parts are assembled by spars and ribs.
The 2024 alloy according to the designations of the Aluminum Association or standard EN 573-3 has the following chemical composition (% by weight):
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<0.15.
Various alternative solutions have been proposed to improve the compromise between the various required properties, particularly mechanical strength and toughness. Boeing has developed the 2034 alloy with the following 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=0.08-0.15.
This alloy is described in patent EP 0031605 (equivalent to U.S. Pat. No. 4,336,075). Compared with the 2024 alloy in the T351 state, it has a higher specific yield strength due to the increased content of manganese and the addition of another anti-recrystallizing agent (Zr), and improved toughness and fatigue resistance.
U.S. Pat. No. 5,652,063 (Alcoa) concerns an aircraft structure element made starting from an alloy with the following composition (% by weight):
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 and Cu/Mg is between 5 and 9.
The yield strength of the sheet metal made from this alloy in the T8 temper is >77 ksi (531 MPa) This alloy is intended particularly for supersonic aircraft.
U.S. Pat. No. 5,593,516 (Reynolds) relates to an alloy for aeronautical applications containing 2.5 to 5.5% Cu and 0.1 to 2.3% Mg, in which Cu and Mg contents are kept below their solubility limit in aluminum and are related by the following equations:
Cu
max
=5.59-0.91 Mg and Cu
min
=4.59-0.91 Mg
The alloy may also contain Zr<0.20%, V<0.20%, Mn<0.80%, Ti<0.05%, Fe<0.15%, Si<0.10%.
U.S. Pat. Nos. 5,376,192 and 5,512,112 originating from the same initial patent application are applicable to alloys of this type containing 0.1 to 1% of silver. Note that the use of silver in this type of alloy increases the production cost and creates difficulties in recycling manufacturing scrap.
Furthermore, for many years, “AU6MGT” type alloys have been known, according to the old alloy designations in France. Patent FR 1379764 filed by Pechiney in 1963 applies to the use of an alloy of this type with 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 for the manufacture of compressed gas cylinders.
The Aluminum Association registered the 2001 alloy in 1976, with the following 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.
To the best knowledge of the inventors, there is no other industrial use of this alloy apart from compressed gas cylinders manufactured by reverse extrusion.
Problem Posed
The current trend in commercial aircraft construction is to use an increasing number of very thick products, with structure elements being machined in the body of these parts. For example, for some small aircraft, wing skins are machined from relatively thick plates to enable in-depth machining of wing stringers, although these stringers are usually made from sections or folded plates and are then mechanically fixed to the skin. Integral in-depth machining of the skin-stringer assembly can reduce manufacturing costs, since there are fewer parts and assembly is avoided. Furthermore, the use of an unassembled structure reduces the weight of the assembly.
Therefore it is desirable that, in addition to the properties normally required for aircraft structure elements, namely high mechanical strength, good tolerance to damage, good fatigue resistance and good resistance to the different forms of corrosion, plates need uniform mechanical properties throughout their thickness, in other words their properties should not vary significantly as a function of the thickness, typically between 10 and 120 mm. Furthermore, the more machining is necessary, the more desirable it becomes to maintain good stability under machining, and this is achieved by a low level of internal stresses. It is known that the mechanical properties for a thick plate are more uniform and internal stresses are lower if the plate is less sensitive to quenching.
Finally, aircraft wings, particularly for high capacity aircraft, have a curved wing profile with curvature in the longitudinal and in the transverse directions. This complex shape can be obtained in an autoclave during the aging process by forming on a mold, by applying a partial relative vacuum on the surface of the mold side of the plate, lower than the pressure on the other side. It is essential that this operation is successful to avoid expensive scrapping of parts with high added value, and particularly large parts. The key to success is in the lowest possible springback effect for a given mold shape, since springback is frequently the most difficult factor to be controlled.
The purpose of this invention is to supply aircraft structure elements with properties at least equivalent to the properties of the same elements made from a 2024 alloy in the T351 temper concerning static mechanical properties, toughness, crack propagation rate and resistance to corrosion, by using rolled, extruded or forged products with low residual stresses, low quench sensitivity and good formability during aging.
Purpose of the Invention
The purpose of the invention is a structure element, particularly a lower wing element, manufactured from a rolled, extruded or forged product made of an alloy with composition (% by weight):
Cu=4.6-5.3, Mg=0.10-0.50, Mn=0.15-0.45, Si<0.10, Fe<0.15, Zn<0.20, Cr<0.10, other elements <0.05 each and <0.15 total, the remainder being Al treated by solution heat treating, quenching, controlled tension to more than 1.5% permanent deformation and aging.
This element has at least one of the following properties:
yield strength R
0.2
(TL direction)>350 MPa, and preferably>370 MPa,
toughness K
lc
(L-T direction)>42 MPam
resistance of P type to intercrystalline corrosion according to standard ASTM G110.
Another purpose of the invention is a manufacturing process for a structure element comprising:
a) casting a plate or a billet with the composition mentioned above,
b) homogenization of this plate or billet,
c) hot transformation of this plate by rolling or of this billet by extrusion or forging to obtain a product thicker than 10 mm,
d) quenching of the hot transformed product,
e) solution heat treating of this product, preferably at a temperature of less than 10° C. at the incipient melting temperature of the alloy,
f) controlled tension of the product to obtain a permanent deformation of more than 1.5%,
g) aging of the product at a temperature greater than 160° C., possibly together with forming,
h) machining of the

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