Structural member for aeronautical construction with a variation of usage properties

This invention relates to a process for manufacturing an aluminium alloy part with structural hardening as well as to structural members including monolithic structural members and to products prepared from such structural members. A suitable process of the present invention involves annealing in a linear furnace with a controlled temperature profile comprising at least two zones or groups of zones Z1, Z2. The length parallel to the axis of the linear furnace of each of the at least two zones or groups of zones Z1 and Z2 is generally at least about one meter.

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Description
CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Application No. 60/555,304, filed Mar. 23, 2004, the content of which is incorporated herein by reference in its entirety.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to strain hardened products and structural members, particularly for aeronautical construction, made of a heat treatment aluminium alloy. In particular, the present invention relates to so-called long products, in other words products having a length that is significantly more than their width or thickness, typically with a length equal to at least twice their width, and typically at least 5 meters long. These products may be, for example, rolled products (such as thin plates, medium plates, thick plates), extruded products (such as bars, sections, tubes or wires), and forged products.

2. Description of Related Art

Very large aircraft have very particular construction problems. For example, the assembly of structural members becomes more and more critical, firstly because of the cost factor (riveting is a very expensive process), and secondly because they generate discontinuities in the properties of assembled parts.

To minimise assemblies, structural members can be prepared by integral machining in thick plates; different functions such as wing skin and wing stiffener can then be integrated into these single-piece (monolithic) structural members. At the same time, the dimensions of monolithic structural members can be increased. This introduces new manufacturing problems for these parts made by rolling, extrusion, forging or casting, since it is more difficult to guarantee uniform properties in very large parts.

The preparation of monolithic parts with a controlled variation of properties has also been mentioned, which in theory provides a means of better adapting properties of parts to the manufacturer's needs. EP 0 630 986 (Pechiney Rhenalu) describes a process for manufacturing aluminium alloy plates with structural hardening with a continuous variation in usage properties, in which final annealing is done in a furnace with a special structure comprising a hot chamber and a cold chamber, connected by a heat pump. This process has been used to obtain small parts with a length of about one meter made of a 7010 alloy, one end of which is in the T651 state, while the other end is in the T7451 state, wherein the process uses an isochronous annealing treatment. This process has never been developed industrially, since it is difficult to control compatibly with quality requirements necessary in the aeronautical construction field. These difficulties tend to increase even further as the size of the parts increases, knowing that the integration of two or more functions into one single structural member is especially interesting for very large pieces. Moreover, there is no real need for small mechanical parts with a continuous variation of usage properties. Another problem that arises with this process, for example as described in EP 0 630 986, is that the optimum durations of the T651 and T7451 treatments are different. Another problem that arises is that a 7010 product in the T7451 state is typically obtained by an annealing treatment with two plateaus, whereas the T651 state is obtained by an annealing treatment with a single plateau.

SUMMARY OF THE INVENTION

A problem addressed by the present invention was to develop a process for manufacturing structural members, particularly for aeronautical construction, with a variation of usage properties for the manufacture of very long parts, that is sufficiently controllable, stable and reproducible under strict quality assurance and statistical process control conditions that are typically required by aeronautics.

An object of the present invention was the provision of a process for manufacturing an aluminium alloy part with structural hardening, comprising:

    • solution heat treating a semi-finished rolled, extruded and/or forged product, followed by quenching,
    • optionally conducting controlled tension with permanent elongation of at least 0.5%, and
    • annealing,
    • wherein at least a portion of the annealing is done in a furnace with a controlled temperature profile comprising at least two zones or groups of zones Z1, Z2 with initial temperatures T1 and T2 in which the temperature variation around the set temperature for each of the temperatures T1 and T2 does not exceed about ±5° C. (preferably ±4° C. and even better ±3° C.) within the length of the zones or groups of zones, and further wherein the difference between the set values of the initial temperatures T1 and T2 are greater than or equal to about 5° C. (preferably from about 10° C. to about 80° C. and even better from about 10° C. to about 50° C., and still better from about 20° C. to about 40° C.). The zones or groups of zones can optionally be separated by a zone or a group of zones Z1,2 called a “transition group” within which the initial temperature varies from T1 to T2 and wherein the length parallel to the axis of the furnace of each of at least two zones or groups of zones Z1 and Z2 is at least about one meter (and preferably at least about two meters).

In further accordance with the present invention, there is provided a monolithic structural member comprising an aluminium alloy with structural hardening having a length L greater than its width B and/or thickness E. The structural member is particularly adapted for aeronautical construction, and advantageously includes at least two segments P1 and P2 located on a different length of the structural member that have mechanical properties (measured at mid-thickness) selected from the group consisting of:

    • a) P1: KIC(L-T)≧38 MPa{square root}m and P2: Rm(L)≧580 MPa (and preferably ≧590 MPa and even better ≧600 MPa
    • b) P1: KIC(L-T)≧40 MPa{square root}m and P2: Rm(L)≧580 MPa (and preferably ≧590 MPa)
    • c) P1: KIC(L-T)≧41 MPa{square root}m and P2: Rm(L)≧580 MPa (and preferably ≧590 MPa)
    • d) P1: KIC(L-T)≧42 MPa{square root}m and P2: Rm(L)≧590 MPa
    • e) P1: KIC(L-T)≧39 MPa{square root}m and P2: Rm(L)≧580 MPa and P2: Rm(TL)≧550 MPa
    • f) P1: KIC(L-T)≧39 MPa{square root}m and P2: Rm(L)≧580 MPa and P2: Rp0.2(L)≧550 MPa
    • i) P1: KIC(L-T)≧39 MPa{square root}m and P1: Rm(L)≧530 MPa, and P2: Rm(L)≧580 MPa
    • j) P1: KIC(L-T)≧40 MPa{square root}m and P1: Rm(L)≧540 MPa, and P2: Rm(L)≧590 MPa
    • k) P1: Kapp(L-T)(CCT406)≧125 MPa{square root}m and P2: Rm(L)≧590 MPa.

In yet further accordance with the present invention, there is provided an aircraft comprising at least one wing manufactured from a structural member according to this invention wherein segment P1 is located close to the fuselage and segment P2 is located close to a geometric tip of the wing.

Additional objects, features and advantages of the invention will be set forth in the description which follows, and in part, will be obvious from the description, or may be learned by practice of the invention. The objects, features and advantages of the invention may be realized and obtained by means of the instrumentalities and combination particularly pointed out in the appended claims.

BRIEF DESCRIPTION OF THE FIGURES

The accompanying drawings, which are incorporated in and constitute a part of the specification, illustrate a presently preferred embodiment of the invention, and, together with the general description given above and the detailed description of a preferred embodiment given below, serve to explain principles of the invention.

FIG. 1 diagrammatically shows the variation of static mechanical properties (curve 1) for example tensile or compression strength, and dynamic properties (curve 2), for example tolerance to damage, within the length of a wing panel according to the invention.

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

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

a) Terminology

Unless mentioned otherwise, all indications about the chemical composition of alloys are expressed in mass percentage by weight based on the weight of the alloy. Consequently, in a mathematical expression, “0.4 Zn” means 0.4 times the content of zinc, expressed as a mass percentage; this is applicable with any necessary changes to other chemical elements. The designation of alloys follows The Aluminum Association rules, known to those skilled in the art. Metallurgical states are defined in European standard EN 515. The chemical composition of normalized aluminium alloys is defined for example in standard EN 573-3. Unless mentioned otherwise, static mechanical characteristics, in other words the ultimate strength Rm, the yield stress Rp02 and the elongation at failure A, are determined by a tensile test according to standard EN 10002-1, the location at which these pieces are taken and their direction being defined in standard EN 485-1. The toughness KIC was measured according to standard ASTM E 399. The R curve is determined according to standard ASTM 561. The critical stress intensity factor KC, in other words the intensity factor that makes the crack unstable, is calculated starting from the R curve. The stress intensity factor KCO is also calculated by assigning the initial crack length to the critical load, at the beginning of the monotonous load. These two values are calculated for a test piece of the required shape. Kapp denotes the KCO corresponding to the test piece that was used to make the R curve test. Resistance to exfoliation corrosion was determined according to the EXCO test described in standard ASTM G34.

Definitions given in European standard EN 12258-1 are applicable unless mentioned otherwise. The term “plate” is used in this patent for all thicknesses of rolled products.

The term “machining” includes any process for removal of material such as turning, milling, drilling, trimming, electroerosion, grinding, polishing, and chemical milling.

The term “extruded product” also includes products that have been drawn after extrusion, for example by cold drawing through an extrusion die. It also includes hard drawn products.

The term “structural member” refers to an element used in mechanical construction for which the static and/or dynamic mechanical properties are particularly important for the performance and integrity of the structure, and for which a structure calculation is generally specified or done. It is typically a mechanical part that could endanger the safety of the said construction and its users, passengers or others, if it fails. For an aircraft, these structural members include particularly elements making up the fuselage (such as the fuselage skin, fuselage stiffeners or stringers, bulkheads, fuselage circumferential frames, wings (such as wing skins), stringers or stiffeners, ribs and spars and the tail fin composed particularly of horizontal and vertical stabilisers, and floor beams, seat tracks and doors.

The term “monolithic structural member” refers to a structural member made from a single piece of rolled, extruded, forged or cast semi—finished product, without assembly, such as riveting, welding, bonding with another part.

A problem existing in the prior art can addressed by employing a method wherein the temperature in a furnace, the internal length of which is greater than the length of the piece to be heat-treated, is kept approximately constant in at least two zones of a furnace of at least one meter long, while it is significantly different in at least one other zone of at least one meter long. This type of temperature profile can be obtained by subdividing the furnace lengthwise into several temperature zones.

The present inventive process is applicable to all long metallic products, in other words, products with one dimension (called the length) being significantly longer than the other two dimensions (width, thickness). The length is the largest dimension of the product. Typically, within the context of this invention, the length is at least twice as large as the other two dimensions. In preferred embodiments, the length is five or even ten times as large as the other two dimensions. Length normally applies to the longitudinal manufacturing direction (rolling or extrusion direction); but it may be different in some cases. Products according to the present invention may be rolled products (such as plates or thick plates), extruded products (such as bars, tubes or sections), and forged products; these products may be as manufactured or as machined.

For the purposes of this description, the “segments with extreme properties” of a product mean the segments with the greatest difference in properties. Depending on the chosen manufacturing methods, these segments may be located close to the “ends in the geometric sense” (or “geometric ends”) of the product, or they may be elsewhere. The present invention can also be used to make parts in which at least one of the two segments with the greatest difference in properties is closer to the geometric center than to the geometric end of the part.

For the purposes of this description, a “zone” of a furnace is the smallest thermal unit along the length of the furnace and characterized by an approximately constant temperature, in other words by a temperature variation parallel to the axis of the furnace that is small (typically less than one third) compared with the temperature difference that characterizes the variation of the furnace temperature over its total length. This type of furnace zone includes heating and control means that keep the temperature at an approximately constant value within the zone. That is, the temperature variation around the set temperature inside such a zone preferably does not exceed about ±5° C., and more preferably does not exceed about ±4° C. In a preferred embodiment, this difference does not exceed about ±3° C. Certain products may require a temperature variation not exceeding about ±2° C. In the other directions of the furnace, the temperature should be as constant as possible. In any case, the temperature variation around the set temperature within one zone should preferably be smaller than the variation of temperature between the coolest and the hottest zone of the furnace.

Several contiguous zones may form a “group of zones”, in other words a thermal unit within which the temperature is approximately constant (as defined above), forming a controlled temperature gradient. For example, a group of zones in a linear furnace could contain 9 furnace zones (numbered from 1 to 9), wherein two groups of temperature zones are formed, each comprising three furnace zones (numbered 1, 2, 3, 7, 8 and 9 successively) separated by a central group of zones in which there is a controlled temperature gradient obtained using three furnace zones (numbered 4, 5 and 6 successively). For the purposes of the present description, the term “zone group” may include only a single furnace zone.

According to observations made by the applicant, the minimum temperature difference that results in differences in properties that can be used industrially between two segments with extreme properties of a product according to the present invention, is preferably not less than about five degrees. A difference of at least ten degrees is preferred in some cases. The temperature difference may be much greater, e.g. up to 80° C. or 100° C., or even more, but this can cause problems in control of the temperature and the temperature profile parallel to the axis of the furnace, particularly in the case of relatively small parts. If age-hardened tempers are to be obtained, the temperature difference should typically not exceed fifty degrees. A temperature difference of more than fifty degrees can advantageously be used to make a part for which one of the segments with extreme properties is in a temper close to T3 or T4. For alloys of the Al—Zn—Cu—Mg type (series 7xxx), a rather small temperature difference (from about ten to about thirty degrees) leads to an effect which can be exploited, if desired, in structural members for aircraft construction, while alloys of the Al—Cu type (series 2xxx) usually involve larger temperature differences, such as a value from about 50 to about 100 degrees, or even higher.

The applicant has observed that it is not only the temperature difference between the two segments with extreme properties that matters, but also the temperature control of the temperature between the segments with extreme properties. This is why the present invention preferably uses a furnace comprising a plurality of contiguous furnace zones. “A plurality” means at least two, and preferably at least three furnace zones. A partition between two contiguous zones, as recommended in EP 630 986, is not necessary or required. That is, the use of a partition may not enable sufficient control over the temperature between two zones. Similarly, the use of a heat pump connecting the cold chamber to the hot chamber, as suggested in EP 630 986, may make the temperature profile inside the furnace too unstable. Within the context of the present invention, good control of the temperature profile within the furnace is desirable in order to be able to manufacture structural members compatibly with quality assurance requirements for aeronautical products.

For this purpose, it is highly advantageous to be able to control, and preferably regulate, the temperature in each furnace zone. In one advantageous embodiment of this invention, the furnace comprises at least three furnace zones with a unit length of at least about one meter. For example, to manufacture structural members with a length of about thirty-four meters, the inventors preferably use a furnace with a total length of thirty-six meters with thirty furnace zones with approximately equal lengths, preferably adjustable independently of each other. Advantageously, these thirty furnace zones are grouped so as to form a small number of groups of temperature zones, for example three to five groups.

A process according to the invention advantageously includes the production of a strain hardened part made of an aluminium alloy with structural hardening, solution heat treatment, quenching, possibly tension with a permanent elongation of at least 0.5%, and an annealing treatment in a furnace with a controlled temperature profile. The annealing treatment in a furnace with a controlled temperature profile may comprise one or several temperature plateaus, and typically two or three, or a more or less continuous temperature ramp with no clearly defined plateau, for at least one of the groups of temperature zones making up the temperature profile. Optionally, the annealing treatment in a furnace with a temperature profile is preceded or is followed by another annealing treatment step in a homogeneous furnace (that may be the same furnace, adjusted so as to obtain a uniform temperature in all zones, or another furnace). Such final annealing in a homogeneous furnace is particularly useful when the objective is to obtain a temper which can be used for age forming. In this case, the final anneal is used for age forming. In another embodiment, a part may be annealed in a furnace with a controlled temperature profile, following by at least one forming or machining operation, and then an annealing treatment step in a homogeneous furnace.

The invention can be used to make a monolithic structural member made of an aluminium alloy with structural hardening with a length L greater than its width B and thickness E, particularly for aeronautical construction, the monolithic structural member preferably wherein at least two segments P1 and P2 on different lengths of the structural member have physical properties (measured at mid-thickness) selected from the group formed of:

    • a) P1: KIC(L-T)>38 MPa{square root}m and P2: Rm(L)>580 MPa (and preferably >590 MPa and even better >600 MPa
    • b) P1: KIC(L-T)>40 MPa{square root}m and P2: Rm(L)>580 MPa (and preferably >590 MPa)
    • c) P1: KIC(L-T)>41 MPa{square root}m and P2: Rm(L)>580 MPa (and preferably >590 MPa)
    • d) P1: KIC(L-T)>42 MPa{square root}m and P2: Rm(L)>590 MPa
    • e) P1: KIC(L-T)>39 MPa{square root}m and P2: Rm(L)>580 MPa and P2: Rm(TL)>550 MPa
    • f) P1: KIC(L-T)>39 MPa{square root}m and P2: Rm(L)>580 MPa and P2: Rp0.2(L)>550 MPa
    • i) P1: KIC(L-T)>39 MPa{square root}m and P1: Rm(L)>530 MPa, and P2: Rm(L)>580 MPa
    • j) P1: KIC(L-T)>40 MPa{square root}m and P1: Rm(L)>540 MPa, and P2: Rm(L)>590 MPa
    • k) P1: Kapp(L-T)(CCT406)>125 MPa{square root}m et P2: Rm(L)>590 MPa.

It is preferable if the process is carried out such that the elongation at failure A(L) is greater than 9% and preferably >10% in segments P1 and P2. This is advantageous particularly when the parts are to be subjected to forming operations after aging. Similarly, it is preferable that A(L) is more than 9% outside these segments P1 and P2. It is possible to manufacture semi-products in which (measured at mid-thickness)

    • a) Rp0.2, determined in the L direction or in the LT direction, has a difference p0.2(P2)-Rp0.2(P1) of at least 50 MPa and preferably of at least >75 MPa, and/or
    • b) Rp0.2, determined in the ST direction, has a difference Rp0.2(P2)-Rp0.2(P1) of at least 30 MPa and preferably at least 50 MPa, and/or
    • c) KIC, measured in the L-T direction, has a difference KIC(P1)-KIC(P2) of at least 5 MPa{square root}m and preferably of at least 7 MPa{square root}m, and/or
    • d) Kapp, measured in the L-T direction, has a difference Kapp(P1)-Kapp(P2) of at least 10 MPa{square root}m and preferably of at least 15 MPa{square root}m.

A process according to the invention may be used to produce semi-finished products made of any alloy with structural hardening, such as aluminium alloys in the 2xxx, 4xxx, 6xxx and 7xxx series, and alloys with structural hardening such as those in the 8xxx series containing lithium.

A process according to the invention may be used, in the case of Al—Zn—Cu—Mg-type alloys (series 7xxx), for example, to put one of the segments with extreme properties in a temper close to T6, and another segment with extreme properties in a temper close to T74 or T73.

In alloys of the 2xxx or 6xxx series, as well as in lithium-containing alloys of the 8xxx series, a process according to the invention may be used, for example, to put one of the segments with extreme properties in a temper close to T3 or T4, and the other segments with extreme properties in a temper close to T6 or T8.

In one advantageous embodiment of the invention, the alloy comprises from about 7 to about 15% of zinc, from about 1 to about 3% of copper and from about 1.5 to about 3.5% of magnesium. In other advantageous embodiments, the zinc content is at least about 7%, and preferably from about 8 to about 13%, and more preferably from about 8.5 to about 11%. The copper content is advantageously from about 1.3 to about 2.1%, and the magnesium content is preferably from about 1.8 to about 2.7%. These alloys, including 7449, 7349 and 7056, can result in a very high mechanical strength (for example in the T651 or T7951 state) and very high toughness (for example in the T76, T7651 or T74 state, or in the T7451, T73 or T7351 state) while keeping acceptable corrosion resistance and compromise between mechanical strength and toughness, as well as an acceptable (i.e. at least EA rating) resistance to exfoliation corrosion (EXCO test) in the two states corresponding to two segments with extreme properties of the product and in intermediate zones.

In one advantageous embodiment of this invention, annealing is carried out on a plate, section or a forged part subjected to solution heat treatment, quenched and stretched, for example, in at least two steps:

A first homogenous step at a temperature between 115° C. and 125° C. for a duration of between 2 and 12 hours, and a second step during which one segment or end is treated at a temperature between 115° C. and 125° C., while the another segment or the other end is treated at a temperature between 150° C. and 160° C., both for a duration of between 8 and 24 hours.

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

In another advantageous embodiment of this invention, annealing is done at about 120° C. (i.e. under-aging) on one segment or end P1 of a product made of 2xxx alloy (such as 2024 or 2023), while annealing to the peak mechanical strength (temper T851) at about 190° C. is carried out on another segment or the other end P2. In a variant of this embodiment, the segment or end which is not peak-aged (i.e. P1) is aged at about 100° C. (or 80° C.).

In another advantageous embodiment, annealing to the peak mechanical strength (temper T651) is carried out on a segment or end of product made of a 7xxx alloy (such as 7349, 7449 or 7056) at about 120° C., while over-annealing (temper T7651, T7451 or T7351) is carried out at another segment or the other end in two plateaux at 120° C. and 150-165° C.

In yet another advantageous embodiment, annealing to the peak mechanical strength (state T6) is carried out on a product made of a 6xxx alloy (such as 6056) at about 190° C., while over annealing (state T7851) is carried out in two plateaux at the other end.

Metallic parts obtained by the process according to the invention can be used as structural members in aeronautical construction. These structural members may be bi-functional or multi-functional, in other words they may combine different functions in a single monolithic part that processes in prior art could only combine by assembly of different parts. These structural members of the present invention can also enable simpler and lighter weight construction and manufacturing of aircraft, particularly very high capacity freight or passenger aircraft.

One specific advantage of the process according to the invention is that optimum properties are achieved at each segment with extreme properties or at each end, over a well-controlled length of the product. Therefore the aircraft designer knows exactly the length over which the product will have the recommended and guaranteed optimum properties. In one particularly preferred embodiment, a process according to the invention is used to make structural members that do not have a continuous variation of properties along their entire length, but in which there are at least two zones in which the physical properties (or at least some of the physical properties) are constant over a certain length of the product. In one advantageous embodiment of the invention, the length of this zone is at least one meter, and preferably at least two meters. Such a product, as well as its use as a structural element in an aircraft wing, is schematically represented on FIG. 1.

Another specific advantage of the process according to the invention is precise control of properties in the transition segment P1,2 between two groups of segments P1 and P2 (there may be two or more groups, depending on the number of groups of temperature zones), wherein P1 and P2 may be segments with extreme properties. The aircraft designer does not need maximum properties in the transition zone for any particular property (or groups of properties) to be optimised, for example the ultimate strength in the longitudinal direction Rm(L) and the toughness KIC(L-T). But he does need a certain compromise between these properties or groups of properties, since in this transition zone the structural member actually plays a structural role and must satisfy precise specifications.

In particular, structural members include:

    • upper or lower wing (skin) panels;
    • upper or lower wing stringers;
    • wing spars;
    • fuselage stiffeners;
    • butt straps, particularly butt straps for upper and lower wing (skin) panels;
    • fuselage panels.

The process according to the invention can be used for heat treatment of long parts or structural members. Usually, their 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 could be treated. The process could also be used to treat cast parts, but long cast parts are very unusual and difficult to make. In one preferred embodiment, the length of the part is at least 5 meters, preferably at least 7 meters, but a length of 15 meters or at least 25 meters is preferable, to take full advantage of the possibilities of creating several functionalised segments distributed over the length of the part. Thus, structural members have been made with at least two zones P1 and P2 in which the length FP1 and FP2 (expressed in percent of the total length L) of the said at least two segments P1 and P2 is such that FP1>25% and FP2>25% and preferably FP1>30% and FP2>30%. In other embodiments, FP1>35% and FP2>30% or FP1>40% and FP2>30%.

Structural members according to the invention may advantageously be used in aeronautical construction. For example, a high capacity aircraft including at least one wing including at least one structural member according to the invention could be used, characterised in that segment P1 is located close to the fuselage, and segment P2 is close to the geometric tip of the wing (see FIG. 1). In one advantageous embodiment, the said wing (skin) panels are at least 15 meters long, and preferably at least 25 meters long. As described in the example below, the inventors have made wing (skin) panels more than 30 meters long.

The parts and structural members of the present invention may be monolithic. The process according to the invention can also be used for heat treatment of parts or structural members that are not monolithic, but are assembled from at least two rolled, extruded or forged parts or semi-finished parts (preferably made from an aluminium alloy with structural hardening), for example by welding, riveting or bonding. It is also possible that one or several parts in such an assembly could be made from a base material other than an aluminium alloy.

In this embodiment, it would, for example be possible to start by making an assembly between at least one aluminium alloy plate with structural hardening and at least one aluminium alloy section with structural hardening by riveting, welding or bonding, the said assembly then being treated by the process according to the invention. In one advantageous embodiment of this variant of the process according to the invention, the plates and sections are in the T351 state, and the assembly is made 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 a welded assembly after welding by the process according to the invention, instead of treating the semi-finished products (plates and sections) that will be used in the said assembly before welding, since this can improve the mechanical strength of the welded joint and its resistance to corrosion. This effect is significant when the welded joint is spread over a long length of the structural member (for example approximately parallel to the longitudinal direction of the product).

The invention will be better understood after reading the following example that is in no way limiting.

EXAMPLE

A 36-meter long, 2.5-meter wide and 30 mm thick plate is made by hot rolling of a rolling plate.

The alloy composition was:

    • Zn 9.1%, Mg 1.89%, Cu 1.57%, Fe 0.06%, Si 0.03%, Ti 0.03%, Zr 0.11%, other elements <0.01 each.

The rolling plate was homogenised for 14 hours at 475° C. The input temperature to the hot roller was 428° C., and the output temperature of the hot rolled plate was 401° C. The plate was solution heat treated, quenched and tensioned under the following conditions: holding for 6 hours at 471° C., quenching in water at a temperature between about 15 and 16° C., then controlled tension with a permanent elongation of about 2.5%. The plate was then cropped to give a 34-meter long plate. It was placed lengthwise in a furnace composed of thirty 1200 mm long zones. All annealing temperatures were adjusted within an interval of less than ±3° C. around the set value.

The annealing treatment consisted of a first homogenous treatment step for 6 hours at 120° C. (“first plateau”) and was immediately followed by a second step during which one 18-meter geometric tip (called Z1, corresponding to 15 furnace zones) was treated for 15 hours at 155° C. (“second plateau” preceded by an adjustment period of about 1 hour), while the other 10.8-meter geometric tip (called Z2, corresponding to 9 furnace zones) was held for 16 hours at 120° C. The transition zone between these two tips was 7.2 meters long (called Z1,2 corresponding to 6 furnace zones).

After this second step, the electrical conductivity of the plate was measured at different locations:

    • Segment P1: between 18.2 and 19.5 MS/m
    • Segment P2: between 22.5 and 23.5 MS/m
    • Segment P1,2: between 18.2 and 23.6 MS/m.

The plate was then subjected to a third annealing step, namely homogeneous annealing consisting of a temperature increase to 148° C. for 1h30, followed by holding at 150° C. for 15 hours. This third step was intended to simulate age forming or annealing after the structural member was shaped.

The plate was cut and characterised. Table 1 summarises the static mechanical properties obtained by a tension test. These are averages obtained from measurements made at mid-thickness and at different locations distributed along the plate width. No significant variation of properties was observed in the plate width. For RP0.2 in the L and LT direction, values have also been obtained by compression; these values are put between brackets in table 1.

TABLE 1 LT TC Position [mm] L (long) (long transverse) (short transverse) in the length direction direction direction of a 34 m Rm Rp0.2 A Rm Rp0.2 A Rm Rp0.2 A panel [MPa] [MPa] [%] [MPa] [MPa] [%] [MPa] [MPa] [%]   0 (P1) 561 517 13.5 550 506 12.5 550 495 8.5 (509) (519) 13600 (P1) 565 522 13.5 553 511 12.5 548 502 8.5 (513) (528) 16000 (P1) 556 509 13.5 547 501 12.5 540 500 8.5 (500) (514) 18400 (P1,2) 566 523 13.5 559 519 12.5 546 498 7.5 (527) (538) 20800 (P1,2) 612 587 12.0 598 575 11.5 590 545 7.0 (573) (593) 25600 (P2) 621 598 12.5 607 585 11.5 595 554 6.5 (590) (605) 34000 (P2) 624 602 12.1 608 586 11.5 599 558 6.1 (594) (607)

The toughness results KIC and Kapp (the latter obtained on a CT127 type test piece as well as on a CCT406 type test piece) are given in table 2

TABLE 2 Position [mm] along Kapp(L − T) Kapp(L − T) the length of a KIC (L − T) KIC (T − L) (CT127) (CCT406) 34 m panel [MPa{square root}m] [MPa{square root}m] [MPa{square root}m] [MPa{square root}m]   0 (P1) 43.8 36.1 106 132 13600 (P1) 45.8 38.1 108 — 16000 (P1) 46.7 37.3 99 — 18400 (P1,2) 43.0 34.2 102 — 20800 (P1,2) 39.4 32.9 88 — 25600 (P2) 36.1 34.9 89 34000 (P2) 34.9 29.1 94 110

This 34-meter long plate can be used as a wing (skin) panel for very high capacity cargo or passenger aircraft. For this application, the segment with extreme properties X of the plate (corresponding to a high toughness KIC, the static mechanical strength being lower) is fitted on the fuselage side and the segment with extreme properties Z of the plate (corresponding to a high static mechanical strength with a lower toughness KIC) is at the geometric tip of the wing.

The temperature set points as well as the temperature measure on the plate and in the air of the furnace zones during the second aging step are shown in table 3. It includes the temperature profile during the annealing step at 120° C. and 155° C. at a steady temperature state. The temperature of the plate was measured using about forty thermocouples; the values given in table 3 were measured at mid-width.

TABLE 3 Furnace Set Plate Air temperature zone temperature [° C.] temperature [° C.] [° 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 22 155 155.1 154.8 30 155

Additional advantages, features and modifications will readily occur to those skilled in the art. Therefore, the invention in its broader aspects is not limited to the specific details, and representative devices, shown and described herein. Accordingly, various modifications may be made without departing from the spirit or scope of the general inventive concept as defined by the appended claims and their equivalents.

As used herein and in the following claims, articles such as “the”, “a” and “an” can connote the singular or plural.

All documents referred to herein are specifically incorporated herein by reference in their entireties.

Claims

1. A process for manufacturing an aluminium alloy part with structural hardening, comprising:

solution heat treating a semi-finished rolled, extruded or forged product, followed by quenching,
optionally controlling tension with permanent elongation of at least 0.5%, and annealing,
wherein at least a portion of the annealing is conducted in a furnace with a controlled temperature profile comprising at least two zones or groups of zones Z1, Z2 with initial temperatures T1 and T2 and having a temperature variation around the set temperature for each of the temperatures T1 and T2 that does not exceed about ±5° C. within the length of the zones or groups of zones, and wherein the difference between the set values of the initial temperatures T1 and T2 is greater than or equal to about 5° C., and the zones or groups of zones are optionally separated by a zone or a group of zones Z1,2, within which the initial temperature varies from T1 to T2,
and wherein the length parallel to the axis of the furnace of each of the at least two zones or groups of zones Z1 and Z2 is at least one meter.

2. A process according to claim 1, wherein the temperature variation around the set temperature for each of the temperatures T1 and T2 does not exceed about ±4° C. within the length of the at least two zones or groups of zones Z1 and Z2.

3. A process according to claim 1, wherein the difference between the set temperatures T1 and T2 is from about 10° C. to about 80° C.

4. A process according to claim 1, wherein the temperature in at least one of the zones or groups of zones Z1 or Z2 varies as a function of time according to at least two temperature plateaus, and/or according to a temperature ramp with no clearly defined plateau.

5. A process according to claim 1, wherein the annealing in a linear furnace with controlled temperature gradient is followed by at least one forming or machining operation and annealing in a homogeneous furnace.

6. A process according to claim 1, wherein the annealing in a linear furnace with a controlled temperature gradient is preceded by annealing in a homogeneous furnace.

7. A process according to claim 1, wherein the length of the part is at least 7 meters.

8. A process according to claim 1, wherein the aluminium alloy part with structural hardening is monolithic.

9. A process according to claim 1, wherein the aluminium alloy part with structural hardening is assembled starting from at least two aluminium alloy parts with structural hardening.

10. A process according to claim 9, wherein assembly of said at least two parts are made by riveting, bonding, laser beam welding, friction stir welding and/or electron beam welding.

11. A process according to claim 1, wherein the annealing comprises a first homogeneous treatment at a temperature between 115° C. and 125° C. for a duration of from about 2 to about 12 hours, a second treatment during which one end of said part is treated at a temperature from about 115° C. to about 125° C., while the other end of said part is treated at a temperature from about 150° C. to about 160° C., both for a duration of between 8 and 24 hours.

12. A monolithic structural member comprising an aluminium alloy with structural hardening having a length L greater than a width B and thickness E, suitable for aeronautical construction, said monolithic structural member comprising at least two segments P1 and P2 each located on a different length of said structural member, wherein at least one physical property (measured at mid-thickness) of P1 and/or P2 selected from the group consisting of:

a) P1: KIC(L-T)≧38 MPa{square root}m and P2: Rm(L)≧580 MPa
b) P1: KIC(L-T)≧40 MPa{square root}m and P2: Rm(L)≧580 MPa
c) P1: KIC(L-T)≧41 MPa{square root}m and P2: Rm(L)≧580 MPa
d) P1: KIC(L-T)≧42 MPa{square root}m and P2: Rm(L)≧590 MPa
e) P1: KIC(L-T)≧39 MPa{square root}m and P2: Rm(L)≧580 MPa and P2: Rm(TL)≧550 MPa
f) P1: KIC(L-T)≧39 MPa{square root}m and P2: Rm(L)≧580 MPa and P2: Rp0.2(L)≧550 MPa
i) P1: KIC(L-T)≧39 MPa{square root}m and P1: Rm(L)≧530 MPa, and P2: Rm(L)≧580 MPa
j) P1: KIC(L-T)≧40 MPa{square root}m and P1: Rm(L)≧540 MPa, and P2: Rm(L)≧590 MPa
k) P1: Kapp(L-T)(CCT406)>125 MPa{square root}m et P2: Rm(L)>590 MPa.

13. A structural member according to claim 12, wherein A(L)≧9% in segments P1 and P2.

14. A structural member according to claim 13, wherein A(L)≧9% outside segments P1 and P2.

15. A structural member according to claim 12, wherein the length FP1 and FP2 (expressed as a percent of the length L) of said at least two segments P1 and P2 is such that FP1≧25% and FP2>25%.

16. A structural member according to claim 15, wherein FP1≧35% and FP2≧30%.

17. A structural member according to claim 16, wherein FP1≧40% and FP2≧30%.

18. A structural member according to claim 12, wherein the alloy comprises from about 7 to about 15% of zinc, from about 1 to about 3% of copper and/or from about 1.5 to about 3.5% of magnesium.

19. A structural member according to claim 18, wherein zinc is from about 8 to about 13%.

20. A structural member according to claim 19, wherein copper is from about 1.3 to about 2.1%.

21. A structural member according to claim 20, wherein magnesium is from about 1.8 to about 2.7%.

22. A structural member according to claim 12, wherein the length of the part is at least 7 meters.

23. A method for making an aircraft wing panel, wing stringers, wing spars, fuselage stiffeners, fuselage panels and/or butt straps comprising using a structural member according to claim 12.

24. An aircraft comprising at least one wing panel made from a structural member according to claim 12, wherein said segment P1 is located close to the fuselage, and said segment P2 is close to the geometric tip of the wing.

25. A method for forming a hardened aluminium alloy part comprising treating said part in a furnace having at least two zones, each at least one meter in length at a temperature that is maintained approximately constant in said at least two zones.

26. A monolithic structural member prepared from a process of claim 25.

27. An aircraft comprising a structural member of claim 26.

28. A semi-product in which (measured at mid-thickness) comprising an aluminium alloy with structural hardening having a length L greater than a width B and thickness E, suitable for aeronautical construction, said semi-product comprising at least two segments P1 and P2 each located on a different length of said semi-product, wherein at least one physical property (measured at mid-thickness) of P1 and/or P2 selected from the group consisting of:

a) Rp0.2, determined in the L direction or in the LT direction, has a difference p0.2(P2)-Rp0.2(P1) of at least 50 MPa and preferably of at least >75 MPa, and/or
b) Rp0.2, determined in the ST direction, has a difference Rp0.2(P2)-Rp0.2(P1) of at least 30 MPa and preferably at least 50 MPa, and/or
c) KIC, measured in the L-T direction, has a difference KIC(P1)-KIC(P2) of at least 5 MPa{square root}m and preferably of at least 7 MPa{square root}m, and/or
d) Kapp, measured in the L-T direction, has a difference Kapp(P1)-Kapp(P2) of at least 10 MPa{square root}m and preferably of at least 15 MPa{square root}m.

29. A single monolithic structural member that is at least bi-functional.

30. A structural member of claim 29 comprising an alloy selected from the group consisting of 7449, 7349 and 7056.

31. A semi-product of claim 28, wherein said alloy is selected from the group consisting of 7449, 7349 and 7056.

32. A structural member of claim 29 that does not have a continuous variation of properties along its entire length, and said structural member comprises at least two segments in which at least some physical properties thereof are constant over a predetermined length of the segment.

33. A member of claim 32, wherein said predetermined length is at least one meter.

34. A member of claim 32, wherein said predetermined length is at least two meters.

35. A method of claim 25 wherein the product produced thereby does not have a continuous variation of properties along its entire length, and said product produced comprises at least two segments in which at least some physical properties thereof are constant over a predetermined length of the segment.

36. A method of claim 35, wherein said predetermined length is at least one meter.

37. A method of claim 35, wherein said predetermined length is at least two meters.

38. A structural member of claim 26 that does not have a continuous variation of properties along its entire length, and said structural member comprises at least two segments in which at least some physical properties thereof are constant over a predetermined length of the segment.

39. A member of claim 38, wherein said predetermined length is at least one meter.

40. A member of claim 38, wherein said predetermined length is at least two meters.

41. An aircraft comprising a structural member of claim 38.

42. An aircraft comprising a structural member of claim 39.

43. An aircraft comprising a structural member of claim 40.

44. A process of claim 1, wherein said aluminum alloy is selected from the group consisting of 2xxx, 4xxx, 6xxx, 7xxx and 8xxx alloys.

45. A structural member of claim 12 wherein said aluminum alloy is selected from the group consisting of 2xxx, 4xxx, 6xxx, 7xxx and 8xxx alloys.

46. A method of claim 25 wherein said aluminum alloy part comprises 2xxx, 4xxx, 6xxx, 7xxx and/or 8xxx alloys.

47. A monolithic structural member prepared using a method of claim 46.

48. An aircraft comprising a monolithic structural member of claim 47.

Patent History
Publication number: 20050217770
Type: Application
Filed: Mar 23, 2005
Publication Date: Oct 6, 2005
Inventors: Philippe Lequeu (Veyre-Monton), David DuMont (Romans Sur Isere)
Application Number: 11/086,984
Classifications
Current U.S. Class: 148/698.000; 148/701.000; 148/415.000; 148/417.000