CN100507066C - Structural element for aircraft engineering exhibiting a variation of performance characteristics - Google Patents

Abstract

The invention relates to a method for producing a part made of aluminium alloy with structural hardening consisting a) in melting a semifinished extruded or forged rolled stock followed by water quenching, b) in optionally carrying out a controlled drawing with a permanent stretching which is equal to or greater than 5 % and c) in tempering. Said invention is characterised in that at least one tempering operation is carried out in a controlled thermal gradient linear furnace comprising at least two areas or groups of areas (Z1, Z2) at initial temperatures (T1, T2) in which the variation of each temperature (T1, T2) around a specified temperature is equal or less than <5 DEG C through the length of said areas of the groups of areas, the difference between the specified temperatures of the initial temperatures (T1, T2) is equal or greater than 5 DEG C and said areas or the groups of areas are separable by a transition area or group of areas (Z1,2) inside of which the initial temperature varies from T1 to T2 and a length which is parallel to the axis of the linear furnace of each at least two areas or the groups of areas is equal to or greater than 1 meter.

Description

Structural members for aeronautical construction with service performance variations

technical field

The present invention relates to a strain hardened product and structural member made of a heat-treated aluminum alloy, especially for aeronautical construction. The invention relates in particular to so-called long products, in other words products whose length is significantly greater than their width or thickness, usually equal to at least twice their width and usually at least 5 meters long. These products may be rolled products (such as sheet, plate, heavy plate), extruded products (such as rods, profiles, tubes or wires) and forged products.

Background technique

Very large aircraft have very specific construction problems. For example, the assembly of structural members is becoming more and more critical, firstly because of cost factors (riveting is a very expensive process) and secondly because of the discontinuities created in the performance of the assembled parts.

To minimize assembly, structural members can be prepared by finishing in thick plates; different functions such as wing skin functions and wing stiffener functions can then be integrated into these one-piece (monolithic) structures in the component. At the same time, the size of the monolithic structural member can be increased. A new processing problem is introduced here for these parts made by rolling, extruding, forging or casting, since it is more difficult to guarantee uniform properties for very large parts.

The production of monolithic components with controlled property variation has been mentioned, which theoretically provides a better way of adapting the properties of the component to the needs of the manufacturer. Patent EP0630986 (Pechiney Rhenalu) describes a process for the manufacture of aluminum alloy plates with a structure hardened with continuously varying properties in use, in which the final annealing is carried out in a furnace with a special structure consisting of a heat chamber connected by a heat pump and cold room. The process is used to obtain small parts made of 7010 alloy with a length of about one meter, which is in the T651 temper at one end and in the T7451 temper at the other end, using isochronous annealing. The method has never been developed industrially due to its difficulty in controlling to meet the quality requirements necessary in the field of aeronautical construction; these difficulties increase with the size of the part, realizing the integration of two or more functions into a single structural member Medium is very meaningful for very large objects. Another problem arising from this process, such as mentioned in this patent, is that the optimum durations for T651 and T7451 processing are different. Another problem that arises is that the 7010 product in the T7451 temper is usually obtained by annealing with two plateaus, while the T561 temper is obtained by annealing with a single plateau.

The problem solved by the present invention is to develop a method for the manufacture of structural elements with variations in service performance, which are used for the manufacture of very long parts, especially for aeronautical construction, which are subject to strict quality assurance and aeronautical Generally required to be sufficiently controllable, stable and reproducible under the conditions of statistical process control.

purpose of invention

A first object of the invention is a process for the manufacture of aluminum alloy parts with structure hardening, comprising:

a) solution heat-treated rolled, extruded or forged semi-finished products, followed by quenching,

b) optionally, controlled stretching with at least 0.5% permanent elongation,

c) annealing treatment,

It is characterized in that at least one step of the annealing treatment is carried out in a linear furnace with a controlled temperature profile comprising at least two zones or groups of zones Z ₁ , Z ₂ with initial temperatures T ₁ and T ₂ , wherein the temperature variation around the set temperature for each temperature _T1 and _T2 does not exceed ±5°C (preferably ±4°C, and even more preferably ±3°C) over the length of said zone or group of zones, initially The difference between the set temperatures of the temperatures _T1 and _T2 is greater than or equal to 5°C (preferably between 10°C and 80°C, more preferably between 10°C and 50°C, and more preferably from 20°C to 40°C) . Said zone or set of zones may be separated by a zone or set of zones _Z1,2 called a transition zone or set of zones within which the initial temperature changes from _T1 to _T2 and said Each of the at least two zones or groups of zones _Z1 and _Z2 has a length parallel to the axis of the linear furnace of at least one meter (and preferably at least two meters).

A second object of the invention is a monolithic structural member made of an aluminum alloy having a hardened structure and a length L greater than its width B and thickness E, especially for aeronautical construction, said monolithic structural member being characterized in that, at the At least two segments _P1 and _P2 on different lengths of said structural member have physical properties (measured at intermediate thicknesses) selected from the following group:

a) P ₁ : K _IC(LT) >38MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa, even better than 600MPa)

b) P ₁ : K _IC(LT) >40MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa)

c) P ₁ : K _IC(LT) >41MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa)

d) P ₁ : K _IC(LT) >42MPa√m and P ₂ : R _m (L)>590Mpa

e)P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _m (TL)>550Mpa

f) P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _p0.2 (L)>550Mpa

i) P ₁ : K _IC(LT) >39MPa√m and P ₁ : R _m (L)>530Mpa and P ₂ : R _m (L)>580Mpa

j) P ₁ : K _IC(LT) >40MPa√m and P ₁ : R _m (L)>540Mpa and P ₂ : R _m (L)>590Mpa

k) P ₁ : K _{app(LT)(CCT406)} >125 MPa√m and P ₂ : R _m (L) >590 MPa.

Yet another object of the invention is an aircraft comprising at least one wing panel manufactured from a structural member according to the invention, wherein the segment P ₁ is close to the fuselage and the segment P ₂ is close to the geometric end of the wing, opposite to the fuselage.

Description of drawings

Figure 1 graphically shows the variation of static mechanical properties (curve 1), such as tensile and compressive strength, and dynamic properties (curve 2), such as failure tolerance, over the length of a wing according to the invention.

Figure 2 shows the mechanical strength of a 34 meter long structural member according to the invention.

invention disclosure

a) Terminology

All indications regarding the chemical composition of the alloys are expressed as weight percents unless otherwise mentioned. Therefore, in the mathematical expression, "0.4 Zn" is expressed as a mass percent of zinc content of 0.4; applies mutatis mutandis to other chemical elements. Alloy designations follow Aluminum Association rules well known to those skilled in the art. Metallurgical states are defined in European Standard EN515. The chemical composition of normalized aluminum alloys is defined, for example, in Standard 573-3. Unless mentioned otherwise, the static mechanical properties, in other words the ultimate strength R _m , the yield stress R _p02 and the elongation at failure A were determined by tensile tests according to standard EN10002-1, the position of these specimens obtained and their orientation by Standard EN485-1 limits. Toughness K _IC is determined according to standard ASTME399. The R curve is determined according to the standard ASTM561. The critical stress intensity factor K _C , in other words the crack non-plateau intensity factor is calculated from the R curve. The stress intensity factor K _C0 can also be calculated at the beginning of individual value loading by assigning the initial crack length to the critical load. Calculate these two values for a test specimen of the desired shape. K _app represents the K _co corresponding to the sample used for the R-curve test. Exfoliation resistance is determined according to the EXCO test described in standard ASTM G34.

The definitions given in European Standard EN 12258-1 may be used unless mentioned otherwise. The term "plate" is used in this patent for rolled products of all thicknesses.

The term "machining" includes any process that removes material, such as turning, milling, drilling, trimming, electroerosion, grinding and polishing, chemical milling.

The term "extruded product" also includes products that are drawn after extrusion, for example cold drawn through an extrusion die. It may also include deadlift products.

The term "structural member" refers to an element used in a mechanical construction, the static and/or dynamic mechanical properties of which are of particular importance to the integrity and performance of the structure, and for which typically specific or completed structural calculations are available. Failure of such mechanical components can often compromise the safety of the structure and its users, passengers or others. For aircraft, these structural members include, inter alia, elements that make up the fuselage (such as fuselage skin, fuselage stiffeners or stringers, bulkheads, frames around the fuselage, wings (such as wing skin), stringers or Reinforcements, ribs and spars and especially rear wings made of horizontal or vertical stabilizers, and floor beams, seat rails and doors).

The term "integral structural member" refers to a structural member made of a single piece of rolled, extruded, forged or cast semi-finished product, which does not require assembly, such as riveting, welding, bonding, with other parts.

Detailed Description of the Invention

The problem according to the invention is solved by a method in which the temperature in the furnace (its internal length being greater than the length of the workpiece to be heat treated) is kept approximately constant in at least two zones of the furnace which are at least one meter long, while at least one At least one other region of meter length is significantly different in temperature. This type of temperature distribution can be obtained by dividing the furnace into several temperature zones along its length.

The invention can be applied to all long metal products, in other words products whose length in one dimension (called length) is significantly longer than the other two dimensions (width, thickness). Length is the largest dimension of this product. Typically, within the context of the present invention, the length is at least twice the other two dimensions. In preferred embodiments, this length is five or even ten times the length of the other two dimensions. It usually applies to the longitudinal machine direction (rolling or extrusion direction); but in some cases it can be different. Products according to the invention may be rolled products such as plates or slabs, extruded products such as rods, tubes or profiles, and forged products; these products may be fabricated or machined.

For illustrative purposes, a "segment with extreme performance" of a product refers to the segment with the greatest difference in performance. Depending on the manufacturing method chosen, these areas may be close to the "geometric end" (or "geometric end") of the product, or they may be elsewhere; the invention can also be used to manufacture parts where two At least one of the regions with the greatest difference in properties is not near the geometric ends of the part but near the geometric center.

For purposes of illustration, a "zone" of a furnace is the smallest thermal unit defined by a segment along the length of the furnace and characterized by an approximately constant temperature, in other words, compared to the temperature difference that characterizes the variation of the furnace temperature distribution over the entire length of the furnace , the temperature variation parallel to the axis of the furnace is small. Furnace zones of this type are characterized by heating and control means which maintain the temperature in said zone at a greater than constant value. The temperature fluctuation around the set temperature in this region is preferably not more than ±5°C, more preferably not more than ±4°C. In a preferred embodiment, the difference is no more than about ±3°C. Certain products require temperature variations not to exceed approximately ±2°C. In the other direction of the furnace, the temperature should be as constant as possible. In general, the temperature variation around the set temperature within a zone must be smaller than the temperature variation between the coldest and hottest zones of the furnace.

Several adjacent zones may form a "zone group", in other words a thermal unit within which the temperature is approximately constant, resulting in a controlled temperature distribution. For example, in a linear furnace comprising nine furnace zones (numbered 1 to 9), two temperature zone groups may be formed, each temperature zone group comprising three furnace zones (numbered sequentially 1 , 2, 3, 7, 8 and 9), in said central zone group, three furnace zones (numbered sequentially 4, 5 and 6) are used to obtain a controlled temperature gradient. In this patent, the term "group of zones" may include only a single furnace zone.

According to the applicant's observations, the minimum temperature difference that produces an industrially applicable difference in properties between two segments with extreme properties of the product according to the invention is five degrees. A difference of at least ten degrees is preferred. The temperature difference can be larger, up to 80°C or 100°C, or more, but this can cause problems in the control of the temperature and temperature distribution parallel to the axis of the linear furnace, especially for relatively small parts. If the age-hardened state annealed state is to be obtained, the temperature difference usually does not exceed fifty degrees. A temperature difference of greater than fifty degrees can be advantageously used to manufacture components in which a segment with extreme properties is in a state close to T3 or T4. For Al-Zn-Cu-Mg type (series 7xxx) alloys, the relatively small temperature difference (from about ten degrees to about thirty degrees) allows it to be applied to structural members for aircraft construction, while Al-Cu type (Series 2xxx) alloys generally require larger temperature differences, such as values between about fifty to about one hundred degrees, or even higher.

The applicant has observed that not only the temperature difference between the two segments with extreme performance ends is relevant, but also the control of the temperature between the segments with extreme performance ends. This is why the present invention uses a furnace comprising a plurality of adjacent furnace zones. "Multiple" means at least two, and preferably at least three furnace zones. The spacing between two adjacent areas suggested in patent EP630986 is neither necessary nor useful. It does not allow adequate control of the temperature between the two zones. Similarly, the use of a heat pump connecting the cold chamber to the hot chamber as proposed in EP630986 makes the temperature distribution in the furnace very unstable. In the context of the present invention, a good control of the temperature distribution inside the furnace is necessary for the manufacture of structural components suitable for the quality assurance requirements of aeronautical products.

For this purpose, it must be possible to control and preferably regulate the temperature of each furnace zone. In an advantageous embodiment of the invention, the furnace comprises at least three furnace zones having a unit length of at least one meter. For example, to manufacture a structural element having a length of about thirty-four meters, the inventors used a linear furnace with a total length of thirty-six meters and thirty furnace zones of approximately equal length, adjustable independently of each other. Advantageously, the thirty furnace zones are grouped such that a smaller number of temperature zone groups is formed, for example three to five groups.

The process according to the invention consists in the production of strain hardened components made of aluminum alloys, utilizing structure hardening, solution heat treatment, quenching, possible stretching with a permanent elongation of at least 0.5% and controlled temperature in a furnace Distributed annealing treatment. Said annealing treatment in a furnace with a controlled temperature profile may comprise one or several temperature plateaus, and usually two or three plateaus, or more or less continuous temperature ramps with no clearly defined plateaus, for constituting the temperature at least one of the gradient's set of temperature zones. Optionally, an annealing treatment in a furnace with a temperature profile can precede or follow another annealing treatment in a uniform furnace (which can also be a linear furnace, adjusted to obtain a uniform temperature in all zones, or other furnaces) . This final annealing in a homogeneous furnace is especially useful when the aim is to obtain a condition that can be used for age forming. In this case, final annealing is used for age forming. In another embodiment, the part may be annealed in a furnace with a controlled temperature profile, followed by at least one forming or machining operation, followed by an annealing treatment step in a uniform furnace.

The invention can be used for the manufacture of structurally hardened monolithic structural elements made of aluminum alloys, having a length L greater than their width B and thickness E, especially suitable for aeronautical construction, said monolithic structural elements being characterized in that At least two segments _P1 and _P2 at different lengths of said structural member have physical properties (measured at intermediate thicknesses) selected from these groups:

a) P ₁ : K _IC(LT) >38MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa, even better than 600MPa)

b) P ₁ : K _IC(LT) >40MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa)

c) P ₁ : K _IC(LT) >41MPa√m and P ₂ : R _m (L) >580Mpa (and preferably >590Mpa)

d) P ₁ : K _IC(LT) >42MPa√m and P ₂ : R _m (L)>590Mpa

e)P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _m (TL)>550Mpa

f) P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _p0.2 (L)>550Mpa

i) P ₁ : K _IC(LT) >39MPa√m and P ₁ : R _m (L)>530Mpa and P ₂ : R _m (L)>580Mpa

j) P ₁ : K _IC(LT) >40MPa√m and P ₁ : R _m (L)>540Mpa and P ₂ : R _m (L)>590Mpa

k) P ₁ : K _{app(LT)(CCT406)} >125 MPa√m and P ₂ : R _m (L) >590 MPa.

It is preferred if the process is carried out such that the elongation at failure A(L) in segments P ₁ and P ₂ is greater than 9% and preferably >10%. This is particularly advantageous when the workpiece is to be subjected to a forming operation after aging. Similarly, it is more preferable that the A(L) of these fragments other than _P1 and _P2 is greater than 9%. Semi-finished products can be manufactured where (measured at intermediate thickness)

a) R _p0.2 , in L direction or in LT direction, with a difference R _p0.2(P2) - R _p0.2(P1) of at least 50 MPa, and preferably at least >75 MPa, and/or

b) R _p0.2 , in the TC direction, has a difference R _p0.2(p2) - R _p0.2(P1) of at least 30 MPa, and preferably at least 50 MPa, and/or

c) K _IC , measured in the LT direction, has a difference K _IC(P1) -K _IC(P2) of at least 5 MPa√m, and preferably at least 7 MPa√m, and/or

d) K _app , measured in the LT direction, has a difference K _{app(P1 )} - K _app(P2) of at least 10 MPa√m, and preferably at least 15 MPa√m.

The process according to the invention can be used to produce semi-finished products made of any alloy with structure hardening, such as aluminum alloys of the 2xxx, 4xxx, 6xxx and 7xxx series, and alloys with structure hardening of the 8xxx series containing lithium.

In the case of Al-Zn-Cu-Mg type alloys (7xxx series), the process according to the invention can be used to make one of the segments have extreme properties in a near T6 temper and the other in a near T74 or T73 temper extreme performance.

In alloys of the 2xxx or 6xxx series, as well as in lithium-containing 8xxx series alloys, the process according to the invention can be used to make one of the segments have extreme properties in a state close to T3 or T4, while the other has a state close to T6 Or extreme performance in T8 status.

In an advantageous embodiment of the invention, the alloy comprises between 7 and 15% zinc, between 1 and 3% copper and between 1.5 and 3.5% magnesium. In another advantageous embodiment, the zinc content is at least 7%, preferably between 8 and 13%, and especially preferably between 8.5 and 11%. The copper content is advantageously between 1.3 and 2.1%, and the magnesium content is between 1.8 and 2.7%. These alloys include 7449, and 7349 and 7056, both of which can produce very high mechanical strength (such as in the T651 or T7951 temper) and very high toughness (such as in the T76, T7651 or T74 temper, or in the T7451, T73 or T7351 temper) , while maintaining acceptable corrosion resistance with a balance of mechanical strength and toughness, and acceptable (that is, at least EA grade) anti-exfoliation performance in the two-state and intermediate regions corresponding to the two segments of the product with extreme properties ( EXCO test).

In an advantageous embodiment of the invention, annealing is carried out on the sheet, profile or forging which has been solution heat treated, quenched and drawn, in two steps:

The first homogenization step lasts 2 to 12 hours at a temperature between 115°C and 125°C, during the second step a fragment or end is treated between 115°C and 125°C, and simultaneously, at 150°C and 160°C The temperature between processing the other piece or the other end, both for 8 to 24 hours.

This annealing is particularly suitable for products made of 7xxx alloys, and especially for products made of 7349, 7449 or 7056 alloys.

In another advantageous embodiment of the invention, annealing is carried out at about 120° C. on one section or end P ₁ of a product made of a 2xxx alloy such as 2024 or 2023, while on the other section or other end P ₂ Annealed at about 190°C to reach peak mechanical strength (state T851). In a variation of this embodiment, the segment or end that is not peak aged (ie P ₁ ) is aged at about 100°C (or 80°C), which is the partially aged state.

In another advantageous embodiment, annealing to peak mechanical strength (temper T651) is carried out at about 120°C on a segment or one end of a product made of a 7xxx alloy (such as 7349, or 7449 or 7056), simultaneously at two plateaus at 120°C And 150-165 ° C on the other fragment or the other end has been annealed (state T7651, T7451 or T7351).

In another advantageous embodiment, however, annealing is performed at about 190°C to peak mechanical strength (temper T6) on products made of 6xxx alloys such as 6056, while the two plateaus are over-annealed at the other end (temper T7851 ).

Metal parts obtained by the process according to the invention can be used as structural members in aerospace construction. These structural elements may be bifunctional or multifunctional, in other words they may combine different functions in a single monolithic part which in the prior art could only be combined by assembling different parts. These structural members may also simplify and reduce the weight of aircraft construction and manufacture, especially for very high capacity cargo and passenger aircraft.

A particular advantage of the process according to the invention is that the optimum properties are obtained at each end or each segment with extreme properties over a well-controlled length of the product. Therefore, the aircraft designer knows the exact length over which the product has the recommended and guaranteed optimum performance. In a particularly recommended preferred embodiment, the process according to the invention is used to manufacture a structural member of performance that does not have a continuous variation along its entire length, but wherein has physical properties (or at least some physical properties) over a specific length of the product ) remains constant for at least two regions. In an advantageous embodiment of the invention, the area has a length of at least one meter, and preferably at least two meters. Such a product, and its use as a structural element in an aircraft wing, is schematically shown in FIG. 1 .

Another particular advantage of the process according to the invention is the precise control of the transition segment P1 between two segment groups _P1 and _P2 (here also two or more groups, depending on the number of temperature zones) _{, 2} , where P ₁ and P ₂ can be fragments with extreme performance. For any particular performance (or group of performances) to be optimized, the aircraft designer does not require the maximum performance in the transition region, eg ultimate strength R _m(L) and toughness K _IC(LT) in the longitudinal direction. But he needs some kind of compromise between these properties or groups of properties, because in this transition zone the structural elements actually play a structural role and must meet precise specifications.

In particular, structural members are:

- upper or lower wing (skin) panels;

- upper or lower wing stringers;

- wing stiffeners;

- Fuselage reinforcements;

- splices, especially for upper or lower wing (skin) pieces;

- Fuselage panels.

The method according to the invention can be used for heat treatment of long parts or structural components. Typically, their cross-section perpendicular to the length is approximately constant over their length, but this is not necessarily the case. Similarly, parts may or may not be straight; for example slightly curved forged structural members may be handled. This method can also be used to process castings, but long castings are very unusual and difficult to manufacture. In a preferred embodiment, the length of the part is at least 5 meters, preferably at least 7 meters, but preferably 15 meters or at least 25 meters in length, to take full advantage of the ability to create several functional segments distributed over the entire length of the part. possibility. Thus, the structural member is made having at least two regions _P1 and _P2 , wherein the lengths F _P1 and F _P2 of said at least two segments _P1 and _P2 (expressed as a percentage of the total length L) are such that F _P1 >25% and F _P2 >25% and preferably F _P1 >30% and F _P2 >30%. In another embodiment, _FP1 >35% and _FP2 >30% or _FP1 >40% and _FP2 >30%.

The structural member according to the invention can advantageously be used in aeronautical construction. For example, it can be used for high-capacity aircraft comprising at least one wing comprising at least one (skin) panel made of a structural member according to the invention, characterized in that segment P ₁ is located close to fuselage, while segment _P2 is near the geometric end of the wing (see Figure 1). In an advantageous embodiment, said wing (skin) panels are at least 15 meters long, and preferably at least 25 meters long. As discussed in the examples below, the inventors made wing (skin) panels that can be longer than 30 meters.

The components and structural members may be integral. The method according to the invention can also be used for the heat treatment of parts or structural elements which are not monolithic, but which can be made of at least two rolled, extruded or forged parts or semi-finished parts (preferably made of a structure-hardened aluminum alloy made) assembled, for example by welding, riveting or bonding. It is also possible that one or more components in such an assembly may be made from other base materials than aluminum alloys.

In this embodiment, it is possible, for example, to initially assemble at least one aluminum alloy plate with structure hardening and at least one aluminum alloy profile with structure hardening by means of riveting, welding or bonding, and then use the method according to the invention to assemble the The components are processed. In an advantageous embodiment of this variant of the method according to the invention, the plates and profiles are in the T351 temper and the assembly is produced by laser beam welding (LBW), friction stir welding (FSW) or electron beam welding (EBW) . The applicant has observed that it may be preferable to treat such welded assemblies with the method according to the invention after welding, rather than pre-welding the semi-finished products (plates or profiles) to be used in said assemblies, as this improves the quality of the welded joint. Mechanical strength and improve its corrosion resistance. This effect is significant when the welded joint extends over the entire length of the structural member (eg approximately parallel to the longitudinal direction of the product).

The invention will be better understood after reading the following examples, which are not limiting.

Example :

A 36-meter-long, 2.5-meter-wide and 30-mm-thick plate was made by hot-rolling the rolled plate.

The alloy composition is:

Zn 9.1%, Mg 1.89%, Cu 1.57%, Fe 0.06%, Si 0.03%, Ti 0.03%, Zr 0.11%,

The other elements are <0.01% respectively.

The rolled panels were homogenized at 475°C for 14 hours. The input temperature to the hot rolls was 428°C, while the output temperature of the hot rolled sheet was 401°C. The panels were solution heat treated, quenched and stretched at 471°C for 6 hours, quenched in water between about 15 and 16°C, and then subjected to controlled stretching to about 2.5% permanent elongation. The slabs were then cut into 34 meter long slabs. It was placed elongated into a linear furnace consisting of thirty 1200mm long zones. All annealing temperatures are adjusted within an interval of less than ±3°C around the set value.

The annealing process consisted of a first homogenization step ("first plateau") at 120°C for 6 hours, followed by a second step during which an 18-meter geometric tip (called Z ₁ , corresponding to 15 furnace zones) at 155°C for 15 hours (the "second plateau" is preceded by an approximately 1-hour conditioning phase), while another 10.8-meter geometric end (called Z ₂ , corresponding to 9 furnaces area) at 120°C for 16 hours. The transition zone between these two ends is 7.2 meters long (called Z _1,2 corresponding to 6 furnace zones).

After this second step, the conductivity is measured at different locations on the plate:

Fragment P ₁ : between 18.2 and 19.5 MS/m

Fragment P ₂ : Between 22.5 and 23.5 MS/m

Fragment P _1,2 : Between 18.2 and 23.6 MS/m

The plate was then subjected to a third annealing treatment, ie the temperature was raised to 148°C for 1 hour and 30 minutes followed by a uniform anneal at 150°C for 15 hours. This third step is intended to simulate aging forming or annealing after the structural member has been formed.

The board was cut and characterized. Table 1 summarizes the static mechanical properties obtained from the tensile tests. They are averages obtained at various locations along the width of the plate measured on the mid-thickness. No significant variation in performance was observed across the width of the board.

The values for R _P0.2 in the L and LT directions are obtained from compression; these values are placed in brackets in Table 1.

Table 1

The toughness results K _IC and K _app (the latter obtained on the CT127 type test piece as well as on the CCT406 type test piece) are given in Table 2

Table 2

Position on length of 34m plate [mm] KIC(LT)[MPa√m] K_IC(TL)[MPa√m] K_app(LT)(CT127)[MPa√m] K_app(LT)(CCT406)[MPa√m] 0(P₁) 43.8 36.1 106 132 13600(P₁) 45.8 38.1 108 - 16000(P₁) 46.7 37.3 99 - 18400(P_{1, 2}) 43.0 34.2 102 - 20800(P_{1, 2}) 39.4 32.9 88 - 25600(P₂) 36.1 34.9 89 34000(P₂) 34.9 29.1 94 110

The 34 meter long panels can be used for wing (skin) panels of high capacity cargo or passenger aircraft. For this application, the segment of the panel with the extreme performance X-edge (corresponding to high toughness K _IC , low static mechanical properties) is mounted on the fuselage side, while the segment of the panel with the extreme performance Z-edge (corresponding to high static mechanical strength and Low toughness K _IC ) at the geometric end of the wing.

In Table 3 are shown the temperature settings of the linear furnace used for annealing during the second aging step and the distribution of the temperatures measured on the plate and in the air in the furnace region. It includes the temperature profile during the annealing steps at 120° C. and 155° C. during the stabilization temperature phase; the values given in Table 3 were measured at the middle width.

table 3

furnace area Set temperature [°C] Plate temperature [°C] Air temperature [°C] 1 120 3 120 120.5 6 120 120.8 120.8 9 120 124.4 124.3 10 123 125.9 126.7 11 129 129.9 129.7 14 147 147.7 148.3 16 155 157.2 156.6 17 155 156.8 156.6 18 155 155.3 154.9 twenty two 155 155.1 154.8 30 155

Claims (49)

1. Integral structural members made of aluminum alloys with structure hardening and whose length L is greater than their width B and thickness E, The monolithic structural member is characterized in that at least two segments _P1 and _P2 located on different lengths of the structural member have physical and mechanical properties measured at intermediate thicknesses selected from the following group: a) P ₁ : K _IC(LT) >38MPa√m and P ₂ : R _m (L)>580Mpa b) P ₁ : K _IC(LT) >40MPa√m and P ₂ : R _m (L)>580Mpa c) P ₁ : K _IC(LT) >41MPa√m and P ₂ : R _m (L)>580Mpa d) P ₁ : K _IC(LT) >42MPa√m and P ₂ : R _m (L)>590Mpa e)P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _m (TL)>550Mpa f) P ₁ : K _IC(LT) >39MPa√m and P ₂ : R _m (L)>580Mpa and P ₂ : R _p0.2 (L)>550Mpa i) P ₁ : K _IC(LT) >39MPa√m and P ₁ : R _m (L)>530Mpa and P ₂ : R _m (L)>580Mpa j) P ₁ : K _IC(LT) >40MPa√m and P ₁ : R _m (L)>540Mpa and P ₂ : R _m (L)>590Mpa k) P ₁ : K _{app(LT)(CCT406)} >125 MPa√m and P ₂ : R _m (L) >590 MPa. 2. Structural member according to claim 1, wherein in group a), R _m (L) > 590 MPa. 3. Structural member according to claim 1, wherein in group a), R _m (L) > 600 MPa. 4. The structural member according to claim 1, wherein in group b), _Rm (L) > 590 MPa. 5. The structural member according to claim 1, wherein in group c), R _m (L) > 590 MPa. 6. Structural member according to claim 1, wherein A _(L) >9% in segments _P1 and _P2 . 7. Structural member according to claim 6, wherein A _(L) >10% in segments _P1 and _P2 . 8. Structural element according to claim 6, characterized in that A _(L) > 9% outside segments _P1 and _P2 . 9. Structural member according to claim 1, wherein the lengths _FP1 and _FP2 of said segments _P1 and _P2 expressed as a percentage of the length L are such that _FP1 > 25% and _FP2 > 25%. 10. Structural member according to claim 9, wherein F _P1 >30% and F _P2 >30%. 11. Structural member according to claim 9, wherein F _P1 >35% and F _P2 >30%. 12. Structural member according to claim 11, wherein F _P1 >40% and F _P2 >30%. 13. The structural member of claim 1, wherein the alloy comprises between 6 and 15% zinc, between 1 and 3% copper and between 1.5 and 3.5% magnesium. 14. A structural member according to claim 13, wherein the zinc content is higher than 7%. 15. A structural member according to claim 13, wherein the zinc content is between 8 and 13%. 16. A structural member according to claim 15, wherein the zinc content is between 8.5 and 11%. 17. A structural member according to claim 15, wherein the copper content is between 1.3 and 2.1%. 18. A structural member according to claim 17, wherein the magnesium content is between 1.8 and 2.7%. 19. A structural element as claimed in any one of claims 1 to 18, wherein the length of the element is at least 7 metres. 20. A structural member according to claim 19, wherein the length of the member is at least 15 metres. 21. A structural member according to claim 19, wherein the length of the member is at least 25 meters. 22. A structural member according to claim 1, wherein said member is used in aerospace construction. 23. Integral structural members made of aluminum alloys with structure hardening and whose length L is greater than their width B and thickness E, wherein said monolithic structural member is made of an Al-Cu type alloy, and wherein it comprises at least two segments _P1 and _P2 located on different lengths of said structural member, one of which is in the T8 temper and the other is not In full aging condition. 24. A structural member according to claim 23, wherein said member is used in aerospace construction. 25. Integral structural members made of aluminum alloys with structure hardening and whose length L is greater than their width B and thickness E, The monolithic structural member is characterized in that it comprises two segments _P1 and _P2 , wherein: a) R _p0.2 measured in the L direction or in the LT direction has a difference R _p0.2(P2) - R _p0.2(P1) of at least 50 MPa, and/or b) R _p0.2 measured in the TC direction has a difference R _p0.2(P2) - R _p0.2(P1) of at least 30 MPa, and/or c) K _IC , measured in the LT direction, has a difference K _IC(P1) -K _IC(P2) of at least 5 MPa√m, and/or d) K _app , measured in the LT direction, has a difference K _{app(P1 )} −K _app(P2) of at least 10 Mpa√m. 26. A structural member according to claim 25, wherein said member is used in aerospace construction. 27. A structural member according to claim 25, wherein R _p0.2 measured in the L direction or in the LT direction has a difference R _p0.2(P2) - R _p0.2(P1) of at least >75 MPa. 28. A structural member according to claim 25, wherein _Rp0.2 measured in the TC direction has a difference _Rp0.2(P2) -Rp0.2 _(P1) of at least 50 MPa. 29. A structural member according to claim 25, wherein _KIC , measured in the LT direction, has a difference KIC _{(P1) -KIC(P2)} of at least 7 MPa√m. 30. A structural member according to claim 25, wherein _Kapp , measured in the LT direction, has a difference Kapp _(P1) -Kapp _(P2) of at least 15 MPa√m. 31. A process for manufacturing a structural member as claimed in any one of claims 1-30, comprising: a) solution heat-treated rolled, extruded or forged semi-finished products, followed by quenching, b) optionally, controlled stretching with at least 0.5% permanent elongation, c) annealing treatment, characterized in that at least one step of said annealing treatment is carried out in a furnace with a controlled temperature profile comprising at least two zones or groups of zones Z ₁ , Z ₂ with initial temperatures T ₁ and T ₂ , where the temperature variation around the set temperature for each temperature _T1 and _T2 does not exceed ±5°C over the length of the zone or group of zones and the difference between the set temperatures for the initial temperatures _T1 and _T2 greater than or equal to 5°C, said zone or zone group may be separated by a zone or zone group Z _1,2 called a transition zone or zone group within which the initial temperature varies from _T1 to _T2 , Also, each of said at least two zones or groups of zones _Z1 and _Z2 has a length parallel to the axis of the linear furnace of at least one meter. 32. A process according to claim 31, wherein the temperature variation around the set temperature of each temperature _T1 and _T2 does not exceed ±4°C over the length of said at least two zones or groups of zones _Z1 and _Z2 . 33. A process according to claim 31 _, wherein the temperature variation around the set temperature of each temperature _T1 and _T2 does not _exceed ± 3°C. 34. A process as claimed in any one of claims 31 to 33, wherein the difference between set temperatures _T1 and _T2 is between 10°C and 80°C. 35. The process according to claim 34, wherein the difference between the set temperatures _T1 and _T2 is between 10°C and 50°C. 36. The process according to claim 34, wherein the difference between the set temperatures _T1 and _T2 is between 20°C and 40°C. 37. The process according to any one of claims 31 to 33, wherein the temperature in at least one of the zones or zone groups _Z1 or _Z2 varies as a function of time according to at least two temperature plateaus, or according to a temperature ramp without a plateau Variety. 38. Process according to any one of claims 31 to 33, wherein said annealing treatment in a furnace with controlled temperature profile is followed by at least one forming or machining operation and an annealing treatment step in a homogeneous furnace. 39. A process according to any one of claims 31 to 33, wherein said annealing treatment in a furnace with a controlled temperature profile is preceded by an annealing treatment step in a homogeneous furnace. 40. A process according to any one of claims 31 to 33, wherein the length of the part is at least 5 metres. 41. The process of claim 40, wherein the length of the part is at least 7 meters. 42. The process of claim 40, wherein the length of the part is at least 15 meters. 43. The process of claim 40, wherein the length of the part is greater than 25 meters. 44. A process according to any one of claims 31 to 33, wherein said aluminum alloy part having structure hardening is monolithic. 45. A process according to any one of claims 31 to 33, wherein said aluminum alloy part having structural hardening is assembled from at least two aluminum alloy parts having structural hardening. 46. The process of claim 45, wherein the assembly is done by riveting, bonding, welding, laser beam welding, friction welding or electron beam welding. 47. A process according to any one of claims 31 to 33, wherein said annealing treatment comprises a first homogenization step at between 115°C and 125°C for 2 to 12 hours, and a second step, in the second step , one end was treated at a temperature between 115°C and 125°C, while the other end was treated at a temperature between 150°C and 160°C, both for 8 to 24 hours. 48. Use of a structural member according to any one of claims 1 to 30 for the manufacture of aircraft wing panels, wing stringers, spars, fuselage stiffeners, fuselage panels or splices. 49. Aircraft comprising at least one wing made of a structural element according to any one of claims 1 to 30, characterized in that segment _P1 is close to the fuselage and segment _P2 is close to the geometric end of the wing.

Images