U.S. patent application number 11/398664 was filed with the patent office on 2006-08-17 for al-zn-mg-cu alloys and products with high mechanical characteristics and structural members suitable for aeronautical construction made thereof.
Invention is credited to Frank Eberl, Christophe Sigli, Sjoerd van der Veen, Timothy Warner.
Application Number | 20060182650 11/398664 |
Document ID | / |
Family ID | 28052134 |
Filed Date | 2006-08-17 |
United States Patent
Application |
20060182650 |
Kind Code |
A1 |
Eberl; Frank ; et
al. |
August 17, 2006 |
Al-Zn-Mg-Cu alloys and products with high mechanical
characteristics and structural members suitable for aeronautical
construction made thereof
Abstract
The present invention further relates to 7xxx alloys and
products produced therewith that can be flat rolled, extruded or
forged, as well as associated methods. Al--Zn--Mg--Cu alloys of the
present invention preferably comprise (in mass percentage): a) Zn
8.3-14.0 Cu 0.3-2.0 Mg 0.5-4.5 Zr 0.03-0.15 Fe+Si<0.25 b) at
least one element selected from the group consisting of Sc, Hf, La,
Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y and Yb, where the content of
each of the elements, if selected, is between 0.02 and 0.7%, c)
remainder aluminum and inevitable impurities. The present invention
further is directed to products wherein Mg/Cu>2.4 and (7.9-0.4
Zn)>(Cu+Mg)>(6.4-0.4 Zn). The disclosed products can be used
for example, as structural members in aeronautical construction,
especially as stiffeners capable for use in fuselages of civilian
and other aircrafts as well as in related applications.
Inventors: |
Eberl; Frank; (Grenoble,
FR) ; Sigli; Christophe; (Grenoble, FR) ;
Warner; Timothy; (Voreppe, FR) ; Veen; Sjoerd van
der; (Clermont-Ferrand, FR) |
Correspondence
Address: |
BAKER DONELSON BEARMAN CALDWELL & BERKOWITZ, PC
555 11TH STREET, NW
6TH FLOOR
WASHINGTON
DC
20004
US
|
Family ID: |
28052134 |
Appl. No.: |
11/398664 |
Filed: |
April 6, 2006 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
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10406610 |
Apr 4, 2003 |
|
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11398664 |
Apr 6, 2006 |
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Current U.S.
Class: |
420/532 |
Current CPC
Class: |
C22C 21/10 20130101;
C22F 1/053 20130101 |
Class at
Publication: |
420/532 |
International
Class: |
C22C 21/10 20060101
C22C021/10 |
Foreign Application Data
Date |
Code |
Application Number |
Apr 5, 2002 |
FR |
02 04250 |
Claims
1. (canceled)
2. A fuselage structure of according to claim 79, wherein
Mg/Cu>2.8.
3. A fuselage structure according to claim 79, wherein
Mg/Cu>3.5.
4. A fuselage structure according to claim 79, wherein
Mg/Cu>4.0.
5. A fuselage structure according to claim 79, wherein the maximum
content of the elements Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho,
Er, Y and Yb does not exceed 1.5% in total.
6. A fuselage structure according to claim 79, comprising only Ti
from said group consisting of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb,
Dy, Ho, Er, Y and Yb.
7. A fuselage structure comprising a stringer comprising an
Al--Zn--Mg--Cu alloy, comprising (in mass percentage): a) Zn
9.5-14.0 Cu 0.3-2.0 Mg 0.5-4.5 Fe+Si<0.25 b) at least one
element selected from the group consisting of Zr Sc, Hf, La, Ti,
Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y and Yb, the content of each of
said elements, if selected, being between 0.02 and 0.7%, c)
remainder aluminum and inevitable impurities, and wherein
Mg/Cu>2.4 and (7.9-0.4 Zn)>(Cu+Mg)>(6.4-0.4 Zn).
8. A fuselage structure according to claim 7, wherein the maximum
content of the elements Zr, Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy,
Ho, Er, Y and Yb does not exceed 1.5% in total.
9. A fuselage structure according to claim 79, wherein
Zn>9.0%.
10. A fuselage structure according to claim 9, wherein
Zn>9.5%.
11. A fuselage structure according to claim 9, wherein zinc is
between 9.0% and 11.0%.
12. A fuselage structure according to claim 79, wherein
Cu>0.6%.
13. A fuselage structure according to claim 7, wherein
Cu>0.6%.
14. A fuselage structure according to claim 12, wherein Cu 0.6-1.2%
and Mg 2.2-3.0%.
15. A fuselage structure according to claim 12, wherein Cu 0.8-1.5%
and Mg 2.2-3.0%.
16. A fuselage structure according to claim 79, wherein Mg
0.5-3.6%.
17. A fuselage structure according to claim 7, wherein Mg
0.5-3.6%.
18. A fuselage structure according to claim 7, wherein
Mg/Cu>2.8.
19. A fuselage structure according to claim 7, wherein
Mg/Cu>3.5.
20. A fuselage structure according to claim 7, wherein
Mg/Cu>4.0.
21. A fuselage structure according to claim 18 wherein said alloy
comprises a rolled, forged or extruded material comprising an alloy
according to claim 18.
22. A fuselage structure according to claim 79, wherein
Mg>1.95+0.5 (Cu-2.3)+0.16 (Zn-6)+1.9 (Si-0.04).
23. A fuselage structure according to claim 7, wherein
Mg>1.95+0.5 (Cu-2.3)+0.16 (Zn-6)+1.9 (Si-0.04).
24. A fuselage structure according to claim 79, wherein the maximum
mass percentage of the following elements is not exceeded: Cr 0.40
Mn 0.60 Sc 0.50 Zr 0.15 Hf 0.60 Ti 0.15 Ce 0.35 Nd 0.35 La Eu 0.35
Gd 0.35 Tb 0.35 Dy 0.40 Ho 0.40 Er 0.40 Yb 0.40 and Y 0.20.
25. A fuselage structure according to claim 7, wherein the maximum
mass percentage of the following elements is not exceeded: Cr 0.40
Mn 0.60 Sc 0.50 Zr 0.15 Hf 0.60 Ti 0.15 Ce 0.35 Nd 0.35 La Eu 0.35
Gd 0.35 Tb 0.35 Dy 0.40 Ho 0.40 Er 0.40 Yb 0.40 and Y 0.20.
26. A fuselage structure according to claim 79, further comprising
at least one element selected from the group consisting of Ag, Sn,
Cd, Ge and In, the content of each of said at least one element, if
selected, being present in an amount of between 0.02% and
0.15%.
27. A fuselage structure, according to claim 7, further comprising
at least one element selected from the group consisting of Ag, Sn,
Cd, Ge and In, the content of each of said at least one element, if
selected, being present in an amount of between 0.02% and
0.15%.
28. A fuselage structure according to claim 79, wherein said alloy
is formed into an extruded product prepared from said alloy,
wherein said product exhibits in temper T6511, measured on test
pieces cut from a plane zone of the profile, a) a bending angle,
determined at 130.degree. C. by means of a three point bending test
according to DIN 50 111 (section 3.1) on a specimen of 1.6 mm
thickness and expressed as the mean value computed from several
individual measurements performed on specimens cut from different
locations over the length of the profile, of at least 34.degree.,
and b) a tensile yield strength (TYS) of at least 720 MPa,
29. A fuselage structure according to claim 7, wherein said alloy
is formed into an extruded product prepared from said alloy,
wherein said product exhibits in temper T6511, measured on test
pieces cut from a plane zone of the profile, a) a bending angle,
determined at 130.degree. C. by means of a three point bending test
according to DIN 50 111 (section 3.1) on a specimen of 1.6 mm
thickness and expressed as the mean value computed from several
individual measurements performed on specimens cut from different
locations over the length of the profile, of at least 34.degree.,
and b) a tensile yield strength (TYS) of at least 720 MPa,
30. A fuselage structure according to claim 28, wherein said
extruded product has a bending angle of at least 35.degree. and a
TYS of at least 750 MPa.
31. A fuselage structure according to claim 29, wherein said
extruded product has a bending angle of at least 35.degree. and a
TYS of at least 750 MPa.
32. A fuselage structure according to claim 28, wherein said
product exhibits in temper T76511, measured on test pieces cut from
a plane zone of the profile, a) a bending angle, determined at
130.degree. C. by means of a three point bending test according to
DIN 50 111 (section 3.1) on a specimen of 1.6 mm thickness and
expressed as the mean value computed from several individual
measurements performed on specimens cut from different locations
over the length of the profile, of at least 37.degree., and b) an
ultimate tensile strength (UTS) of at least 670 MPa.
33. A fuselage structure according to claim 29, wherein said
product exhibits in temper T76511, measured on test pieces cut from
a plane zone of the profile, a) a bending angle, determined at
130.degree. C. by means of a three point bending test according to
DIN 50 111 (section 3.1) on a specimen of 1.6 mm thickness and
expressed as the mean value computed from several individual
measurements performed on specimens cut from different locations
over the length of the profile, of at least 37.degree., and b) an
ultimate tensile strength (UTS) of at least 670 MPa.
34. A fuselage structure according to claim 32, wherein said
product has a stress corrosion resistance, determined by an EXCO
test according to ASTM G34 in the T6511 temper on unmachined test
pieces, of at least level EB.
35. A fuselage structure according to claim 33, wherein said
product has a stress corrosion resistance, determined by an EXCO
test according to ASTM G34 in the T6511 temper on unmachined test
pieces, of at least level EB.
36. A structural aircraft member manufactured from a fuselage
structure of claim 79.
37. A structural aircraft member manufactured from a fuselage
structure of claim 7.
38. A structural aircraft member according to claim 36, wherein
said structural member comprises a fuselage stringer.
39. A structural aircraft member according to claim 37, wherein
said structural member comprises a fuselage stringer.
40. (canceled)
41. (canceled)
42. (canceled)
43. A structural member according to claim 36 , having a ultimate
tensile strength in the T76511 temperof at least 670 MPa.
44. (canceled)
45. A structural member according to claim 37, having a ultimate
tensile strength in the T76511 temper of at least 670 MPa.
46. (canceled)
47. (canceled)
48. (canceled)
49. A structural member according to claim 36, having a tensile
yield strength in the T76511 temper of at least 640 MPa.
50. A structural member according to claim 37, having a tensile
yield strength in the T76511 temper of at least 640 MPa.
51. (canceled)
52. (canceled)
53. A structural member according to claim 36, having a tensile
yield strength of at least 660 MPa.
54. A structural member according to claim 37, having a tensile
yield strength of at least 660 MPa.
55. A structural aircraft member according to claim 36, wherein
said structural member comprises a floor beam.
56. A structural aircraft member according to claim 37, wherein
said structural member comprises a floor beam.
57. An aircraft comprising a fuselage assembled from a plurality of
stringers and a plurality of sheets, wherein at least part of said
stringers comprise structural members according to claim 36.
58. An aircraft comprising a fuselage assembled from a plurality of
stringers and a plurality of sheets, wherein at least part of said
stringers comprise structural members according to claim 37.
59. A fuselage structure comprising an extruded product comprising
a bending angle of at least 34.degree. in the T6511 temper
determined at 130.degree. C. by a 3 point bending test according to
DIN 50111 on a sample thereof 1.6 mm in thickness cut from a plane
area of said product, and said product further comprising a TYS of
at least 720 MPa.
60. A fuselage structure comprising an extruded product comprising
a bending angle of at least 360 in the T76511 temper determined at
130.degree. C. by a 3 point bending test according to DIN 50111 on
a sample thereof 1.6 mm in thickness cut from a plane area of said
product, and said product further comprising a TYS of at least 660
MPa.
61. A fuselage structure of claim 60 wherein said extruded product,
further comprises a corrosion resistance rating of at least EB
(EXCO test according to ASTM G34).
62. (canceled)
63. (canceled)
64. (canceled)
65. (canceled)
66. A fuselage structure of claim 79 comprising a rolled
product.
67. A fuselage structure of claim 7 comprising a rolled
product.
68. A fuselage structure of claim 79 comprising an extruded
product.
69. A fuselage structure of claim 7 comprising an extruded
product.
70. A fuselage structure of claim 79 comprising a forged
product.
71. A fuselage structure of claim 7 comprising a forged
product.
72. A stringer comprising a 7xxx alloy wherein after being
subjected to homogenization and scalping and being extruded at
400.degree. C. at a speed of below 0.50m/min, extrusion forces
required to obtain extrusion profiles of said alloy decrease as the
content of magnesium is increased in said alloy.
73. A stringer comprising a 7xxx alloy comprising up to 0.15% Sc,
up to 12% Zn, and a Mg/Cu ratio of >2.4, said alloy possessing
an R.sub.p0,2(L) of from 680-700 MPa and R.sub.M(L) of from 686-700
MPa.
74. A fuselage structure comprising at least one stringer that is
not 2024, wherein if prepared with a stringer pitch of 200 mm and a
stiffening ratio of 0.25, said structure exhibits an SIF that is
reduced up to 5% as compared to a fuselage structure with 2024 T3
alloy stringers.
75. A fuselage structure comprising at least one stringer that is
not 2024, wherein if prepared with a stringer pitch of 200 mm and a
stiffening ratio of 0.25, said structure exhibits an SIF that is
reduced up to 15% as compared to a fuselage structure with 2024
plastic domain stringers.
76. A stringer that is not 2024, wherein if utilized in a fuselage
with a stringer pitch of 200 mm and a stiffening ratio of 0.25,
said structure exhibits an SIF that is reduced up to 5% as compared
to a fuselage structure with 2024 T3 alloy stringers.
77. A stringer that is not 2024, wherein if utilized in a fuselage
with a stringer pitch of 200 mm and a stiffening ratio of 0.25,
said structure exhibits an SIF that is reduced up to 15% as
compared to a fuselage structure with 2024 plastic domain
stringers.
78. A stringer geometry selected from one or more of the following:
TABLE-US-00020 Z1 Z2 Z3 Z4 Z5 Z6 Z7 Z8 Free flange 12.7 12.7 12.7
12.7 12.7 12.7 12.7 12.7 width [mm] Fastened flange 25.4 25.4 25.4
25.4 25.4 25.4 25.4 25.4 width [mm] Height [mm] 38.1 38.1 38.1 38.1
38.1 38.1 38.1 38.1 Free flange 1.0 1.5 1.5 2.0 1.0 1.5 1.5 1.5
thickness [mm] Fastened flange 1.0 1.0 1.5 1.5 1.0 1.0 1.0 1.5
thickness [mm] Web thickness 1.0 1.0 1.0 1.0 1.5 1.5 1.5 1.5 [mm]
Section [mm.sup.2] 76 83 95 102 95 102 102 114 Equivalent 1.0 1.1
1.3 1.3 1.3 1.3 1.3 1.5 thickness [mm]
79. A fuselage structure comprising at least one stringer
comprising an Al--Zn--Mg--Cu alloy, comprising (in mass
percentage): a) Zn 8.3-14.0 Cu 0.3-2.0 Mg 0.5-4.5 Zr 0.03-0.15
Fe+Si<0.25 b) at least one element selected from the group
consisting of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y and
Yb, the content of each of said elements, if selected, being
between 0.02 and 0.7%, c) remainder aluminum and inevitable
impurities, and wherein Mg/Cu>2.4 and (7.9-0.4
Zn)>(Cu+Mg)>(6.4-0.4 Zn).
80. A fuselage structure according to claim 79, wherein if prepared
with a stringer pitch of 200 mm and a stiffening ratio of 0.25,
said structure exhibits an SIF that is reduced up to 5% as compared
to a fuselage structure with 2024 T3 alloy stringers.
81. A fuselage structure according to claim 79, wherein if prepared
with a stringer pitch of 200 mm and a stiffening ratio of 0.25,
said structure exhibits an SIF that is reduced up to 15% as
compared to a fuselage structure with 2024 plastic domain
stringers.
82. A fuselage structure according to claim 79 wherein said at
least one stringer geometry is selected from one or more of the
following: TABLE-US-00021 Z1 Z2 Z3 Z4 Z5 Z6 Z7 Z8 Free flange 12.7
12.7 12.7 12.7 12.7 12.7 12.7 12.7 width [mm] Fastened flange 25.4
25.4 25.4 25.4 25.4 25.4 25.4 25.4 width [mm] Height [mm] 38.1 38.1
38.1 38.1 38.1 38.1 38.1 38.1 Free flange 1.0 1.5 1.5 2.0 1.0 1.5
1.5 1.5 thickness [mm] Fastened flange 1.0 1.0 1.5 1.5 1.0 1.0 1.0
1.5 thickness [mm] Web thickness 1.0 1.0 1.0 1.0 1.5 1.5 1.5 1.5
[mm] Section [mm.sup.2] 76 83 95 102 95 102 102 114 Equivalent 1.0
1.1 1.3 1.3 1.3 1.3 1.3 1.5 thickness [mm]
Description
CLAIM FOR PRIORITY
[0001] The present invention claims priority under 35 U.S.C. .sctn.
119 from French Patent Application No. 02 04250 filed Apr. 5, 2002,
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 Al--Zn--Mg--Cu
type alloys with high mechanical characteristics, typically having
a Zn content greater than 8.3%, as well as to products and
structural members suitable for aeronautical products and/or
constructions manufactured from such products.
[0004] 2. Description of the Related Art
[0005] Al--Zn--Mg--Cu alloys (belonging to the family of 7xxx
alloys) are currently in use in aeronautical construction, and are
particularly used in the construction of civilian aircraft wings.
For the wing exterior, a skin of plate made in alloys such as 7150,
7055, 7449 is often used, and optionally includes stiffeners (also
called stringers) made from profiles in 7150, 7055, or 7449 alloys.
7150, 7050 and 7349 alloys are also commonly used for the
manufacture of fuselage stiffeners or stringers.
[0006] Some of these alloys have been known for decades, such as
for example 7075 and 7175 alloys (zinc content between 5.1 and 6.1%
by weight), 7050 (zinc content between 5.7 and 6.7%), 7150 (zinc
content between 5.9 and 6.9%) and 7049 (zinc content between 7.2
and 8.2%). Such alloys have a high tensile yield strength, as well
as good fracture toughness and good resistance to stress corrosion
and to exfoliation corrosion. More recently, it has appeared that
for certain applications the use of an alloy with a higher zinc
content can have advantages, since such an alloy allows the tensile
yield strength to be increased further. 7349 and 7449 alloys
contain between 7.5 and 8.7% zinc. Wrought alloys still higher in
zinc have been described in the literature, but have not been
deemed in aeronautical construction.
[0007] The article "Microstructure and Properties of a New
Super-High-Strength Al--Zn--Mg--Cu alloy C912" by Y. L. Wu et al.,
published in Materials & Design, vol 18, p. 211-215 (1998)
describes an alloy including Zn 8.7%, Mg 2.6%, Cu 2.5%, Si and Fe
each <0.05%, which is for the manufacture of structural members
for wings and fuselage.
[0008] U.S. Pat. No. 5,560,789 (Pechiney) discloses an alloy
including Zn 10.7%, Mg 2.84% and Cu 0.92% which is transformed by
extrusion. The alloying elements of this alloy are very high in
zinc, magnesium and copper, and as such, are difficult to put into
solid solution because the temperature of solution heat treatment
is limited by the melting point of intermetallic phases, which have
the lowest incipient melting point. As a consequence, products
formed with such an alloy have a high mechanical strength, but a
very low elongation at fracture due to the presence of coarse
precipitates. Thus, such a product has a low formability.
[0009] U.S. Pat. No. 5,221,377 (Aluminum Company of America)
discloses several Al--Zn--Mg--Cu alloys with a zinc content of up
to 11.4% and with a rather high copper content. These alloys are
difficult to cast and the alloying elements are difficult to put
into solid solution, which favours the presence of coarse
precipitates, which are not welcome.
[0010] Moreover, it has been proposed to utilise Al--Zn--Mg--Cu
alloys with a high zinc content to manufacture hollow bodies
intended to resist high pressures, such as for example cylinders
for compressed gas. European Patent Application No. EP 020 282 Al
(Societe Metallurgique de Gerzat) discloses alloys with a zinc
content of between 7.6% and 9.5%. European Patent Application No.
EP 081 441 A1 (Societe Metallurgique de Gerzat) discloses a process
for obtaining such cylinders. European Patent Application No. EP
257 167 A1 (Societe Metallurgique de Gerzat) states that none of
the known Al--Zn--Mg--Cu alloys can safely and reproducibly satisfy
the strict technical demands imposed by this specific applications.
This document proposes moving towards a lower zinc content, namely
between 6.25% and 8.0%.
[0011] The teaching of the above described documents is specific to
the problem of cylinders for compressed gas, particularly
concerning maximizing of the bursting pressure of these cylinders,
and thus, is not relevant to other wrought products because the
required mechanical properties would be completely different, among
other things.
[0012] In general, in Al--Zn--Mg--Cu alloys, a high zinc content,
and also a high Mg and Cu content are typically required in order
to obtain good static mechanical characteristics (ultimate tensile
strength (R.sub.m or UTS) and tensile yield strength (R.sub.p0.2 or
TYS)), but this is possible only if these elements can be put into
solid solution. It is also well known (see for example U.S. Pat.
No. 5,221,377) that discloses that when the zinc content of an
alloy of the 7xxx family is increased beyond around 7 to 8%, then
problems associated with insufficient resistance to exfoliation
corrosion and stress corrosion will arise. More generally, it is
known that the most charged Al--Zn--Mg--Cu alloys are likely to
pose corrosion problems. These problems are generally resolved by
using particular thermal or thermo-mechanical treatments,
especially by pushing the aging treatment beyond the peak, for
example during T7 type treatment. But such treatments can then
cause a corresponding drop in the static mechanical
characteristics. In other words, for a minimum level of resistance
to corrosion to be obtained, when optimising an Al--Zn--Mg--Cu
alloy one must find a compromise between static mechanical
characteristics (TYS R.sub.p0.2, UTS R.sub.m, elongation at
fracture A) and damage tolerance characteristics (fracture
toughness, crack propagation rate etc.). According to the minimal
level of resistance to corrosion to be envisaged, either (i) a
temper close to peak strength is utilised (T6 tempers), which
generally offers (ii) acceptable toughness--TYS compromise
favouring static mechanical characteristics, or (iii) annealing is
pushed beyond peak strength (T7 tempers), by seeking a compromise
favouring fracture toughness.
[0013] Whichever approach is used, the manufacture and use of such
products poses at least two problems. On one hand, alloys with a
high zinc and magnesium content are difficult to cast and to
transform, especially by extrusion, rolling or forging. For
example, the maximum force that can be supplied by an extrusion
press can be a limiting factor. In particular, among 7xxx series
alloys, 7349 and 7449 alloys require very high extrusion forces. On
the other hand, for certain applications, the formability of rolled
and extruded products is critical. This is especially true with
respect to fuselage stringers. Therefore, up until now, alloys that
could be developed having a mechanical strength still higher than
7349 and 7449 alloys would likely be difficult to cast and to
transform, and products made therefrom would tend to have a low
formability.
SUMMARY OF THE INVENTION
[0014] A problem which the present invention attempted to resolve
was therefore to obtain a novel alloy and associated novel wrought
Al--Zn--Mg--Cu type products with a high zinc content, (i.e.
greater than 8.3%) and especially extruded products, as well as
their associated methods. Products of the present invention
preferably possess a very high ultimate tensile strength (UTS), a
very high tensile yield strength (TYS) as well as adequate
resistance to corrosion and a high formability, and also are
capable of being manufactured industrially under conditions of
highest reliability compatible with the severe requirements of the
aeronautical industry.
[0015] The present inventors have found that these and other
problems can be addressed inter alia by finely adjusting the
concentration of Zn, Cu and/or Mg as well as controlling the
content of certain impurities (particularly Fe and Si), and further
by optionally adding other elements.
[0016] In accordance with these and other objects, one embodiment
of the present invention is directed to an Al--Zn--Mg--Cu alloy
that can be rolled, extruded and/or forged, characterised in that
it comprises (in mass percentage): [0017] a) Zn 8.3-14.0 Cu 0.3-2.0
[0018] Mg 0.5-4.5, preferably 0.5-3.6 [0019] Zr 0.03-0.15
Fe+Si<0.25 [0020] b) at least one element selected from the
group consisting of Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er,
Y and Yb, the content of each of said elements, if selected, being
between 0.02 and 0.7%, [0021] c) remainder aluminum and inevitable
impurities, [0022] and wherein [0023] Mg/Cu>2.4 and [0024]
(7.9-0.4 Zn)>(Cu+Mg)>(6.4-0.4 Zn).
[0025] In yet further accordance with the present invention, there
is provided another embodiment directed to an Al--Zn--Mg--Cu alloy
that can be rolled, extruded and/or forged, characterised in that
it comprises (in mass percentage): [0026] a) Zn 9.5-14.0 Cu 0.3-2.0
[0027] Mg 0.5-4.5 and preferably 0.5-3.6 [0028] Fe+Si<0.25
[0029] b) at least one element selected from the group consisting
of Zr, Sc, Hf, La, Ti, Ce, Nd, Eu, Gd, Tb, Dy, Ho, Er, Y, Yb, Cr
and Mn, the content of each of said elements, if selected, being
between 0.02 and 0.7%, [0030] c) remainder aluminum and inevitable
impurities, [0031] and wherein [0032] Mg/Cu>2.4 and [0033]
(7.9-0.4 Zn)>(Cu+Mg)>(6.4-0.4 Zn).
[0034] In yet further accordance with the present invention, there
is provided another embodiment directed to a structural member for
aeronautical construction incorporating at least one product, and
particularly to a structural member suitable for the construction a
fuselage of an aircraft, such as a fuselage stringer.
[0035] 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 FIGURES
[0036] FIG. 1 shows the section of profile T1 according to an
embodiment of the present invention.
[0037] FIG. 2 shows the section of profile T2 according to an
embodiment of the present invention.
[0038] FIG. 3 shows the section of profile T3 according to an
embodiment of the present invention.
[0039] FIG. 4 shows the section of profile T4 according to an
embodiment of the present invention.
[0040] FIG. 5 shows the section of profile T5 according to an
embodiment of the present invention.
[0041] FIG. 6 diagrammatically illustrates the zone of a fuselage
stringer which has been formed by joggling according to an
embodiment of the present invention.
[0042] FIG. 7 diagrammatically illustrates the location of sampling
on profile T1, where a test piece for the 3 point bending test is
cut.
[0043] FIG. 8 diagrammatically illustrates the definition of the
bending angle.
[0044] FIG. 9 diagrammatically illustrates the important parameters
of the 3 point bending test.
[0045] FIG. 10 diagrammatically illustrates a two stringer bay
crack.
[0046] FIG. 11a and 11b diagrammatically illustrate a buckling
test.
[0047] FIG. 12 compares the predicted crippling stress for
different Z stringers according to the invention (grey bars) and
according to prior art (white bars) for the same geometry.
DETAILED DESCRIPTION OF THE INVENTION
[0048] In FIGS. 1, 2, 3, 4 and 5, the dimensions are approximate
values expressed in millimetres. In FIGS. 1, 2, 3 and 4, letter (a)
designates the foot section, and letter (b) the top section of the
profile.
[0049] In FIG. 6, the reference letters are as follows [0050] a
Joggling depth [0051] b Joggling width [0052] c Upper foot:
important plane deformation [0053] d Lower foot: important plane
deformation
[0054] Unless indicated otherwise, the chemical compositions are
given as percentages by weight based on total weight of the article
being described. Therefore, in a mathematical formula, "0.4 Zn"
means "0.4 times the zinc content, expressed in percentage by
weight". This also applies to other chemical elements as well as
Zn. The alloy designations follow the rules of The Aluminum
Association. The metallurgical tempers are as defined in the
European Standard EN 515. Unless indicated otherwise, the static
mechanical characteristics, i.e. the ultimate tensile strength
R.sub.m, the tensile yield strength R.sub.p0.2, elongation at
fracture A, are determined by a tensile test according to the
standard EN 10002-1. The term "extruded product" includes so-called
"drawn" products obtained by extrusion, followed by drawing.
[0055] During preparatory studies, the present inventors arrived at
a conclusion that a novel material exhibiting a significantly
improved compromise between mechanical strength and formability
should preferably possess a sufficiently high zinc content,
typically above 8.3%, and advantageously above 9.0%. According to
the present invention, the inventors have found a very specific
domain of composition which presents formation of wrought products,
and especially extruded products, which at the same time have, high
static mechanical properties, sufficient resistance to corrosion,
and good formability.
[0056] In connection with the present invention, extruded products
have been developed which can advantageously be used, for example,
as fuselage stringers in aircraft, particularly civilian aircraft.
For this use, damage tolerance is generally not a limiting factor.
It is therefore feasible to optimise UTS and TYS properties while
sacrificing some damage tolerance, so long as corrosion resistance
is not affected to any measurable extent. However, pushing the UTS
and TYS to maximum values usually leads to a decrease in
formability. It is known to one skilled in the art that fuselage
stringers can be often subjected to complex and very peculiar
shaping operations, usually utilizing a technique called
"joggling." When developing an alloy for fuselage stringers that
has high mechanical strength, it is therefore advantageous that the
formability be at least as good as, or preferably better than, the
mechanical strengths of conventional alloys.
[0057] According to one embodiment of the present invention, this
task can be solved inter alia by fine adjustment of the content of
the elements of the alloys and certain impurities, as well as by
optionally adding a controlled concentration of certain other
elements to the alloy composition.
[0058] The present invention includes Al--Zn--Mg--Cu alloys
comprising: [0059] Zn 8.3-14.0 Cu 0.3-2.0 Mg 0.5-4.5 as well as
certain other elements specified hereinbelow, and the rest being
aluminum with its inevitable impurities.
[0060] Alloys according to some embodiments of the present
invention should preferably include at least 0.5% magnesium, since
it may be not possible to obtain satisfactory static mechanical
characteristics with a lower magnesium content. In alloys with a
zinc content of less than 8.3%, the inventors have not found much
improvement with respect to conventional alloys. Preferably, the
zinc content is higher than 9.0%, and still more preferably higher
than 9.5%. However, specific arithmetic relationships between
certain alloying elements should generally be respected in some
embodiments, as will be explained below. In another advantageous
embodiment of the invention, the zinc content is comprised between
9.0 and 11.0%. In any case, the zinc content should preferably not
exceed a value of about 14%, because beyond about 14%, irrespective
of the magnesium and copper content, the results may be less than
satisfactory in some instances.
[0061] The preferable addition of at least 0.3% of copper serves to
improve resistance to corrosion. A minimum copper content of about
0.6% is generally preferred. To ensure satisfactory solution heat
treatment, the Cu content should preferably not exceed about 2%,
and the Mg content should preferably not exceed about 4.5%. A
maximum content of about 3.6% is preferred for magnesium. In a
preferred embodiment, the copper content is between 0.6% and 1.2%,
while the magnesium content is between 2.5% and 3.4%. In another
preferred embodiment, the copper content is between 0.8% and 1.2%,
while the magnesium content is between 2.2% and 3.0%. As will be
explained below, the ratio between the magnesium and copper content
should advantageously conform to certain criteria.
[0062] The present inventors have found that to address certain
problems in the art regarding Al--Zn--Mg--Cu alloys, several
additional technical features can be considered if desired.
[0063] First of all, the alloy should typically be sufficiently
loaded with alloying elements likely to precipitate during
maturation or during on annealing treatment, in order for the alloy
to be capable of presenting advantageous static mechanical
characteristics. As such, in addition to the minimum and maximum
limits for the zinc, magnesium and copper contents indicated
hereinabove, the content of these alloy additions should
advantageously satisfy the condition Mg+Cu>6.4-0.4 Zn in some
embodiments. This was a finding that was completely unexpected
based on the teachings of the prior art.
[0064] To reinforce the effect achieved using the disclosed
preferred alloy composition(s), disclosed above, a sufficient
content of so-called anti-recrystallising elements can also
advantageously be added. More precisely, for alloys with more than
9.5% zinc, at least one element selected from the group consisting
of Zr, Sc, Hf, La, Ti, Y, Ce, Nd, Eu, Gd, Th, Dy, Ho, Er, Yb, Cr
and Mn can preferably be added. And each of these elements, if
added, should preferably be present in a concentration of between
0.02 and 0.7%. It is preferred that the total concentration of the
elements of this group not exceed 1.5%, based on the total weight
of the alloy.
[0065] These presence of one or more anti-recrystallising elements,
in the form of fine precipitates formed during thermal or
thermomechanical treatment, serve to block or at least minimize
recrystallisation. However, it has unexpectedly been found that
when the alloy is highly charged with zinc (Zn>9.5%) excessive
precipitation should be avoided when a wrought product is being
quenched. A compromise then was found for the anti-recrystallising
elements content by the present inventors. Namely, according to one
embodiment of the invention, for alloys with a zinc content of
between 8.3% and 9.5%, zirconium between 0.03% and 0.15% should
advantageously be added, preferably along with at least one element
selected from the group consisting of Sc, Hf, La, Ti, Y, Ce, Nd,
Eu, Gd, Tb, Dy, Ho, Er and Yb. Each element present in this group,
is preferably present in a concentration of between 0.02 and 0.7%.
In a preferred embodiment, Ti is present, alone or together with
one or more other elements from the above group.
[0066] For anti-recrystallising elements, it is advantageous in
some embodiments, irrespective of the zinc content, for such
elements not to exceed the following maximum contents: Cr 0.40; Mn
0.60; Sc 0.50; Zr 0.15; Hf 0.60; Ti 0.15; Ce 0.35 and preferably
0.30; Nd 0.35 and preferably 0.30; Eu 0.35 and preferably 0.30; Gd
0.35; Tb 0.35; Ho 0.40; Dy 0.40; Er 0.40; Yb 0.40; Y 0.20; La 0.35
and preferably 0.30. It is further preferred that the total
concentration of the elements of this group not exceed about
1.5%.
[0067] The inventors have unexpectedly found that in order to
optimize the UTS and TYS, preferably a ratio Mg/Cu>2.4 should be
adopted, more preferably a ratio of at least 2.8, and still more
preferably, 3.5 or even 4.0.
[0068] Another technical feature is associated with the need to be
able to manufacture wrought products industrially under conditions
of very high or even the highest reliability compatible with the
severe requirements of the aeronautical industry, as well as under
satisfactory economic conditions. So it is highly advantageous to
choose a chemical composition which minimises the appearance of hot
cracks or splits during solidification of the plates or billets.
Hot cracks or splits are crippling defaults leading to the plates
or billets that are discarded. It has been noted during numerous
tests that the appearance of hot cracks or splits was unexpectedly
much more probable when the 7xxx alloys finished solidifying below
470.degree. C. To significantly reduce the probability of hot
cracks or splits during casting to an acceptable industrial level,
it was determined according to the present invention that a
chemical composition such as Mg>1.95+0.5 (Cu-2.3)+0.16
(Zn-6)+1.9 (Si-0.04) may be advantageous in some instances.
[0069] Within the scope of the present invention, this empirical
criterion is called the "castability criterion." The alloys
produced according to this variant of the invention typically
complete their solidification at a temperature of between about
473.degree. C. and 478.degree. C., and thus allow an industrial
reliability of casting processes (that is, a constant and excellent
quality of the cast ingots) to be reached that is generally
compatible with some if not all of the severe requirements of the
aeronautical industry.
[0070] Another technical feature of one embodiment of the invention
is associated with a need to substantially minimise to the extent
possible the quantity of insoluble precipitates following
homogenisation and aging treatments. This is because the presence
of such insoluble precipitates decreases the fracture toughness,
the elongation at rupture, and especially the formability of
certain products. Thus, it may be advantageous to employ, a Mg, Cu
and Zn content such as Mg+Cu<7.9-0.4 Zn. The inventors have
found that there likely are little or no disadvantages associated
with selecting a composition close to the borderline represented by
the relationship Mg+Cu<7.9-0.4 Zn. However, beyond this
borderline, the excellent aptitude to deep formability by joggling
(which is one of the main advantages of the present invention) may
decrease rapidly.
[0071] And finally, the inventors have noted that incorporating a
small quantity, of between 0.02 and 0.15% per element, of one or
more elements chosen from the group consisting of Sn, Cd, Ag, Ge
and In, improves the response of the alloy to an annealing
treatment, and also has beneficial effects in terms of mechanical
resistance and resistance to corrosion of products made from such
alloys. A concentration between 0.05% and 0.10% is preferred. Among
these elements, silver is advantageous in some embodiments.
[0072] For profiles, adding one or more anti-recrystallization
elements such as scandium may be particularly advantageous., Such
an effect is also seen, for example, in alloys that may be used to
prepare thick plates. An increase of mechanical strength in
profiles is observed which is higher when the width or thickness of
the section is low. This is known as "press effect" to one skilled
in the art. The inventors have found that when scandium is employed
as an anti-recrystallization element, a concentration between 0.02%
and 0.50% is preferred.
[0073] The present invention is especially advantageous for use in
extruded products. Extruded products can be used advantageously to
produce structural members suitable for use in aeronautical
construction. Preferred applications for products according to the
present invention include their use as structural members for a
fuselage of civilian or other aircrafts or related products. Such
structural members, in particular stringers, are principally
dimensioned for mechanical strength, whereas damage tolerance,
(provided that is has an acceptable minimum value), is normally not
a property used for dimensioning such structural members.
Therefore, if needed and up to a certain point in terms of
structural members, it is often desirable to optimise mechanical
strength of such materials while accepting a certain loss of damage
tolerance, without decreasing the usefulness of the final product.
However, corrosion resistance should always be maintained at an
acceptable minimum level. The increase of mechanical strength of
fuselage stringers allows, at the discretion of the manufacturer
operator, one to reduce the weight of a fuselage being manufactured
while maintaining the same strength and/or permits the formation a
stronger fuselage structure at the same weight as compared to
alloys now utilized by aircraft manufacturers. By increasing the
distance between two adjacent stringers (within the limits of
bending of the fuselage skin sheet), the number of fuselage
stringers can be decreased. This leads to a corresponding decrease
in the number of assembly points between a stringer and a fuselage
skin. Such a decrease in number of assembly points can be very
advantageous, because the assembly points, such as bolts or rivets,
are very significant contributors to the overall manufacturing cost
of such a structure.
[0074] In preferred embodiments of the present invention, it is
possible to obtain a fuselage skin and stringer assembly that is
novel over those of the prior art in terms of the location and
number of assembly points. Thus, a particularly advantageous use of
a product according to the present invention is the use of an alloy
described herein as a structural member in aeronautical
constructions. In particular such alloys are suitable, inter alia,
in the construction of aircrafts comprising a fuselage assembled
from a plurality of stringers and a plurality of sheets, wherein at
least part of the stringers comprise products according to the
present invention. Such an aircraft will generally either (i) have
a structure that is a reduced weight as compared with fuselages
prepared from known materials, which is at least as strong, or (ii)
will have a stronger structure, which will not be much heavier, if
at all, than fuselages of aircrafts made according to the state of
the art.
[0075] It is not only advantageous to minimize the number of
assembly points between structural members of different type (such
as fuselage stringers and fuselage skin), but it is also often
desirable to minimize the number of assembly points between
structural members of the same type, and especially between two
stringers. In order to achieve this goal, it may be advantageous to
utilize sheets or extruded profiles with as large a relevant size
parameter as possible. In the case of extruded profiles, this
relevant size parameter is essentially the length.
[0076] The manufacture of very long profiles in Al--Zn--Mg--Cu
alloys with a high content of alloying elements may in some cases
require fairly detailed process control during casting, extrusion
and thermal treatment, and may require the adjustment of the
chemical alloy composition. In particular, the present inventors
have found that a product according to the present invention can be
obtained by using a reduced extrusion force with respect to known
products having a comparable zinc content. Thus products of the
present invention are generally capable of extruding longer
profiles.
[0077] Airworthiness authorities require that stringers be designed
to resist limit load with large damage. It is generally recommended
that a 2-stringer-bay crack is taken for evaluation of the required
damage tolerance. This is a crack extending over two stringer bays,
with the center stringer broken (see FIG. 10). It was recognized by
the present inventors that the residual strength of fuselage shells
working in tension could benefit from the high strength of
stringers according to the present invention. The use of stringers
according to the present invention as structural members in
aircraft fuselage panels can therefore improve the residual
strength of the structure, inter alia because they close the crack
in the skin, thus preventing unstable fracture. This leads to a
higher residual strength of the panel after damage. This effect can
be used either to increase the safety margin of constructions by
using stringers according to the present invention instead of
stringers of the prior art, and/or the weight of the construction
can be reduced by using reduced size and/or weight stringer
sections and thinner skin panels, and/or by increasing the spacing
between adjacent stringers.
[0078] Airworthiness authorities also generally require that such
structures be designed to resist ultimate load for 3 seconds
without excessive deformation. However, yielding is permitted. This
usually leads to post-buckling designs for fuselage shells in
stability critical locations. Although buckling of perfect columns
(Euler theory) or real-life structure that is very slender is
essentially an elastic phenomenon (governed by Young's modulus),
post-buckling designs display plastic deformation and can therefore
benefit from an increase in yield strength. The buckling test is
shown on FIG. 11.
[0079] It was recognized by the present inventors that the shear-
and compression stability of fuselage shells working in compression
and/or shear could benefit from the high strength of stringers
according to the present invention. The use of stringers according
to the present invention as structural members in aircraft fuselage
panels can improve the shear- and compression stability of fuselage
cells, because these stringers exhibit a higher local buckling
stability. This effect can be used either to increase the safety
margin of constructions in which stringers according to prior art
are substituted by stringers according to the invention, and/or to
decrease the weight of the construction, by using reduced stringer
sections and thinner skin panels, and/or increased stringer
spacing. Alternatively, increased rivet pitch can be obtained,
leading to a lower assembly cost.
[0080] Another problem that arises when using extruded profiles in
Al--Zn--Mg--Cu alloys as fuselage stringers is their formability.
One technique which is commonly used during the industrial
manufacture of fuselage stringers from profiles is joggling.
Joggling involves the introduction of a step localised over a zone
of a few millimetres (see FIG. 6). This can be achieved, for
certain products according to the present invention, either at
elevated temperature (preferably at about 130.degree. C.), and/or
at room temperature. When joggling is performed at room
temperature, it is preferable to employ a solution heat treatment
of the profile delivered in the instable W temper, followed by
quenching and joggling. Joggling at room temperature generally does
not allow forming to be as deep as joggling at elevated
temperature, but when applicable, it is often more practical.
[0081] Joggling as an industrial process typically does not readily
lend itself to the characterisation of materials under development.
However it is known that defects appearing during joggling are
related to the maximum plane deformation which can be supported by
the material. Thus, it is possible to evaluate the aptitude of the
material to joggling using a 3 point bending test such as DIN
standard 50111. According to DIN standard 50111 (September 1987,
see especially section 3.1), the specimen should have a width which
is sufficient with respect to its thickness, in order to be in the
centre of the test piece under conditions of plane deformation.
[0082] The formability of products of the present invention are
evaluated at 130.degree. C. (i.e. warm formability of the product
in its temper of use), wherein a flat test piece is deformed in a
furnace at 130.degree. C. until a drop of the applied force is
detected. This drop indicates that a crack has initiated. The
temperature should be precisely controlled during this test. Since
deformation takes place at elevated temperature, the rate of
deformation is a parameter which influences the result. In the
present case, this rate was fixed at 50 mm/mn. The higher the
bending angle as described in FIG. 8, the better the aptitude of
the material to deformation by joggling. For mechanical reasons, it
is highly recommended that all test pieces to be compared have the
same thickness. Therefore when comparing two test pieces of
different thicknesses, the face which will be under compressive
stress should be machined down to the thickness of the thinnest
test piece. In the case of a profile, the test piece is cut at a
representative location, as shown, for example, in FIG. 7 for the
profile T1.
[0083] A 3 point bending test at 130.degree. C. can be applied to
test pieces cut from products in T6x or T7x temper. It is also
possible to characterize formability in an as-quenched condition W,
if the period of time between quenching and bending testing is
controlled. In the case of extruded products, the bending angle at
130.degree. C. can be expressed as an average value computed from
individual measurements on test pieces taken from different
locations over the length of the profile.
[0084] A product according to the present invention which is
particularly preferred is an extruded product that exhibits a
bending angle in the T6511 temper, determined at 130.degree. C. by
a 3 point bending test according to DIN 50 111 (section 3.1) on a
sample of 1.6 mm thickness cut from a plane area, of at least
34.degree., and a TYS of at least 720 MPa. Even more preferably
extended products of the present invention possess a bending angle
of at least 35.degree. and a TYS of at least 750 MPa. For a
thickness up to 60 mm, the static mechanical properties (UTS, TYS
and A) do not depend much on the thickness of the section.
[0085] Another preferred product is an extruded product that
exhibits a bending angle in the T76511 temper, determined at
130.degree. C. by the 3 point bending test according to DIN 50 111
(section 3.1) on a sample of 1.6 mm thickness cut from a plane
area, of at least 36.degree., and a TYS of at least 660 MPa, and
more preferably of at least 670 MPa. Such a product can be used,
for example, for applications in which a corrosion resistance
rating of at least EB (EXCO test according to ASTM G34) is required
for non machined specimens.
[0086] Extruded products of the present invention including the two
particular ones identified above, can be used advantageously as
fuselage stringers in many applications and are particularly useful
for aircrafts including civilian aircrafts.
[0087] And mentioned above, the present inventors have surprisingly
found that compared to prior art products, and including prior art
products with a comparable zinc content, products according to
certain embodiments of the present invention exhibit a high warm
formability. On the other hand, cold formability in the unstable W
temper after solution heat treatment and quenching may be slightly
less. For the manufacture of structural members of aircrafts, such
as fuselage stringers, warm forming process is preferable,
particularly if deformation during joggling is deep.
[0088] Products according to the present invention can also be
used, for example, as floorbeams for aircrafts. In the form of
extruded profiles, the can also be used, for example, as seat
tracks. In civilian aircrafts, seat tracks are profiles, generally
of great length, which are normally parallel to the length of the
cabin, and on which the seats are mounted. According to the present
invention, seat tracks in T76511 temper can be obtained which
exhibit a UTS in the area where the seats are fixed (i.e. the top
of a "I" shape) of 670 MPa or more, and even of 680 MPa or more,
and a TYS of 640 MPa or more, and oven of 660 MPa or more. Seat
tracks of commercial aircrafts have to be resistant to corrosion by
corrosive liquid foodstuff under high mechanical stress. Indeed,
seat tracks according to the present invention exhibit a good
stress corrosion resistance as determined according to ASTM
G47.
[0089] The use of structural members according to the present
invention for aircraft construction leads to significant weight
savings, which allows to increase the load capacity of said
aircraft, or to decrease fuel consumption.
[0090] The following examples illustrate different embodiments of
the invention and demonstrate its advantages; they do not restrict
this invention.
EXAMPLES
Example 1
[0091] Several Al--Zn--Mg--Cu alloys were prepared by
semi-continuous casting of rolling ingots, and were then subjected
to a range of conventional transformation techniques, comprising a
homogenisation stage, the parameters of which have been determined
according to U.S. Pat. No. 5,560,789, the content of which is
incorporated herein by reference in its entirety. The
homogenisation stage was followed by hot rolling, solution heat
treatment which was followed by quenching and stress relieving
operations, and finally an aging treatment was conducted in order
to obtain a product in temper T651. This process resulted in the
formation of plates in the T651 temper having a thickness of 20
mm.
[0092] Compositions of plates according to this example are
specified in Table 1 below: TABLE-US-00001 TABLE 1 Mg/ Alloy Zn Mg
Cu Fe Si Zr Ti Mn Sc Cu A 8.40 2.11 1.83 0.09 0.06 0.11 0.017 0 0
1.15 B 10.27 3.2 0.71 0.08 0.03 0.11 0.017 0 0 4.57 C 10.08 2.69
0.95 0.08 0.03 0.11 0.014 0 0 2.83 D 9.97 2.14 1.32 0.09 0.03 0.11
0.017 0 0 1.62
[0093] The static mechanical characteristics were determined by a
tensile test according to standard EN 10002-1. The fracture
toughness K.sub.1C was determined according to standard ASTM
E399.
[0094] The results are specified in Table 2 TABLE-US-00002 TABLE 2
Tensile test Tensile test Fracture L direction TL direction
toughness L-T R.sub.p0.2 R.sub.m A R.sub.p0.2 R.sub.m A K.sub.Ic
Alloy MPa MPa % MPa MPa % MPa m A 627 665 14.7 566 623 13.6 31.9 B
716 726.5 6.5 640 696 5.2 21.1 C 700 717 9.2 629 676 8.1 21 D 665
685 12.2 608 649 11 26.8
[0095] It appears that plate C according to the invention presents
a good compromise between mechanical strength and elongation.
Compared to plate D, the mechanical strength is significantly
improved. Compared to plate A in alloy 7449 according to prior art,
plate C exhibits a mechanical strength that is very significantly
improved. The fact that fracture toughness is less good in plate C
than in plate B could potentially limit the possibilities of
application of plate C to those applications for which fracture
toughness is not taken into account when dimensioning the
structural members, but which require both a high mechanical
strength and a good formability. Compared to plate B, the
elongation at fracture of plate C is significantly improved.
[0096] Moreover, in order to obtain the properties given in Table
2, plate B needs to be submitted to a rather long solution heat
treatment, which does not lend itself to the requirements of
industrial production. And yet, coarse phases have been found in
the product which has an adverse effect on the homogeneity of
mechanical properties, both within the same production lot and
within the same individual product (plate or profile). Such
presence of precipitates may preclude the use of product B as a
structural member in aircrafts.
Example 2
[0097] Several rolling ingots whose chemical composition is
specified in Table 3 were cast. The silicon content was
approximately the same for all of them, about 0.04%.
[0098] Alloys G1, G2, G3 and G4, as well as alloys B and D,
described in example 1 are used as comparisions with certain
preferred embodiments. Alloy C is an alloy according to the
invention described in example 1. During testing, all these alloys
exhibited satisfactory castability, that is, no splits or cracks
were observed during casting tests performed on an industrial
scale.
[0099] Alloys G5, G6, G7, G8 were used as comparisons with certain
preferred embodiments of the present invention, and alloy G9 is an
alloy 7060 as per the prior art; these alloys exhibited cracks
during casting tests.
[0100] Difficulties showing up during casting of these alloys do
not necessarily render the wrought products from these plates
unsuitable for use, but they are the cause of extra cost because
their implementation (that is, the quantity of vendible metal
relative to the quantity of charged metal, a parameter directly
associated with the quantity of discarded ingots) will be greater
than for the alloys corresponding to certain preferred domains of
the invention. In addition, the propensity of these alloys to form
splits during their solidification makes it very difficult to
render the casting process reliable within the scope of a quality
assurance program based on statistical process control.
[0101] It is noted that all the 7xxx alloys having a very
pronounced propensity to form splits or cracks in casting generally
have a magnesium content lower than desired magnesium content
typically employed for such alloys. A desirable Mg value was
obtained by calculating the Mg limit value defined by the empirical
castability criterion. TABLE-US-00003 TABLE 3 Zn Mg Cu Observed
Critical Mg > Alloy [%] [%] [%] crackability Mg content Critical
Mg G1 7.5 3 3 low 2.54 yes G2 8.5 3 2.3 low 2.35 yes G3 7.5 3 1.6
low 1.84 yes G4 6.5 3 2.3 low 2.03 yes B 10.27 3.2 0.71 low 1.82
yes C 10.08 2.69 0.95 low 1.91 yes D 9.97 2.14 1.32 low 2.08 yes G5
8.5 2.3 3 high 2.7 no G6 6.5 2.3 3 high 2.38 no G7 8.5 1.6 2.3 high
2.35 no G8 7.5 1.6 1.6 high 1.84 no G9 7 1.65 2.1 high 2.01 no
Example 3
[0102] Extrusion ingots have been cast from alloys whose
composition is summarized in Table 4. Homogenization was carried
out as follows: TABLE-US-00004 Ingots Q1 and Q2: 4 h at 465.degree.
C. + 20 h at 476.degree. C. Ingots Q3 and Q4: 4 h at 465.degree. C.
+ 20 h at 471.degree. C. Ingots P1 through P3: 20 h at 471.degree.
C.
M, T and S phases were completely dissolved by these homogenization
treatments; this was checked by differential enthalpic analysis
(according U.S. Pat. No. 5,560,789, incorporated herein by
reference).
[0103] Ingot diameter was 200 mm for ingots P3 and Q1 through Q4,
and 155 mm for ingots P1 and P2. TABLE-US-00005 TABLE 4 Ingot Zn Mg
Cu Cr Mn Si Fe Zr Ti Mg/Cu P1 8.10 2.48 1.65 0.14 0.17 0.01 0.08
0.15 0.03 1.50 P2 8.45 2.60 1.76 0.18 0.18 0.05 0.14 0.12 0.02 1.48
P3 8.39 2.55 1.71 0.18 0.16 0.04 0.15 0.11 0.02 1.49 Q1 10.20 3.10
0.68 0.17 0.17 0.07 0.08 0.13 0.04 4.56 Q2 10.20 2.84 0.95 0.18
0.17 0.06 0.11 0.13 0.03 2.99 Q3 9.98 2.10 1.24 0.18 0.17 0.06 0.14
0.12 0.03 1.69 Q4 10.00 2.15 1.25 0.18 0.17 0.07 0.14 0.12 0.03
1.72 R1 10.18 2.97 0.66 0.17 0.16 0.07 0.13 0.11 0.02 4.5 R2 10.16
3.12 0.70 0.17 0.16 0.07 0.13 0.11 0.02 4.46
[0104] From these homogenized and scalped ingots, five types of
profiles T1, T2, T3, T4 and T5 were extruded. Their sections are
represented on FIGS. 1, 2, 3, 4 and 5. The temperature of the
container and of the die was above 400.degree. C. , and the
extrusion speed was below 0.50 m/min.
[0105] Maximum extrusion forces are summarized in Table 5. It can
be seen that surprisingly the alloys according to the present
invention do not require a higher extrusion force, and that the
extrusion force surprisingly even decreases for certain types of
profiles with increasing magnesium content. TABLE-US-00006 TABLE 5
Extrusion force Extrusion force Extrusion force Extrusion force
Extrusion force Extrusion force [bars] for [bars] for [bars] for
[bars] for [bars] for Extrusion Profile ingot P1 ingot Q1 ingot Q2
ingot Q3 ingot Q4 ratio T1 179 175 170 164 164 58 T2 151 145 142
137 139 24 T3 203 208 200 193 195 13
[0106] Profiles Q1 through Q4 were solution heat treated at
471.degree. C., while profiles P1 through P3 were solution heat
treated at 472.degree. C. (profiles T1, T2 and T3). All profiles
were water quenched and then stretched with a permanent set
comprised between 1.5% and 2%. Products in tempers T6511 or T76511
were obtained. Their mechanical properties are summarized in Table
6 for specimens of three different thickness values in temper
T6511, cut from a flat area of the profile. This temper has been
obtained by artificial aging under the following conditions: [0107]
Alloys Q1 and Q2: 18 h at 120.degree. C.
[0108] Alloys P1, P2, P3, Q3 and Q4: 36 h at 120.degree. C.
TABLE-US-00007 TABLE 6 Alloy R.sub.m[MPa] R.sub.p0.2 [MPa] A [%]
profile T1 T3 T2 T1 T3 T2 T1 T3 T2 Q1 755 753 788 743 736 783 8.4
7.0 4.7 Q2 746 750 778 731 729 771 9.8 8.7 6.0 Q3 698 699 728 674
673 712 13.6 12.3 9.3 Q4 697 696 723 673 670 704 13.3 11.7 10.7 P1
708 694 745 671 656 718 12.5 11.7 7.7 Sampling: T1 = foot of
profile T1. T2 = top of profile T2. T3 = top of profile T3.
[0109] Mechanical properties in T76511, obtained by articifial
aging under the following conditions: TABLE-US-00008 Q1 through Q4:
12 h at 120.degree. C. + 8 h at 150.degree. C. P1: 12 h at
120.degree. C. + 10 h at 156.degree. C.
[0110] are summarized in Table 7. TABLE-US-00009 TABLE 7 Alloy
R.sub.m [MPa] R.sub.p0.2 [MPa] A [%] profile T1 T3 T2 T1 T3 T2 T1
T3 T2 Q1 694 706 712 674 687 696 99 7.7 8.3 Q2 694 704 708 675 686
693 10.3 9.0 8.3 Q3 674 676 697 662 664 684 9.6 9.7 10.0 Q4 673 677
687 657 663 672 11.1 9.7 10.0 P1 659 644 686 615 589 643 12.1 10.3
9.1 Sampling: T1 = foot of profile T1. T2 = top of profile T2. T3 =
top of profile T3.
[0111] It can be seem that compared to alloy P1, alloys Q1 and Q2
have a mechanical strength significantly higher.
[0112] Corrosion resistance was evaluated by means of the EXCO test
(ASTM G34) of the products Q1 and Q2 in temper T6511 (unmachined
specimens taken from the beginning of the extrusion) as EA or EB,
and was generally at least as good as or better than what was
observed for samples P1, P2, P3, Q3 and Q4.
[0113] For R1 and R2, the following mechanical properties were
found: TABLE-US-00010 TABLE 8 T6511 T76511 R.sub.m R.sub.p0.2 A
R.sub.m R.sub.p0.2 A [MPa] [MPa] [%] [MPa] [MPa] [%] Alloy R1, 753
738 8 688 669 10 profile T4(a) Alloy R1, 756 743 8 686 667 9
profile T4(b) Alloy R2, 755 743 7 676 659 10 profile T5 NOTE:
profile T4(a) = samples cut from the top of the profile, see FIG.
4, feature (a).
Example 4
[0114] Formability of profiles of shape T1 according to example 3
was studied by using the three point bending test according to DIN
50 111 (September 1987, section 3.1). The location of sampling, a
flat zone, is shown on FIG. 7. The important parameters of the
three point bending test are shown on FIG. 9. The test was
performed at 130.degree. C.
[0115] Both tempers T6511 and T76511 were tested. The resulting
values for the bending angle .alpha. (as defined on FIG. 8) are
summarized in Table 9. These are mean values computed from half a
dozen individual determinations using specimens cut at different
locations distributed over the length of the profiles.
TABLE-US-00011 TABLE 9 Bending angle alloy temper T76511 temper
T6511 Q1 43.4.degree. -- Q2 38.1.degree. 36.9.degree. Q3
33.9.degree. 33.8.degree. P1 41.5.degree. 35.2.degree.
[0116] In all cases, the profiles according to the invention (Q1
and Q2) had a formability which was comparable to the formability
of profiles according to prior art (Q3 and P1).
Example 5
[0117] Cold forming of samples similar to those used in the example
4 (in the unstable W temper after solution heat treatment and
quenching) was studied at room temperature by using the same three
point bending test. The variation of the bending angle .alpha. (as
defined on FIG. 8) over the length of the profiles is small. Table
10 refers to values obtained in the W temper. TABLE-US-00012 TABLE
10 Sample Bending angle Q1 27.1.degree. Q2 27.3.degree. Q3
33.6.degree. P1 34.5.degree.
Example 6
[0118] Rolling ingots were elaborated by a process similar to the
one described in example 1. The chemical composition is given in
Table 11. Hot rolled plates with a thickness of 25 mm were obtained
by a process similar to the one described in example 1. The plates
were solution heat treated at a temperature between 472 and
480.degree. C. for 2 hours, quenched and stretched with a permanent
set comprised between 1.5% and 2%. Finally, the stretched plated
were artificially aged at a temperature of 135.degree. C.
TABLE-US-00013 TABLE 11 Al- Mg/ loy Zn Mg Cu Fe Si Zr Ti Mn Sc Cu M
9.94 3.02 0.78 0.04 0.03 0.10 0.063 0 0 3.87 N 10.00 2.72 0.77 0.06
0.04 0.10 0.055 0 0.10 3.53 K 9.90 2.03 1.55 0.03 0.03 0.10 0.05 0
0.10 1.31
[0119] The following mechanical properties were obtained:
TABLE-US-00014 TABLE 12 R.sub.p0.2(L) Duration R.sub.p0.2(L)
(compres- of aging R.sub.m(L) (tensile) sive) K.sub.Ic (or K.sub.q)
Plate [h] [MPa] [MPa] [MPa] A [%] [MPa m].sup.1 N 6 711 687 678
10.4 16.9 N 12 702 695 696 9.7 14.5 M 6 691 676 662 10.0 21.2 M 12
684 675 660 8.9 20.4 K 6 694 666 620 12.9 23.2 K 12 692 674 685
11.7 19.7 1: measured with B = 1 inch and W = 3 inches.
[0120] It was checked that for plates N and K, the aging treatment
of 12 h leads to the T6 temper. For aging times significantly
longer, R.sub.p0,2(L) and R.sub.m(L) decrease.
[0121] It can be seen that for the same Zn content and with a
similar Mg/Cu ratio, plate N (containing 0.10% scandium) exhibits
better static mechanical properties than plate M (no scandium).
[0122] For the same zinc content and for the same scandium content,
plate N (high Mg/Cu ratio) exhibts better R.sub.p0,2(L) and
R.sub.m(L) values than plate K.
Example 7
[0123] For several profiles elaborated according to example 3, the
resistance to stress corrosion was evaluated. The results are
summarized in Table 13. TABLE-US-00015 TABLE 13 Stress Duration of
Sample Temper [MPa] the test Alloy Q1, profile T1, L direction
T76511 530 >30 days Alloy Q1, profile T1, L direction T6511 350
>30 days Alloy P1, profil T4, L direction T76511 430 >30 days
Alloy P1, profile T4, LT direction T76511 400 >30 days Alloy P1,
profile T4, LT direction T6511 280 >30 days Alloy R1, profile
T4, LT direction T76511 Alloy R1, profile T4, LT direction
T76511
[0124] It can be seen that the products according to the invention
show a satisfactory resistance to stress corrosion.
Example 8
[0125] Profiles in alloys 7349 or 7449 were produced with and
without scandium, according to a process similar to the one
described in example 3. Table 14 lists the chemical compositions,
Table 15 the obtained mechanical properties. TABLE-US-00016 TABLE
14 Alloy Zn Mg Cu Fe Si Zr Ti Mn Cr Sc Mg/Cu X1 8.1 2.5 1.7 0.08
0.01 0.15 0.03 0.17 0.14 0 1.47 X2 8.4 2.1 1.9 0.06 0.02 0.10 0.02
0 0 0 1.11
[0126] TABLE-US-00017 TABLE 15 Temper T6 Temper T76 R.sub.m
R.sub.p0.2 A R.sub.m R.sub.p0.2 A product [MPa] [MPa] [%] [MPa]
[MPa] [%] Alloy X1, 713 681 15 650 606 13 profile T1, measured at
(a) Alloy X1, 711 678 11 654 614 10 profile T2, measured at (a)
Alloy X1, 740 708 7 670 628 8.5 profile T2, measured at (b) Alloy
X2, 673 645 17 645 626 14 profile T1, measured at (a) Alloy X2, 680
653 12 646 623 11 profile T2, measured at (a) Alloy X2, 728 699 10
667 632 11 profile T2, measured at (b)
[0127] A comparison with the results of example 3 shows that the
products according to the invention have increased mechanical
strength (R.sub.m, R.sub.p0,2) compared to products X1 and X2
according to prior art.
Example 9
[0128] Seat tracks for aircraft have been manufactured from
extrusion billets of chemical composition R1 and Q1 according to
the previous examples. These profiles are "I type" profiles
including a foot section, a centre section, and a top section on
which the seats are fixed. The thickness of the centre section was
of the order of 2 mm, and the height of the profile was of the
order of 65 mm.
[0129] Table 16 summarizes the static mechanical properties in
temper T76511. TABLE-US-00018 TABLE 16 Alloy Sampling R.sub.m [MPa]
R.sub.p0.2 [MPa] R1 Foot 688 669 R1 Top 686 667 Q1 Foot 672 643 Q1
Top 683 660
[0130] Stress corrosion testing according to ASTM G47 shows good
resistance to stress corrosion.
Example 10
[0131] Numerical models for damage tolerance of fuselage shells
employing high-strength stringers according to the invention were
evaluated in order to determine the residual strength of a fuselage
shells. Airworthiness authorities require that such structure be
designed to resist limit load with large damage; it is recommended
that a 2-stringer-bay crack is taken for evaluation of the required
damage tolerance. This is a crack 12 extending over two stringer
bays 14, 16, with the center stringer 18 broken (see FIG. 10). It
was recognized by the present inventors that the residual strength
of fuselage shells working in tension could benefit from the high
strength of stringers according to the present invention. The use
of stringers according to the present invention as structural
members in aircraft fuselage panels can improve the residual
strength of the structure, because they close the crack 12 in the
skin 20, thus preventing unstable fracture. This leads to a higher
residual strength of the panel after damage. This effect can be
used either to increase the safety margin of constructions in which
stringers according to prior art are substituted by stringers
according to the invention, or to decrease the weight of the
construction, by using reduced stringer sections and thinner skin
panels, and/or increased stringer spacing.
[0132] The fracture of fuselage skin is governed by the stress
intensity factors (SIF) at the crack tips. For a typical fuselage
crown structure with a stringer pitch of 200 mm and a stiffening
ratio (section of stringer/total section) of 0.25, the SIF for a
crack of 2-stringer-bay length in a panel with stringers made of
the present invention shows a reduction of about 5% compared to a
panel with stringers in the widely used 2024 T3 alloy. For longer
cracks, 2024 stringers are more often solicited in the plastic
domain, and the stress in the stringers will not reach the yield
point. In the case of plastic domain 2024 stringers, the SIF of the
present invention is reduced by about 15%. However, it should be
noted that for stringers according to prior art in alloy 2024,
there is also a risk of the stringers reaching their ultimate
tensile stress and failing, whereas stringers according to the
invention will not break under these conditions.
Example 11
[0133] Numerical models of fuselage shells working in compression
and/or shear were evaluated in order to determine the shear- and
compression stability. Airworthiness authorities require that such
structure be designed to resist ultimate load for 3 seconds without
excessive deformation. However, yielding is permitted. This usually
leads to post-buckling designs for fuselage shells in stability
critical locations. Although buckling of perfect columns (Euler
theory) or real-life structure that is very slender is essentially
an elastic phenomenon (governed by Young's modulus), post-buckling
designs display plastic deformation and can therefore benefit from
an increase in yield strength. The buckling test is shown in FIGS.
11a and 11b. FIG. 11b is taken along line A-A rotated 90.degree. A
fuselage skin 20 is shown with two stringers 14, 16 attached
thereto. Rivets 22 attach the stringers 16, 18 and the fuselage
skin 20. The gap 24 between the stringers and skin is clearly shown
in both FIG. 11a and rotated FIG. 11b.
[0134] It was recognized by the present inventors that the shear-
and compression stability of fuselage shells working in compression
and/or shear could benefit from the high strength of stringers
according to the present invention. The use of stringers according
to the present invention as structural members in aircraft fuselage
panels can improve the shear- and compression stability of fuselage
cells, because these stringers exhibit a higher local buckling
stability. This effect can be used either to increase the safety
margin of constructions in which stringers according to prior art
are substituted by stringers according to the invention, or to
decrease the weight of the construction, by using reduced stringer
sections and thinner skin panels, and/or increased stringer
spacing. Alternatively, increased rivet pitch can be obtained,
leading to a lower assembly cost.
[0135] The gain in buckling stability can be obtained by applying a
very general method given in Michael C. Y. Niu, Airframe Stress
Analysis and Sizing, 2.sup.nd edition, chapter 10, incorporated
herein by reference in its entirety. The present inventors have
found that stringer stability of the stringer according to the
invention (with compressive yield strength of 700 MPa and
compressive elastic modulus of 73 GPa compared to the widely used
7150 T77511 stringer (with a typical compressive yield strength of
538 MPa and an elastic modulus of 73 GPa) is increased on the order
of about 15%, for typical fuselage Z-stringers based on the data in
Table 17 and FIG. 12.
[0136] Table 17 shows the parameters of different stringer
geometries analyzed. FIG. 12 compares crippling stress for these
different stringer geometries Z1 to Z8 (from the left to the
right). TABLE-US-00019 TABLE 17 Small Z-stringer designs: Z1 Z2 Z3
Z4 Z5 Z6 Z7 Z8 Free flange 12.7 12.7 12.7 12.7 12.7 12.7 12.7 12.7
width [mm] Fastened flange 25.4 25.4 25.4 25.4 25.4 25.4 25.4 25.4
width [mm] Height [mm] 38.1 38.1 38.1 38.1 38.1 38.1 38.1 38.1 Free
flange 1.0 1.5 1.5 2.0 1.0 1.5 1.5 1.5 thickness [mm] Fastened
flange 1.0 1.0 1.5 1.5 1.0 1.0 1.0 1.5 thickness [mm] Web thickness
1.0 1.0 1.0 1.0 1.5 1.5 1.5 1.5 [mm] Section [mm.sup.2] 76 83 95
102 95 102 102 114 Equivalent 1.0 1.1 1.3 1.3 1.3 1.3 1.3 1.5
thickness [mm]
[0137] 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.
[0138] The priority document, French Patent Application No.02
04250, filed Apr. 5, 2002 is incorporated herein by reference in
its entirety.
[0139] As used herein and in the following claims, articles such as
"the", "a" and "an" can connote the singular or plural.
[0140] All documents referred to herein are specifically
incorporated herein by reference in their entireties.
* * * * *