U.S. patent application number 11/086984 was filed with the patent office on 2005-10-06 for structural member for aeronautical construction with a variation of usage properties.
Invention is credited to DuMont, David, Lequeu, Philippe.
Application Number | 20050217770 11/086984 |
Document ID | / |
Family ID | 35052970 |
Filed Date | 2005-10-06 |
United States Patent
Application |
20050217770 |
Kind Code |
A1 |
Lequeu, Philippe ; et
al. |
October 6, 2005 |
Structural member for aeronautical construction with a variation of
usage properties
Abstract
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 Z.sub.1, Z.sub.2. The length parallel to the
axis of the linear furnace of each of the at least two zones or
groups of zones Z.sub.1 and Z.sub.2 is generally at least about one
meter.
Inventors: |
Lequeu, Philippe;
(Veyre-Monton, FR) ; DuMont, David; (Romans Sur
Isere, FR) |
Correspondence
Address: |
CONNOLLY BOVE LODGE & HUTZ LLP
SUITE 800
1990 M STREET NW
WASHINGTON
DC
20036-3425
US
|
Family ID: |
35052970 |
Appl. No.: |
11/086984 |
Filed: |
March 23, 2005 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
60555304 |
Mar 23, 2004 |
|
|
|
Current U.S.
Class: |
148/698 ;
148/415; 148/417; 148/701 |
Current CPC
Class: |
C22C 21/10 20130101;
C22F 1/053 20130101; C22F 1/04 20130101 |
Class at
Publication: |
148/698 ;
148/701; 148/415; 148/417 |
International
Class: |
C22F 001/04; C22C
021/10 |
Claims
What is claimed is:
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 Z.sub.1,
Z.sub.2 with initial temperatures T.sub.1 and T.sub.2 and having a
temperature variation around the set temperature for each of the
temperatures T.sub.1 and T.sub.2 that does not exceed about
.+-.5.degree. C. within the length of the zones or groups of zones,
and wherein the difference between the set values of the initial
temperatures T.sub.1 and T.sub.2 is greater than or equal to about
5.degree. C., and the zones or groups of zones are optionally
separated by a zone or a group of zones Z.sub.1,2, within which the
initial temperature varies from T.sub.1 to T.sub.2, and wherein the
length parallel to the axis of the furnace of each of the at least
two zones or groups of zones Z.sub.1 and Z.sub.2 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
T.sub.1 and T.sub.2 does not exceed about .+-.4.degree. C. within
the length of the at least two zones or groups of zones Z.sub.1 and
Z.sub.2.
3. A process according to claim 1, wherein the difference between
the set temperatures T.sub.1 and T.sub.2 is from about 10.degree.
C. to about 80.degree. C.
4. A process according to claim 1, wherein the temperature in at
least one of the zones or groups of zones Z.sub.1 or Z.sub.2 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.degree.
C. and 125.degree. 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.degree. C. to about
125.degree. C., while the other end of said part is treated at a
temperature from about 150.degree. C. to about 160.degree. 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
P.sub.1 and P.sub.2 each located on a different length of said
structural member, wherein at least one physical property (measured
at mid-thickness) of P.sub.1 and/or P.sub.2 selected from the group
consisting of: a) P.sub.1: K.sub.IC(L-T).gtoreq.38 MPa{square
root}m and P.sub.2: R.sub.m(L).gtoreq.580 MPa b) P.sub.1:
K.sub.IC(L-T).gtoreq.40 MPa{square root}m and P.sub.2:
R.sub.m(L).gtoreq.580 MPa c) P.sub.1: K.sub.IC(L-T).gtoreq.41
MPa{square root}m and P.sub.2: R.sub.m(L).gtoreq.580 MPa d)
P.sub.1: K.sub.IC(L-T).gtoreq.42 MPa{square root}m and P.sub.2:
R.sub.m(L).gtoreq.590 MPa e) P.sub.1: K.sub.IC(L-T).gtoreq.39
MPa{square root}m and P.sub.2: R.sub.m(L).gtoreq.580 MPa and
P.sub.2: R.sub.m(TL).gtoreq.550 MPa f) P.sub.1:
K.sub.IC(L-T).gtoreq.39 MPa{square root}m and P.sub.2:
R.sub.m(L).gtoreq.580 MPa and P.sub.2: R.sub.p0.2(L).gtoreq.550 MPa
i) P.sub.1: K.sub.IC(L-T).gtoreq.39 MPa{square root}m and P.sub.1:
R.sub.m(L).gtoreq.530 MPa, and P.sub.2: Rm(L).gtoreq.580 MPa j)
P.sub.1: K.sub.IC(L-T).gtoreq.40 MPa{square root}m and P.sub.1:
R.sub.m(L).gtoreq.540 MPa, and P.sub.2: Rm(L).gtoreq.590 MPa k)
P.sub.1: K.sub.app(L-T)(CCT406)>125 MPa{square root}m et P2:
R.sub.m(L)>590 MPa.
13. A structural member according to claim 12, wherein
A.sub.(L).gtoreq.9% in segments P.sub.1 and P.sub.2.
14. A structural member according to claim 13, wherein
A.sub.(L).gtoreq.9% outside segments P.sub.1 and P.sub.2.
15. A structural member according to claim 12, wherein the length
F.sub.P1 and F.sub.P2 (expressed as a percent of the length L) of
said at least two segments P.sub.1 and P.sub.2 is such that
F.sub.P1.gtoreq.25% and F.sub.P2>25%.
16. A structural member according to claim 15, wherein
F.sub.P1.gtoreq.35% and F.sub.P2.gtoreq.30%.
17. A structural member according to claim 16, wherein
F.sub.P1.gtoreq.40% and F.sub.P2.gtoreq.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
P.sub.1 is located close to the fuselage, and said segment P.sub.2
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
P.sub.1 and P.sub.2 each located on a different length of said
semi-product, wherein at least one physical property (measured at
mid-thickness) of P.sub.1 and/or P.sub.2 selected from the group
consisting of: a) R.sub.p0.2, determined in the L direction or in
the LT direction, has a difference .sub.p0.2(P2)-R.sub.p0.2(P1) of
at least 50 MPa and preferably of at least >75 MPa, and/or b)
R.sub.p0.2, determined in the ST direction, has a difference
R.sub.p0.2(P2)-R.sub.p0.2(P1) of at least 30 MPa and preferably at
least 50 MPa, and/or c) K.sub.IC, measured in the L-T direction,
has a difference K.sub.IC(P1)-K.sub.IC(P2) of at least 5 MPa{square
root}m and preferably of at least 7 MPa{square root}m, and/or d)
K.sub.app, measured in the L-T direction, has a difference
K.sub.app(P1)-K.sub.app(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.
Description
CROSS REFERENCE TO RELATED APPLICATION
[0001] 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
[0002] 1. Field of the Invention
[0003] 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.
[0004] 2. Description of Related Art
[0005] 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.
[0006] 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.
[0007] 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
[0008] 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.
[0009] An object of the present invention was the provision of a
process for manufacturing an aluminium alloy part with structural
hardening, comprising:
[0010] solution heat treating a semi-finished rolled, extruded
and/or forged product, followed by quenching,
[0011] optionally conducting controlled tension with permanent
elongation of at least 0.5%, and
[0012] annealing,
[0013] 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 Z.sub.1, Z.sub.2 with initial
temperatures T.sub.1 and T.sub.2 in which the temperature variation
around the set temperature for each of the temperatures T.sub.1 and
T.sub.2 does not exceed about .+-.5.degree. C. (preferably
.+-.4.degree. C. and even better .+-.3.degree. C.) within the
length of the zones or groups of zones, and further wherein the
difference between the set values of the initial temperatures
T.sub.1 and T.sub.2 are greater than or equal to about 5.degree. C.
(preferably from about 10.degree. C. to about 80.degree. C. and
even better from about 10.degree. C. to about 50.degree. C., and
still better from about 20.degree. C. to about 40.degree. C.). The
zones or groups of zones can optionally be separated by a zone or a
group of zones Z.sub.1,2 called a "transition group" within which
the initial temperature varies from T.sub.1 to T.sub.2 and wherein
the length parallel to the axis of the furnace of each of at least
two zones or groups of zones Z.sub.1 and Z.sub.2 is at least about
one meter (and preferably at least about two meters).
[0014] 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 P.sub.1 and P.sub.2 located on a different
length of the structural member that have mechanical properties
(measured at mid-thickness) selected from the group consisting
of:
[0015] a) P.sub.1: K.sub.IC(L-T).gtoreq.38 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.580 MPa (and preferably .gtoreq.590 MPa
and even better .gtoreq.600 MPa
[0016] b) P.sub.1: K.sub.IC(L-T).gtoreq.40 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.580 MPa (and preferably .gtoreq.590
MPa)
[0017] c) P.sub.1: K.sub.IC(L-T).gtoreq.41 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.580 MPa (and preferably .gtoreq.590
MPa)
[0018] d) P.sub.1: K.sub.IC(L-T).gtoreq.42 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.590 MPa
[0019] e) P.sub.1: K.sub.IC(L-T).gtoreq.39 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.580 MPa and P.sub.2:
R.sub.m(TL).gtoreq.550 MPa
[0020] f) P.sub.1: K.sub.IC(L-T).gtoreq.39 MPa{square root}m and
P.sub.2: R.sub.m(L).gtoreq.580 MPa and P.sub.2:
R.sub.p0.2(L).gtoreq.550 MPa
[0021] i) P.sub.1: K.sub.IC(L-T).gtoreq.39 MPa{square root}m and
P.sub.1: R.sub.m(L).gtoreq.530 MPa, and P.sub.2: Rm(L).gtoreq.580
MPa
[0022] j) P.sub.1: K.sub.IC(L-T).gtoreq.40 MPa{square root}m and
P.sub.1: R.sub.m(L).gtoreq.540 MPa, and P.sub.2: Rm(L).gtoreq.590
MPa
[0023] k) P1: K.sub.app(L-T)(CCT406).gtoreq.125 MPa{square root}m
and P2: R.sub.m(L).gtoreq.590 MPa.
[0024] 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 P.sub.1 is located close to the fuselage and segment
P.sub.2 is located close to a geometric tip of the wing.
[0025] 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
[0026] 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.
[0027] 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.
[0028] FIG. 2 shows the mechanical strength of a 34-meter long
structural member according to the invention.
DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT
[0029] a) Terminology
[0030] 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 R.sub.m, the yield stress R.sub.p02 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 K.sub.IC was measured according to standard
ASTM E 399. The R curve is determined according to standard ASTM
561. The critical stress intensity factor K.sub.C, in other words
the intensity factor that makes the crack unstable, is calculated
starting from the R curve. The stress intensity factor K.sub.CO 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.
K.sub.app denotes the K.sub.CO 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.
[0031] 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.
[0032] The term "machining" includes any process for removal of
material such as turning, milling, drilling, trimming,
electroerosion, grinding, polishing, and chemical milling.
[0033] 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.
[0034] 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.
[0035] 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.
[0036] 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.
[0037] 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.
[0038] 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.
[0039] 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.degree. C., and more
preferably does not exceed about .+-.4.degree. C. In a preferred
embodiment, this difference does not exceed about .+-.3.degree. C.
Certain products may require a temperature variation not exceeding
about .+-.2.degree. 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.
[0040] 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.
[0041] 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.degree. C. or
100.degree. 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.
[0042] 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.
[0043] 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.
[0044] 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.
[0045] 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 P.sub.1 and P.sub.2 on
different lengths of the structural member have physical properties
(measured at mid-thickness) selected from the group formed of:
[0046] a) P.sub.1: K.sub.IC(L-T)>38 MPa{square root}m and
P.sub.2: R.sub.m(L)>580 MPa (and preferably >590 MPa and even
better >600 MPa
[0047] b) P.sub.1: K.sub.IC(L-T)>40 MPa{square root}m and
P.sub.2: R.sub.m(L)>580 MPa (and preferably >590 MPa)
[0048] c) P.sub.1: K.sub.IC(L-T)>41 MPa{square root}m and
P.sub.2: R.sub.m(L)>580 MPa (and preferably >590 MPa)
[0049] d) P.sub.1: K.sub.IC(L-T)>42 MPa{square root}m and
P.sub.2: R.sub.m(L)>590 MPa
[0050] e) P.sub.1: K.sub.IC(L-T)>39 MPa{square root}m and
P.sub.2: R.sub.m(L)>580 MPa and P.sub.2: R.sub.m(TL)>550
MPa
[0051] f) P.sub.1: K.sub.IC(L-T)>39 MPa{square root}m and
P.sub.2: R.sub.m(L)>580 MPa and P.sub.2: R.sub.p0.2(L)>550
MPa
[0052] i) P.sub.1: K.sub.IC(L-T)>39 MPa{square root}m and
P.sub.1: R.sub.m(L)>530 MPa, and P.sub.2: Rm(L)>580 MPa
[0053] j) P.sub.1: K.sub.IC(L-T)>40 MPa{square root}m and
P.sub.1: R.sub.m(L)>540 MPa, and P.sub.2: R.sub.m(L)>590
MPa
[0054] k) P1: K.sub.app(L-T)(CCT406)>125 MPa{square root}m et
P2: R.sub.m(L)>590 MPa.
[0055] 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 P.sub.1 and P.sub.2. 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 P.sub.1 and P.sub.2. It is
possible to manufacture semi-products in which (measured at
mid-thickness)
[0056] a) R.sub.p0.2, determined in the L direction or in the LT
direction, has a difference p.sub.0.2(P2)-R.sub.p0.2(P1) of at
least 50 MPa and preferably of at least >75 MPa, and/or
[0057] b) R.sub.p0.2, determined in the ST direction, has a
difference R.sub.p0.2(P2)-R.sub.p0.2(P1) of at least 30 MPa and
preferably at least 50 MPa, and/or
[0058] c) K.sub.IC, measured in the L-T direction, has a difference
K.sub.IC(P1)-K.sub.IC(P2) of at least 5 MPa{square root}m and
preferably of at least 7 MPa{square root}m, and/or
[0059] d) K.sub.app, measured in the L-T direction, has a
difference K.sub.app(P1)-K.sub.app(P2) of at least 10 MPa{square
root}m and preferably of at least 15 MPa{square root}m.
[0060] 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.
[0061] 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.
[0062] 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.
[0063] 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.
[0064] 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:
[0065] A first homogenous step at a temperature between 115.degree.
C. and 125.degree. 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.degree. C. and 125.degree. C., while the
another segment or the other end is treated at a temperature
between 150.degree. C. and 160.degree. C., both for a duration of
between 8 and 24 hours.
[0066] This annealing is particularly suitable for products made of
7xxx alloy, and particularly 7349, 7449 or 7056 alloy.
[0067] In another advantageous embodiment of this invention,
annealing is done at about 120.degree. C. (i.e. under-aging) on one
segment or end P.sub.1 of a product made of 2xxx alloy (such as
2024 or 2023), while annealing to the peak mechanical strength
(temper T851) at about 190.degree. C. is carried out on another
segment or the other end P.sub.2. In a variant of this embodiment,
the segment or end which is not peak-aged (i.e. P.sub.1) is aged at
about 100.degree. C. (or 80.degree. C.).
[0068] 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.degree. C., while over-annealing (temper T7651, T7451 or
T7351) is carried out at another segment or the other end in two
plateaux at 120.degree. C. and 150-165.degree. C.
[0069] 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.degree. C., while
over annealing (state T7851) is carried out in two plateaux at the
other end.
[0070] 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.
[0071] 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.
[0072] Another specific advantage of the process according to the
invention is precise control of properties in the transition
segment P.sub.1,2 between two groups of segments P.sub.1 and
P.sub.2 (there may be two or more groups, depending on the number
of groups of temperature zones), wherein P.sub.1 and P.sub.2 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 R.sub.m(L) and the
toughness K.sub.IC(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.
[0073] In particular, structural members include:
[0074] upper or lower wing (skin) panels;
[0075] upper or lower wing stringers;
[0076] wing spars;
[0077] fuselage stiffeners;
[0078] butt straps, particularly butt straps for upper and lower
wing (skin) panels;
[0079] fuselage panels.
[0080] 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
P.sub.1 and P.sub.2 in which the length F.sub.P1 and F.sub.P2
(expressed in percent of the total length L) of the said at least
two segments P.sub.1 and P.sub.2 is such that F.sub.P1>25% and
F.sub.P2>25% and preferably F.sub.P1>30% and F.sub.P2>30%.
In other embodiments, F.sub.P1>35% and F.sub.P2>30% or
F.sub.P1>40% and F.sub.P2>30%.
[0081] 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 P.sub.1 is located close to the
fuselage, and segment P.sub.2 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.
[0082] 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.
[0083] 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).
[0084] The invention will be better understood after reading the
following example that is in no way limiting.
EXAMPLE
[0085] A 36-meter long, 2.5-meter wide and 30 mm thick plate is
made by hot rolling of a rolling plate.
[0086] The alloy composition was:
[0087] 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.
[0088] The rolling plate was homogenised for 14 hours at
475.degree. C. The input temperature to the hot roller was
428.degree. C., and the output temperature of the hot rolled plate
was 401.degree. C. The plate was solution heat treated, quenched
and tensioned under the following conditions: holding for 6 hours
at 471.degree. C., quenching in water at a temperature between
about 15 and 16.degree. 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.degree. C.
around the set value.
[0089] The annealing treatment consisted of a first homogenous
treatment step for 6 hours at 120.degree. C. ("first plateau") and
was immediately followed by a second step during which one 18-meter
geometric tip (called Z.sub.1, corresponding to 15 furnace zones)
was treated for 15 hours at 155.degree. C. ("second plateau"
preceded by an adjustment period of about 1 hour), while the other
10.8-meter geometric tip (called Z.sub.2, corresponding to 9
furnace zones) was held for 16 hours at 120.degree. C. The
transition zone between these two tips was 7.2 meters long (called
Z.sub.1,2 corresponding to 6 furnace zones).
[0090] After this second step, the electrical conductivity of the
plate was measured at different locations:
[0091] Segment P.sub.1: between 18.2 and 19.5 MS/m
[0092] Segment P.sub.2: between 22.5 and 23.5 MS/m
[0093] Segment P.sub.1,2: between 18.2 and 23.6 MS/m.
[0094] The plate was then subjected to a third annealing step,
namely homogeneous annealing consisting of a temperature increase
to 148.degree. C. for 1h30, followed by holding at 150.degree. C.
for 15 hours. This third step was intended to simulate age forming
or annealing after the structural member was shaped.
[0095] 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 R.sub.P0.2 in the L and LT direction, values have also
been obtained by compression; these values are put between brackets
in table 1.
1TABLE 1 LT TC Position [mm] L (long) (long transverse) (short
transverse) in the length direction direction direction of a 34 m
R.sub.m R.sub.p0.2 A R.sub.m R.sub.p0.2 A R.sub.m R.sub.p0.2 A
panel [MPa] [MPa] [%] [MPa] [MPa] [%] [MPa] [MPa] [%] 0 (P.sub.1)
561 517 13.5 550 506 12.5 550 495 8.5 (509) (519) 13600 (P.sub.1)
565 522 13.5 553 511 12.5 548 502 8.5 (513) (528) 16000 (P.sub.1)
556 509 13.5 547 501 12.5 540 500 8.5 (500) (514) 18400 (P.sub.1,2)
566 523 13.5 559 519 12.5 546 498 7.5 (527) (538) 20800 (P.sub.1,2)
612 587 12.0 598 575 11.5 590 545 7.0 (573) (593) 25600 (P.sub.2)
621 598 12.5 607 585 11.5 595 554 6.5 (590) (605) 34000 (P.sub.2)
624 602 12.1 608 586 11.5 599 558 6.1 (594) (607)
[0096] The toughness results K.sub.IC and K.sub.app (the latter
obtained on a CT127 type test piece as well as on a CCT406 type
test piece) are given in table 2
2TABLE 2 Position [mm] along K.sub.app(L - T) K.sub.app(L - T) the
length of a K.sub.IC (L - T) K.sub.IC (T - L) (CT127) (CCT406) 34 m
panel [MPa{square root}m] [MPa{square root}m] [MPa{square root}m]
[MPa{square root}m] 0 (P.sub.1) 43.8 36.1 106 132 13600 (P.sub.1)
45.8 38.1 108 -- 16000 (P.sub.1) 46.7 37.3 99 -- 18400 (P.sub.1,2)
43.0 34.2 102 -- 20800 (P.sub.1,2) 39.4 32.9 88 -- 25600 (P.sub.2)
36.1 34.9 89 34000 (P.sub.2) 34.9 29.1 94 110
[0097] 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 K.sub.IC, 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 K.sub.IC) is
at the geometric tip of the wing.
[0098] 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.degree. C. and 155.degree.
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.
3TABLE 3 Furnace Set Plate Air temperature zone temperature
[.degree. C.] temperature [.degree. C.] [.degree. 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
[0099] 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.
[0100] As used herein and in the following claims, articles such as
"the", "a" and "an" can connote the singular or plural.
[0101] All documents referred to herein are specifically
incorporated herein by reference in their entireties.
* * * * *