U.S. patent application number 09/773270 was filed with the patent office on 2002-10-17 for aluminum alloy having superior strength-toughness combinations in thick gauges.
Invention is credited to Chakrabarti, Dhruba J., Goodman, Jay H., Liu, John, Sawtell, Ralph R., Venema, Gregory B..
Application Number | 20020150498 09/773270 |
Document ID | / |
Family ID | 25097708 |
Filed Date | 2002-10-17 |
United States Patent
Application |
20020150498 |
Kind Code |
A1 |
Chakrabarti, Dhruba J. ; et
al. |
October 17, 2002 |
Aluminum alloy having superior strength-toughness combinations in
thick gauges
Abstract
A 7XXX series aluminum alloy having reduced quench sensitivity
suitable for use in aerospace structural components, such as
integral wing spars, ribs, extrusions and forgings comprises, in
weight %: 6 to 10 Zn, 1.3 to 1.9 Mg, 1.4 to 2.2 Cu, wherein
Mg.ltoreq.Cu+0.3, one or more of 0 to 0.4 Zr, up to 0.4 Sc, up to
0.2 Hf, up to 0.4 Cr, up to 1.0 Mn and the balance Al plus
incidental additions including Si, Fe, Ti and the like plus
impurities. By controlling the Mg content to 1.3 to 1.7 wt. %,
limiting Mg.ltoreq.Cu+0.3 and 6.5.ltoreq.Zn.ltoreq.8.5, the alloy
provides significantly improved combined strength and fracture
toughness in heavy gauges. For example, in a six-inch thick plate
there is provided a combination of about 75 ksi quarter-plane
tensile yield strength (L) with a fracture toughness (L-T) of about
33 ksi{square root}in which progresses by artificial
aging/tempering to a combined strength and fracture toughness of
about 67 ksi tensile yield strength (L) and a fracture toughness
(L-T) of about 40 ksi{square root}in. The alloy product possesses
equally attractive combinations of strength and fracture toughness
when intentionally quenched slowly following solution heat
treatment so as to lessen dimensional distortion, particularly in
shapes of varying cross section.
Inventors: |
Chakrabarti, Dhruba J.;
(Export, PA) ; Liu, John; (Lower Burrell, PA)
; Goodman, Jay H.; (Murrysville, PA) ; Venema,
Gregory B.; (Bettendorf, IA) ; Sawtell, Ralph R.;
(Brecksville, OH) |
Correspondence
Address: |
Richard L. Byrne
WEBB ZIESENHEIM LOGSDON ORKIN & HANSON, P.C.
700 Koppers Building
436 Seventh Avenue
Pittsburgh
PA
15219-1818
US
|
Family ID: |
25097708 |
Appl. No.: |
09/773270 |
Filed: |
January 31, 2001 |
Current U.S.
Class: |
420/532 ;
148/439; 148/690 |
Current CPC
Class: |
C22C 21/10 20130101;
C22F 1/053 20130101 |
Class at
Publication: |
420/532 ;
148/439; 148/690 |
International
Class: |
C22C 021/16; C22C
021/18; C22F 001/04 |
Claims
We claim:
1. A 7XXX series aluminum alloy possessing superior strength and
fracture toughness properties consisting essentially of, in % by
weight, about: 6 to 10 Zn, 1.3 to 1.9 Mg, 1.4 to 2.2 Cu, 0 to 0.4
Zr 0 to 0.4 Sc 0 to 0.2 Hf 0 to 0.4 Cr 0 to 1.0 Mn, and the balance
being Al and incidental additions and impurities.
2. The aluminum alloy of claim 1 in gauges greater than 1.5
inches.
3. The aluminum alloy of claim 1 in gauges between about 3 to about
8 inches.
4. The aluminum alloy of claim 1 in gauges between about 6 to about
12 inches.
5. The aluminum alloy of claim 1 in gauges of about 12 inches or
greater.
6. The aluminum alloy of claim 1 wherein: Mg.ltoreq.1.7.
7. The aluminum alloy of claim 1 wherein: Mg.ltoreq.Cu+0.3.
8. The aluminum alloy of claim 1 wherein: Mg.ltoreq.1.7, and
Mg.ltoreq.Cu+0.3.
9. The aluminum alloy of claim 1 wherein: Mg.ltoreq.1.7,
Mg.ltoreq.Cu+0.3, and 6.ltoreq.Zn.ltoreq.8.5
10. The aluminum alloy of claim 1 wherein: Mg.ltoreq.Cu.
11. A 7XXX series aluminum alloy possessing superior strength and
fracture toughness properties consisting essentially of, in % by
weight, about: 6 to 8.5 Zn, 1.3 to 1.8 Mg, 1.4to 2.0 Cu, up to
about 0.4 Zr, one or more elements selected from the group
consisting of: up to about 0.4 Sc, up to about 0.2 Hf. up to about
0.4 Cr, up to about 1.0 Mn, and the balance being Al and incidental
additions and impurities.
12. The aluminum alloy of claim 11 in gauges greater than 1.5
inches.
13. The aluminum alloy of claim 11 in gauges between about 3 to
about 8 inches.
14. The aluminum alloy of claim 11 in gauges between about 6 to
about 12 inches.
15. The aluminum alloy of claim 11 in gauges of about 12 inches or
more.
16. A 7XXX series aluminum alloy possessing superior strength and
fracture toughness combination properties consisting essentially
of, in % by weight, about: 6.5 to 8.5 Zn, 1.3 to 1.7 Mg, 1.4 to 2.0
Cu, 0 to 0.4 Zr, 0 to 0.4 Sc, 0 to 0.2 Hf, 0 to 0.4 Cr, 0 to 1 Mn,
and the balance being Al and incidental additions and
impurities.
17. The aluminum alloy of claim 16 in gauges greater than 1.5
inches.
18. The aluminum alloy of claim 16 in gauges between about 2 inches
to about 12 inches.
19. The aluminum alloy of claim 16 in gauges of about 6 inches to
about 12 inches or more.
20. The aluminum alloy of claim 16 having a quarter-plane tensile
yield strength (TYS) in the longitudinal direction (L) greater than
about 69 ksi and a fracture toughness (K.sub.Ic) in the L/-T
direction greater than about 30 to 34 ksi{square root}in wherein
said alloy has been processed in a thickness of about 6 inches and
in a solution heat treated, quenched, stretched and artificially
aged condition.
21. An aluminum alloy structural component made from plate,
extrusion or forging for a commercial jet aircraft, said structural
component made from a 7XXX series aluminum alloy consisting
essentially of, in % by weight, about: 6 to 10 Zn, 1.3 to 1.9 Mg,
1.4 to 2.2 Cu, 0 to 0.4 Zr 0 to 0.4 Sc 0 to 0.2 Hf 0 to 0.4Cr 0 to
1.0 Mn, and the balance being Al and incidental additions and
impurities.
22. The aluminum alloy structural component of claim 21 wherein:
Mg.ltoreq.1.7.
23. The aluminum alloy structural component of claim 21 wherein:
Mg.ltoreq.Cu+0.3.
24. The aluminum alloy structural component of claim 21 wherein:
Mg.ltoreq.1.7, and Mg.ltoreq.Cu+0.3.
25. The aluminum alloy structural component of claim 21 wherein:
Mg.ltoreq.1.7, Mg.ltoreq.Cu+0.3, and 6.ltoreq.Zn.ltoreq.8.5.
26. The aluminum alloy structural component of claim 21 wherein:
Cu.gtoreq.Mg.
27. The aluminum structural component of claim 21 having a
thickness of between about 4 to about 8 inches or more.
28. An aluminum alloy structural component made from plate,
extrusion or forging for a commercial jet aircraft, said structural
component made from a 7XXX series aluminum alloy consisting
essentially of, in % by weight, about: 6 to 8.5 Zn, 1.3 to 1.8 Mg,
1.4 to 2.0 Cu, up to about 0.4 Zr up to about 0.4 Sc up to about
0.2 Hf up to about 0.4 Cr up to about 1.0 Mn, and the balance being
Al and incidental additions and impurities.
29. The aluminum structural component of claim 28 having a
thickness of greater than about 1.5 inches.
30. The aluminum structural component of claim 28 having a
thickness of greater than 6 inches.
31. The aluminum structural component of claim 28 in the form of an
integral or built-up structural component for use in an
aircraft.
32. The aluminum structural component of claim 28 in the form of an
integral spar for use in an aircraft wing structure.
33. The aluminum structural component of claim 28 having a
quarter-plane tensile yield strength, TYS (L), greater than about
69 ksi and a fracture toughness, K.sub.Ic, (L-T), greater than
about 30 to 34 ksi{square root}in wherein said alloy has been
processed in a thickness of about 6 inches and in a solution heat
treated, quenched, stretched and artificially aged condition.
34. An aluminum alloy structural component made from plate,
extrusion or forging for a commercial jet aircraft, said structural
component made from a 7XXX series aluminum alloy consisting
essentially of, in % by weight, about: 6 to 8.5 Zn, 1.3to 1.7 Mg,
1.4 to 2.0 Cu, up to about 0.4 Zr up to about 0.4 Sc up to about
0.2 Hf up to about 0.4 Cr up to about 1.0 Mn, and the balance being
Al and incidental additions and impurities.
35. The aluminum structural component of claim 34 having a
thickness of greater than about 1.5 inches.
36. The aluminum structural component of claim 34 having a
thickness of greater than 6 inches.
37. An aluminum alloy structural component for a commercial jet
aircraft, said structural component made from a 7XXX series
aluminum alloy consisting essentially of, in % by weight, about: 6
to 8.5 Zn, 1.3 to 1.8 Mg, 1.4 to 2.0 Cu, wherein Cu.gtoreq.Mg, and
(Cu+Mg+Zn)<12.5, up to about 0.4 Zr, one or more elements
selected from the group consisting of: up to about 0.4 Sc, up to
about 0.2 Hf, up to about 0.4 Cr, up to about 1.0 Mn, and the
balance being Al and incidental additions and impurities.
38. The aluminum structural component of claim 37 in the form of an
integral spar for use in an aircraft wing structure.
39. The aluminum structural component of claim 37 in the form of an
integral wing panel and stringer for use in an aircraft wing
structure.
40. The aluminum structural component of claim 37 in the form of a
rib for use in an aircraft wing structure.
41. The aluminum alloy structural component of claim 37 in the form
of a stepped extrusion or extruded spar for use in an aircraft
structure.
42. The aluminum alloy structural component of claim 37 in the form
of a forged spar for use in an aircraft wing structure.
43. The aluminum structural component of claim 37 wherein the
component is a wrought product having both thick and thin sections
and is intentionally quenched slowly following a solution heat
treatment to minimize quench distortion.
44. The aluminum structural component of claim 37 having a
quarter-plane tensile yield strength, TYS (L), greater than about
69 ksi and a fracture toughness, K.sub.Ic (L-T), greater than about
30 to 34 ksi{square root}in wherein said alloy has been processed
in a solution heat treated, quenched, stretched and artificially
aged condition.
45. The aluminum structural component of claim 37 wherein the alloy
contains Mg.ltoreq.Cu+0.3.
46. The aluminum structural component of claim 37 wherein the alloy
contains Cu.gtoreq.Mg.
47. An aluminum alloy structural component for a commercial jet
aircraft, said structural component made from a 7XXX series
aluminum alloy consisting essentially of, in % by weight, about:
6.5 to 8.5 Zn, 1.3 to 1.7 Mg, 1.4 to 2.0 Cu, 0 to 0.4 Zr, 0 to
about 0.38 Sc, 0 to about 0.20 Hf, 0 to about 0.37 Cr, 0 to about
1.0 Mn, and the balance being Al and incidental additions and
impurities.
48. The aluminum structural component of claim 47 having a
thickness of greater than about 1.5 inches.
49. The aluminum structural component of claim 47 having a
thickness of between about 3 to about 8 inches or more.
50. The aluminum structural component of claim 47 wherein the
component is a wrought product and is intentionally quenched slowly
following a solution heat treatment.
51. The aluminum structural component of claim 47 in the form of an
integral spar for use in an aircraft wing structure.
52. The aluminum structural component of claim 47 in the form of a
rib for use in an aircraft wing structure.
53. The aluminum structural component of claim 47 in the form of a
forged product for use in an aircraft wing structure.
54. The aluminum structural component of claim 47 in the form of an
extruded product for use in an aircraft wing structure.
55. The aluminum structural component of claim 47 having a
quarter-plane tensile yield strength, TYS (L), greater than about
69 ksi and a fracture toughness, K.sub.Ic, (L-T), greater than
about 30 to 34 ksi{square root}in wherein said alloy has been
processed in a solution heat treated, quenched, stretched and
artificially aged condition.
56. A mold plate made from an aluminum alloy consisting essentially
of, in % by weight, about: 6 to 10 Zn, 1.3 to 1.9 Mg, 1.4 to 2.2
Cu, 0 to 0.4 Zr 0 to 0.4 Sc 0 to 0.2 Hf 0 to 0.4 Cr 0 to 1.0 Mn,
and the balance being Al and incidental additions and
impurities.
57. A process for making a structural component for a commercial
jet aircraft comprising the steps of: (a) providing a 7XXX series
aluminum alloy consisting essentially of, in % by weight, about 6
to 10 Zn, 1.3 to 1.9 Mg, 1.4 to 2.2 Cu, wherein Mg.ltoreq.Cu+0.3, 0
to about 0.4 Zr, 0 to about 0.4 Sc, 0 to about 0.2 Hf. 0 to about
0.4 Cr, 0 to about 1.0 Mn, and the balance being Al and incidental
additions and impurities; (b) homogenizing and hot forming said
alloy by one of rolling, extruding or forging to a desired
workpiece form; (c) solution heat treating said hot formed
workpiece at a temperature and time sufficient to place into solid
solution all soluble constituents in the alloy; (d) quenching said
solution heat treated workpiece by one of spray quenching or
immersion quenching in water or other quenching media; (e)
artificially aging said quenched workpiece to achieve a desired
temper; and (f) machining said workpiece to provide a desired
configuration of said structural component.
58. The process of claim 57 including the step of stress relieving
the workpiece after said quenching step (d) by one of stretching,
compressing or cold working.
59. The process of claim 57 wherein said quenched workpiece has a
thickness of greater than about 1.5 inches.
60. The process of claim 57 wherein said quenched workpiece has a
thickness of greater than about 2 inches.
61. The process of claim 57 wherein said quenched workpiece has a
thickness of greater than about 3 inches.
62. The process of claim 57 wherein the quenched workpiece has a
thickness of greater than 6 inches, up to about 12 inches.
63. The process of claim 57 wherein t he workpiece is intentionally
quenched slowly following the solution heat treating step (c).
64. The process of claim 57 wherein the provided 7XXX series
aluminum alloy of step (a) consists essentially of: 6.5 to 8.5 Zn,
1.3 to 1.7 Mg, 1.4 to 2.0Cu, wherein Mg.ltoreq.1.7 and
Mg.ltoreq.Cu+0.3 0 to 0.4 Zr, 0 to about 0.38 Sc, 0 to about 0.20
Hf, 0 to about 0.37 Cr, 0 to about 1.0 Mn, and the balance being Al
and incidental additions and impurities.
65. The process of claim 64 wherein said quenched workpiece has a
thickness of greater than about 1.5 inches.
66. The process of claim 64 wherein said quenched workpiece has a
thickness of greater than about 2 inches.
67. The process of claim 64 wherein said quenched workpiece has a
thickness of greater than about 3 inches.
68. The process of claim 64 wherein the quenched workpiece has a
thickness of greater than 6 inches, up to about 12 inches.
69. The process of claim 64 wherein the workpiece is intentionally
quenched slowly following the solution heat treating step (c).
70. The process of claim 57 wherein the structural component has a
quarter-plane tensile yield strength, TYS (L), greater than about
69 ksi and a fracture toughness, K.sub.Ic (L-T), greater than about
30 to 34 ksi{square root}in.
71. A process for making a mold plate comprising the steps of: (a)
providing a 7XXX series aluminum alloy consisting essentially of,
in % by weight, about 6 to 10 Zn, 1.3 to 1.9 Mg, 1.4 to 2.2 Cu,
wherein Mg.ltoreq.Cu+0.3, 0 to about 0.4 Zr, 0 to about 0.4 Sc, 0
to about 0.2 Hf, 0 to about 0.4 Cr, 0 to about 1.0 Mn, and the
balance being Al and incidental additions and impurities; (b)
homogenizing and hot forming said alloy by one of rolling,
extruding or forging to a desired workpiece form; (c) solution heat
treating said hot formed workpiece at a temperature and time
sufficient to place into solid solution all soluble constituents in
the alloy; (d) quenching said solution heat treated workpiece by
one of spray quenching or immersion quenching in water or other
quenching media; (e) artificially aging said quenched workpiece to
achieve a desired temper; and (f) machining said workpiece to
provide a desired configuration of said structural component.
72. The process of claim 71 wherein the workpiece is intentionally
quenched slowly following the solution heat treating of step (c).
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] The present invention relates generally to aluminum alloys
and, more particularly, to 7XXX series aluminum alloys having
superior strength-toughness combinations suitable for thick gauge
(e.g., 2-10 inches) structural parts in aerospace applications.
Parts made from the alloy of the present invention find specific
utility as structural components such as integral spar members and
the like which are integrally machined from thick sections,
including rolled plate, used in the construction of structural
members for high capacity aircraft. The alloy may also be formed by
other known hot forming techniques such as extrusion and forging.
The forging process is particularly suitable for manufacturing high
strength aircraft components, such as, for example, main landing
gear beams.
[0003] 2. Background of the Invention
[0004] As the size of new aircraft, or modification of current
models, gets larger to accommodate heavier payload and/or longer
flight range for improved performance and economy, the demand on
weight savings of structural components, such as fuselage, wing,
spar, etc., continues to increase. The aircraft industry is meeting
this demand through specification of higher strength requirements
of metal parts to enable reduced section thickness as a weight
savings expedient. However, during the late 1960's, those skilled
in the field of fracture mechanics discovered that, in addition to
the strength, the durability and damage tolerance of the material
are also critical to the fail-safe aircraft structural design. Such
consideration of multiple material attributes for aircraft
applications eventually led to today's damage tolerance design,
which combines principles of fail-safe design with periodic
inspection techniques.
[0005] Traditionally, an aircraft wing structure comprises a wing
box designated generally by reference numeral 2 in FIG. 1. The wing
box 2 is attached to upper and lower wing skins 4 and 6,
respectively, connected by vertical structural members called spars
12 and 20. During flight, the upper wing structure of a commercial
aircraft is compressively loaded, calling for high strength with
acceptable fracture toughness as desired material attributes. The
upper wing is typically built with a 7XXX series aluminum alloy
such as 7150 (U.S. Reissue Pat. No. 34,008) or 7055 (U.S. Pat. No.
5,221,377). The lower wing structure of a commercial aircraft is
under tension during flight and, therefore, requires high damage
tolerance. Although it is desirable to design the lower wing with a
high strength alloy to maximize weight efficiency, the damage
tolerance characteristics of the high strength 7XXX series alloys
fall short of the need. Therefore, most commercial aircraft
manufacturers specify a damage-tolerant 2XXX series aluminum alloy
such as 2024 or 2324 for lower wing applications (U.S. Pat. No.
4,294,625), both of which are significantly lower in strength than
the 7XXX series upper wing alloys. The alloy members and temper
designations used herein are in accordance with the well-known
aluminum alloy product standards of the Aluminum Association.
[0006] The upper and lower wing skins 4 and 6, respectively, are
typically stiffened by longitudinally extending stringer members 8,
10 in J, I, T, Z or other shapes which are fastened to the inside
surfaces of the wing skins, as shown in FIG. 1. The upper wing
stringers 8 and upper spar caps 14, 22 are usually manufactured
from a 7XXX series aluminum alloy, and the lower wing stringers 10
and spar caps 16, 24 with a 2XXX series aluminum alloy for the same
structural reasons discussed above regarding their respective
strength and damage-tolerance properties. Vertical web members 18,
26 fastened to both the upper and the lower spar caps run along the
longitudinal direction of the wing, constituting the structural
member spars 12 and 20. This traditional spar design is also called
a "built-up" spar, comprising the upper spar cap 14 or 22, the web
18 or 20 and the lower spar cap 16 or 24 with fasteners (not
shown). Obviously, the fasteners and fastener holes at the joints
are structural weak links. In order to ensure the structural
integrity of the built-up spars 18 and 20, many parts of the web
and spar caps have to be thickened, thereby adding weight to the
structure.
[0007] A desirable design approach to overcome the aforementioned
weight penalty imposed by the built-up spar is to make the upper
spar, the web and the lower spar out of a single piece of aluminum
alloy product. This design eliminates the web-to-upper-spar and
web-to-lower-spar joints, and can provide significant weight
savings with respect to the built-up spar. The one-piece spar is
called an "integral spar," and can be machined to final shape from
a thick plate, a thick extrusion or a thick forging. The integral
spar not only weighs less, but also involves less cost to
manufacture and assemble due to the elimination of fasteners. The
ideal alloy for the integral spar is one which has the same
strength as the upper wing skin with sufficient fracture toughness
to meet the damage tolerant requirements of the lower skin.
However, current alloys do not meet the desired property
requirements. For example, the lower strength of alloy 2024-T351,
commonly used in lower wing applications, will not be able to
safely carry the load transmitted from the highly loaded upper wing
unless its section thickness is significantly increased, which adds
undesirable weight to the aircraft. Alternatively, to design the
upper wing according to the 2XXX strength capability would result
in overall weight penalty. The higher strength 7XXX series-T7X
alloys in thick gauges, such as 7050-T74 (U.S. Pat. No. 3,881,966)
still fall short of the need in strength in the required thickness,
as elaborated below.
[0008] Emerging large jet aircrafts require very large wings which,
in turn, require integral spar alloy products as thick as 6 to 8
inches or more for weight efficiency. Today's industry standards
for thick plate, for example, 6 inch thick 7050-T7451, are given in
the Aerospace Materials Specifications AMS 4050F, setting forth a
minimum yield strength of 60 ksi in the longitudinal (L) direction
and plane strain fracture toughness, K.sub.Ic (L-T) of 24
ksi{square root}in, and those for the transverse direction (LT and
L-T) are, respectively, 60 ksi and 22 ksi{square root}in. However,
the most recently developed upper wing skin specified on the Boeing
777, e.g., aluminum alloy 7055-T7751, is capable of meeting a yield
strength minimum of 86 ksi according to the MIL-HDBK-5H. Therefore,
if an integral spar with 60 ksi yield strength is used, the
strength capability of the upper wing skin will not be able to be
taken full advantage of for maximum weight efficiency. Hence, a
higher strength, thick spar alloy with sufficient fracture
toughness is highly desirable for the construction of an integral
spar. This is but one specific example of the benefits of an
aluminum material with high strength and toughness in thick
sections, but many others exist in modern aircraft.
[0009] It is known that the various tempers resulting from
different artificial aging treatments will impart different levels
of strength and other performance characteristics such as corrosion
resistance and fracture toughness. The 7XXX series aluminum alloys
are nominally in artificially aged tempers such as peak strength
T6-type or over-aged T7-type tempers. U.S. Pat. Nos. 4,863,528;
4,832,758; and 4,477,292, along with U.S. Pat. No. 5,108,520 all
describe tempers for 7XXX series alloys to provide a range of
combinations of strength and performance. All of the aforesaid
patents are fully incorporated herein by reference. It is
well-known to those skilled in the art that for a given 7XXX series
alloy, the peak strength T6-type temper provides the highest
strength combined with the lowest fracture toughness and corrosion
resistance. It is also known that the most over-aged temper such as
T73-type temper for the same alloy provides the highest fracture
toughness and corrosion resistance combined with the lowest
strength. In practice, an appropriate temper is chosen somewhere
between these two extremes to suit a particular application. A more
complete description of the temper description (the "T-XX" suffix)
is given in the Aluminum Standards and Data 2000, published by The
Aluminum Association, Inc.
[0010] The processing of most aerospace alloys generally requires a
solution heat treatment (SHT) followed by quenching and subsequent
artificial aging. In the quest for a stronger material, alloy
designers are faced with two natural phenomena. As the product
shape gets thicker, the quench rate experienced at the interior of
the product cross section naturally decreases, which results in a
loss of strength and fracture toughness in the product. This
important phenomenon is known as "quench sensitivity". Also, there
is an inverse relationship between strength and fracture toughness;
that is, as the alloy gets stronger, its toughness tends to
decrease. Up until now, no one has been able to solve the problem
concerning how to avoid these conflicting phenomena and provide a
stronger, thick alloy product combined with appropriate high
fracture toughness.
[0011] To better understand the present invention, it is helpful to
note certain demonstrated trends in the art concerning 7XXX series
aluminum alloy development efforts to meet demands for improved
properties.
[0012] Type 7050 aluminum alloy employs Zr in place of Cr as a
dispersoid agent and realizes a significant improvement in quench
sensitivity over the prior 7075 alloy and, thus, has been the
mainstay for thick section application as plate, extrusion or
forging in aerospace service. To meet higher strength requirements,
the alloy contents in 7050 alloy are higher than in 7075. For
example, the lower limit of Cu (composition as registered with the
Aluminum Association) increased from 1.2 weight % in 7075 to 2.0
weight % in 7050 alloy and, likewise, Zn from 5.1 to 5.7 weight %.
In 7150 alloy for upper wing application with still higher
strength-toughness requirements, the Mg and Zn lower limits also
slightly increased compared to those in 7050, namely, Zn increased
from 5.7 to 5.9 weight %, and Mg from 1.9 to 2.0 weight %,
respectively, while simultaneously the metal purity was also
improved by lowering the Fe and Si levels. A newer generation upper
wing skin alloy, represented by 7055 alloy, provides a 10 percent
improvement in compression yield strength properties and employs
compositions with a much higher Zn range of 7.6 to 8.4 weight %
while maintaining similar Cu levels and a slightly lower Mg range
at 1.8 to 2.3 weight % compared to either 7050 or 7150 alloys.
[0013] Thus, the past trend against any drive for higher strength
had been to increase alloying additions and composition
optimization as dictated by phase equilibrium relations, while
metal purity increase and microstructure control through
thermal-mechanical processing (TMP) were also sought in order to
obtain simultaneous improvements in toughness and fatigue life,
amongst other properties.
[0014] U.S. Pat. No. 5,865,911 reports a significant improvement in
toughness at equivalent strengths in plates of a 7XXX series alloy
less than 2.5 inches thick. The improvement was obtained through
optimization of alloy compositions by careful use of phase
equilibrium relations, coupled with other innovations. However, the
purported superior strength-toughness advantage was significantly
reduced in this alloy when a quench rate simulating the mid-plane
of a 6-inch product is used after SHT, showing only limited
improvement in properties over that of 7050 alloy. Thus, quench
sensitivity in this alloy has been responsible for the loss in
property advantages in thicker gauges.
[0015] In another example, 7040 alloy registered with the Aluminum
Association reports the following composition, in weight percent,
for the major alloying elements: 5.7 to 6.7 Zn, 1.7 to 2.4 Mg and
1.5 to 2.3 Cu. Published literature (Shahani et al., "High Strength
7XXX Alloys For Ultra-Thick Aerospace Plate: Optimization of Alloy
Composition," Proc. ICAA-6, Vol. 2, 1998; see, also, U.S. Pat. No.
6,027,582 to Shahani et al.) teaches that an optimization balance
was pursued between alloying additions in an effort to improve
strength and other properties while avoiding excess additions to
minimize quench sensitivity. While some property improvements over
7050 alloy were claimed in thicker gauges, the improvements of 7040
still fell short of those desired for newer commercial aircraft
designs.
[0016] The instant invention solves the problems encountered in the
prior art by providing a 7XXX series aluminum alloy which exhibits
significantly reduced quench sensitivity so as to provide
significantly higher strength and fracture toughness levels than
heretofore possible in thick gauge aerospace structural
members.
SUMMARY OF THE INVENTION
[0017] Briefly stated, the present invention is directed to a 7XXX
series aluminum alloy having significantly reduced quench
sensitivity in thick gauges, i.e., greater than about 2 inches and,
more preferably, in the thickness range of about 4 to 8 inches or
greater. A presently preferred broad composition of the alloy of
the present invention consists essentially of, in weight %: about 6
to 10 zinc (Zn), about 1.3 to 1.9 magnesium (Mg), about 1.4 to 2.2
copper (Cu), wherein Mg.ltoreq.Cu+0.3, about 0 to 0.4 zirconium
(Zr), about 0 to 0.4 scandium (Sc), about 0 to 0.2 hafnium (Hf),
about 0 to 0.4 chromium (Cr), about 0 to 1.0 manganese (Mn), the
balance being aluminum (Al) and other incidental elements.
[0018] A more narrow, presently preferred alloy composition
according to the present invention consists essentially of, in
weight %: about 6 to 8.5 Zn, about 1.3 to 1.8 Mg, about 1.4 to 2.0
Cu, wherein, Mg.ltoreq.Cu+0.3, one or more elements selected from
the group consisting of up to about 0.4 Zr, up to about 0.4 Sc, up
to about 0.2 Hf. up to about 0.4 Cr, up to about 1.0 Mn, the
balance being Al , incidental additions and impurities.
[0019] A still more narrowly defined, presently preferred alloy
composition according to the present invention consists essentially
of, in weight %: about 6.5 to 8.5 Zn, about 1.3 to 1.7 Mg, about
1.4 to 2.0 Cu, wherein Mg.ltoreq.1.7 and Mg.ltoreq.Cu+0.3, one or
more elements selected from the group consisting of up to about 0.4
Zr, up to about 0.38 Sc, up to about 0.20 Hf. up to about 0.37 Cr,
up to about 1.0 Mn, the balance being Al and other incidental
additions and impurities. The above defined alloys may contain
impurities and other incidental/intentionally made additions common
and well-known to the 7XXX series family of aluminum alloys, such
as on the order of, for example, in weight %: nominally about 0.03
or up to about 0.12 maximum silicon (Si), nominally about 0.05 or
up to about 0.15 max. iron (Fe), nominally about 0.025 or up to
about 0.15 max. titanium (Ti) and the like. The "other" additions
are generally governed by the 0.05-0.15 ranges as defined in the
alloy designations by the Aluminum Association.
[0020] The elements Zr, Sc, Hf. Cr and Mn are introduced as
dispersoid forming elements aimed at providing an unrecrystallized
or partially recrystallized grain structure in the invention
wrought products. Such grain structures are required to achieve the
highest combination of strength, fracture toughness and stress
corrosion resistance. The above described dispersoid forming
elements are substantially in supersaturation after casting, and
form fine dispersoid particles by solid state reactions during
thermal-mechanical processing. Zr forms Al.sub.3Zr, Sc forms
Al.sub.3Sc, Hf forms Al.sub.3Hf, Cr forms either
Al.sub.12Mg.sub.32Cr or Al.sub.18Mg.sub.3Cr.sub.2 and Mn forms
Al.sub.20Cu.sub.2Mn.sub.3 dispersoid particles. The dispersoid
particles retard or stop recrystallization by exerting a drag force
on the recrystallization nuclei. The homogenization process for the
invention alloys, in fact, is aimed not only at evenly
redistributing and dissolving the cored micro-segregation of the
major alloying elements but also at causing an optimum, copious
distribution of dispersoids for controlling the grain structure
during subsequent thermal-mechanical processing. When multiple
dispersoid elements are present, there may be synergistic effects
for grain structure control. The optimum dispersoid content
depends, in part, on the solidification process and, in part, on
the thermal-mechanical process.
[0021] The alloys of the present invention are conventionally
prepared by melting and may be direct chill (D.C.) cast into ingot
form. Conventional grain refiners such as titanium boride may also
be used as well-known in the art. After conventional scalping and
homogenization (if needed), the ingots are further processed by,
for example, hot rolling into plate or extrusion or forging into
special shaped sections. Generally, the heavy sections are on the
order of greater than 2 inches and, more typically, on the order of
4, 6, 8 or up to 12 inches or more in thickness. In the case of
heavy plate of about 4 to 8 inches in thickness, the plate is
solution heat treated and quenched and mechanically stress relieved
such as by stretching, for example, up to 8%, or compression. A
desired structural shape is then machined from these heat treated
heavy plate sections to form the final part, such as, for example,
an integral wing spar. Similar SHT, quench and often stress relief
operations are also followed in the manufacture of thick sections
made from extrusions and forgings.
[0022] Good combinations of properties are desired in all
thicknesses, but they are particularly useful in thickness ranges
where, conventionally, as the thickness increases, quench
sensitivity of the product also increases. Hence, the alloy of the
present invention finds particular utility in heavy gauges of, for
example, greater than 4 inches in thickness up to 12 inches or
more.
BRIEF DESCRIPTION OF THE DRAWINGS
[0023] FIG. 1 is a transverse cross-sectional view of a typical
wing box construction of an aircraft including front and rear spars
of conventional three-piece built-up design;
[0024] FIG. 2 is a graph showing two calculated cooling curves to
approximate the mid-plane cooling rates for 6- and 8-inch thick
plates under spray quenching, and superimposed on top are two
experimental cooling curves simulating the cooling rates of a
6-inch thick and an 8-inch thick plate;
[0025] FIG. 3 is a graph showing tensile yield strength TYS (L)
versus fracture toughness K.sub.q (L-T) relations for selected
alloys of the present invention and other alloys including type
7150 and 7055 aluminum alloys as controls, all based on simulation
of quarter-plane quench rates of 6-inch thick plate, extrusion or
forging;
[0026] FIG. 4 is a graph similar to FIG. 3 showing tensile yield
strength TYS (L) versus fracture toughness K.sub.q (L-T) relations
for selected alloys of the present invention and other alloys
including 7150 and 7055 controls, all based on simulation of
quarter-plane quench rates of 8-inch thick plate, extrusion or
forging;
[0027] FIG. 5 is a graph showing the influence of Zn content on
quench sensitivity as demonstrated by directional arrows for the
tensile yield strength changes in a 6-inch thick plate quench
simulation;
[0028] FIG. 6 is a graph showing the influence of Zn content on
quench sensitivity as demonstrated by directional arrows for the
tensile yield strength changes in an 8-inch thick plate quench
simulation;
[0029] FIG. 7 is a graph showing cross plots of tensile yield
strength TYS (L) versus plane-strain fracture toughness
K.sub.Ic,(L-T) values at quarter plane of a full-scale production
6-inch thick plate of the invention alloy and of 7050 and 7040
alloys from literature; and
[0030] FIG. 8 is a graph showing the influence of section thickness
on tensile yield strength, as an index of quench sensitivity
property, from a full-scale production die-forging study comparing
alloys of the invention and 7050 alloys.
DETAILED DESCRIPTION OF THE INVENTION
[0031] Mechanical properties of importance for the thick plate,
extrusion or forging for aircraft structural products, as well as
other non-aircraft structural applications, include strength, both
in compression as for the upper wing skin and in tension for the
lower wing skin. Also of importance are the fracture toughness,
both plane-strain and plane-stress, and corrosion resistance such
as, for example, exfoliation and stress corrosion cracking
resistance.
[0032] Integral wing spars and wing skin panels with integral
stringers must be machined from relatively thick plates or other
structural shapes which are extruded or forged from thick ingots or
billets and which have been then solution heat treated, quenched
and artificially aged. It is not always feasible to solution heat
treat and quench the finished structural products because the rapid
cooling of the quenching step would induce residual stress and
cause dimensional distortion. The quench-induced residual stress
could also possibly cause stress corrosion cracking, and re-work to
straighten parts associated with dimensional distortion could
render assembly impracticably difficult. While it is much easier to
obtain better mechanical properties in thinner cross sections
because of the faster cooling which prevents unwanted precipitation
of alloying elements, this cannot be done when quench distortion is
present. Hence, it is necessary to solution heat treat and quench
the thick plates, extrusions or forgings, mechanically straighten
and flatten the wrought product while simultaneously relieving the
residual stress, followed by artificial aging to produce the
desired final temper. The wrought product is then machined to
achieve the desired shape of the finished structural component.
[0033] As alluded to above, in solution heat treating and quenching
thick sections, the quench sensitivity of the aluminum alloy is of
great concern. After solution heat treating, it is desirable to
quickly cool the material in order to retain the various alloying
elements in solid solution rather than to allow them to precipitate
out of solution in coarse form as occurs in slow cooling. The
latter occurrence producing the coarse precipitates results in a
decline in mechanical properties. In thick product cross sections,
i.e., over 3 inches, and more particularly in heavier sections of 4
to 8 inches or more, the quenching medium acting on the exterior
surfaces of the workpiece (such as a plate, forging or extrusion,
for example) cannot efficiently extract heat from the center or
mid-plane region of the material. This is due to the physical
distance to the surface and the fact that heat is extracted through
the metal by conduction which is distance dependent. In thinner
product cross sections, quench rate at the mid-lane is naturally
higher than that in heavier sections. Hence, the quench sensitivity
property of an alloy is not as important in thinner gauge shapes as
it is in heavier gauge workpieces.
[0034] The present invention is, of course, directed at increasing
the strength-toughness properties in a 7XXX series aluminum alloy
in thicker gauges, i.e., greater than about 1.5 inches, and, thus,
the quench sensitivity of the alloy is of extreme importance. In
thicker gauges, the less quench sensitivity the better, with
respect to a material's ability to retain alloying elements in
solid solution (to avoid the formation of coarse precipitates upon
slow cooling) particularly in the slowest-cooled mid-plane region.
The present invention achieves the goal of lower quench sensitivity
by providing a carefully controlled alloy composition which permits
the manufacture of heavier gauges while achieving superior
strength-toughness properties.
[0035] The present invention resulted from an exploration for
possible property advantages in 7XXX series alloys based upon
certain phase equilibrium features that had been identified by the
present inventors in the Al--Zn--Cu--Mg alloy system. The intent
was to explore if increased additions of Zn could expand the matrix
phase field (based on aluminum with a face centered cubic crystal
structure) to allow for increased solubility of other alloying
additions such as Cu and Mg for increased strength.
[0036] The experimental work consisted of the following:
Twenty-eight 11-inch diameter ingots were direct chill (DC) cast,
homogenized and extruded into 1.25 inch thick by 4-inch wide
rectangular bars. The bars were next solution heat treated and were
quenched at different rates to mimic cooling conditions for thin
section as well as those approximating 6-inch and 8-inch thick
sections. The bars were then cold stretched by 1.5 percent for
residual stress relief. The compositions of alloys studied are set
forth in Table 1 below, in which the Zn content ranged from 6.0
weight % to slightly in excess of 11.0 weight %, while Cu and Mg
each were varied between 1.5 and 2.3 weight %, respectively.
1TABLE 1 Invention Composition Invention Composition Specimen Alloy
(wt. %) Specimen Alloy (wt. %) No. Y/N Cu Mg Zn No. Y/N Cu Mg Zn 57
Y 1.57 1.55 6.01 71 N 1.86 1.93 10.93 58 N 1.64 2.29 5.99 72 N 1.98
2.09 11.28 59 N 2.45 1.53 5.86 73 N 1.97 1.86 9.04 60 N 2.43 2.26
6.04 74 Y 1.48 1.50 9.42 61 N 1.95 1.94 6.79 75 N 1.75 2.29 9.89 62
Y 1.57 1.51 7.56 76 N 2.48 1.52 9.60 63 N 1.59 2.30 7.70 77 N 2.19
2.19 9.74 64 N 2.45 1.54 7.71 78 N 1.68 1.55 11.38 65 N 2.46 2.31
7.70 79 N 1.65 2.28 11.04 66 N 2.05 1.92 8.17 80 N 2.38 1.53 11.08
67 Y 1.53 1.52 8.65 81 N 2.22 1.97 9.04 68 N 1.57 2.35 8.62 82 N
1.79 2.00 10.17 69 N 2.32 1.45 8.25 83 N 2.23 2.28 6.62 70 N 2.04
2.19 8.33 84 N 2.48 1.98 8.31
[0037] For all alloys other than the controls: Target Si=0.03,
Fe=0.05, Zr=0.12, Ti=0.025
[0038] For 7150 Control (S# 83): Target Si=0.05, Fe=0.10, Zr=0.12,
Ti=0.025
[0039] For 7055 Control (S# 84): Target Si=0.07, Fe=0.11, Zr=0.12,
Ti=0.025
[0040] Different quenching approaches were explored in order to
obtain at the mid-plane of the 1.25 inch thick extruded bars a
cooling rate mimicking that at the mid-plane of a 6-inch thick
plate subjected to spray quenching in 75.degree. F water in a
full-scale production plant. A second simulation involved
mimicking, under identical circumstances, a bar cooling rate
corresponding to that of an 8-inch thick plate.
[0041] The quenching simulation involved modification of the heat
transfer characteristics of the quenching medium, as well as the
surface of the part by immersion quenching of the extruded bars
with simultaneous incorporation of three practices: a defined warm
water temperature, saturation of the water with CO.sub.2 gas, d a
chemical treatment of the bars to render a bright etch surface
finish to it.
[0042] For mimicking the 6-inch thick plate cooling condition, the
water temperature for immersion quenching was held at 180.degree.
F., the solubility level of CO.sub.2 in the water was about 0.20
lan (a measure of dissolved CO.sub.2 concentration, lan=Standard
volume of CO.sub.2/Volume of water), and the surface had a standard
bright etch finish.
[0043] For the 8-inch thick plate cooling simulation, the water
temperature utilized was 190.degree. F. with a CO.sub.2 solubility
reading varying between 0.17 and 0.20 Ian together with a standard
bright etch surface finish.
[0044] The cooling rates were measured by thermocouples inserted
into the mid-plane of the bar samples. For benchmark reference, the
two calculated cooling curves to approximate the mid-plane cooling
rates under spray quenching at the plants for the 6-and 8-inch
thick plates were plotted, FIG. 2. Superimposed on them were
displayed two groups of plots, the lower group (in the temperature
scale) representing cooling curves simulating mid-plane of the
6-inch and the upper group simulating mid-plane of the 8-inch thick
plate cooling rates. The cooling rates of the simulation were very
similar to those of the plant production plates in the important
temperature range above 500.degree. F., although the simulated
cooling curves for the experimental material differed from those of
the plant plates below 500.degree. F., which was not considered
critical. Hence, the simulation was considered successful.
[0045] After solution heat treating and quenching, aging behaviors
were studied using multiple aging times and aided by EXCO
(exfoliation corrosion resistance), EC (electrical conductivity)
and SCC (stress corrosion cracking resistance) measurements. A
two-step aging practice was used that consisted of a slow heat-up
in 6 hours to 250.degree. F., a 4-hour soak at 250.degree. F.
followed by second step aging soak at 320.degree. F. for variable
times ranging from 4 to 36 hours depending on the alloy composition
in order to obtain acceptable EC and EXCO readings. The different
aging steps were optimized not only to obtain improved
strength-toughness properties, but also to obtain good corrosion
resistance.
[0046] Tensile and compact tension plane-strain fracture toughness
test data were collected on samples that had been given the
different minimum aging times required in order to obtain a visual
EXCO rating of EB or better (EA) for an acceptable exfoliation
corrosion resistance property, and an electrical conductivity (EC)
reading minimum of 36% IACS (International Annealed Copper
Standard), also used as an indicator for degree of over-aging and
corrosion resistance. All tensile tests were performed according to
the ASTM specification E8, and all plane-strain fracture toughness
according to the ASTM specification E399.
[0047] The original intent in the investigation, as stated above,
was to use increased solutes in an expanded alloy phase field, made
possible through high Zn additions, in an effort to obtain improved
strength-toughness properties. However, it was surprising to find
that, contrary to our expectations, the best combinations of
strength-toughness properties were obtained at much lower solute
compositions of Cu and Mg than anticipated or heretofore tried in
the industry.
[0048] We have found that it is desirable to use Zn at higher
levels (above 6.7 wt. %, preferably around 7.5 wt. %), in contrast
with the recent 7040 alloy with Zn specified between 5.7 and 6.7
wt. %, as well as the earlier 7050 and 7010 alloys.
[0049] We have also found that for the best strength-toughness
properties, a concomitant decrease in the Cu and Mg levels is
required. Thus, the Zn levels in the alloy of our invention are
higher, and the Cu and Mg levels are lower than those specified and
practiced in the prior art. The Mg range for 7040 alloy is 1.7 to
2.4 wt. %. The Mg range for 7050 alloy is 1.9 to 2.6 wt. % and for
7010 alloy 2.1 to 2.6 wt. %. The desirable Mg range in the present
invention is 1.3 to 1.9 wt. %, with a more preferred range being
1.3 to 1.7 wt. % for best properties.
[0050] The Cu range registered by the Aluminum Association for 7040
is 1.5 to 2.3 wt. %, while the Cu range for the present invention
is 1.4 to 2.2 wt. %, with a more preferred content at no more than
2.0 wt. % and, still more preferably, no more than 1.8 wt. % Cu for
best properties. We have also discovered that the performance is
enhanced when Cu is greater than Mg, or when Cu+0.3 wt. % is
greater than or equal to Mg, contrary to the disclosure of 7040
(U.S. Pat. No. 6,027,582), where Mg content is required to be
higher than Cu.
[0051] FIG. 3 shows the strength-toughness plot of results from
slow quenching of alloy samples of Table 1 from the SHT temperature
to simulate a 6-inch thick product. As is readily noticed in these
plots, a family of alloy compositions stood out, namely, alloy
numbers 57, 62, 67 and 74 (upper part of the diagram) from the rest
which displayed very high fracture toughness combined with high
strength properties. Surprisingly, all these alloy compositions
belonged to the low-Cu and low-Mg composition ends of our choice,
namely, at around 1.5 wt. % Mg together with 1.5 wt. % Cu, while Zn
levels varied from 6.0 to 9.5 wt. %.
[0052] The Zn levels for the improved alloys illustrated were
about: 6.0 wt. % for alloy #57, 7.5 wt. % for alloy #62, 8.6 wt. %
for alloy #67 and 9.5 wt. % for alloy #74.
[0053] The substantial improvement in the strength-toughness
properties can be seen by comparison against the two control alloys
7150 (alloy #83) and 7055 (alloy #84) which had been processed in
identical manner. In FIG. 3, a dotted line connecting the two data
points has been drawn to show their strength-toughness property
trend which positions the control alloys considerably below the low
solute data points discussed.
[0054] Also included in the plots are results for alloys with about
1.9 wt. % Mg and 2.0 wt. % Cu at varying Zn levels of 6.8 wt. %
(alloy #61), 8.2 wt. % (alloy #66), 9.0 wt. % (alloy #73) and 10.2
wt. % (alloy #82). The results show a dramatic drop in toughness in
these alloys compared to the 1.5 wt. % Mg and 1.5 wt. % Cu
containing alloys at each of the corresponding Zn levels. The
strength-toughness properties in thick gauge product in these
higher Mg and Cu alloys are similar to or marginally better than
the 7150 and 7055 control alloy trend line. The results clearly
demonstrate the strong degradation in properties that occurs with a
moderate increase in Cu and Mg above the low Cu and Mg levels in
the alloys of the present invention, and approaching the levels of
many of the current commercial alloys.
[0055] A similar set of results with similar conclusions as above
is depicted in FIG. 4 for a quench condition which was even slower
than in FIG. 3. The condition of FIG. 4 roughly approximates an
8-inch thick plate mid-plane cooling condition.
[0056] Thus, unlike the past teachings, some of the highest
strength-toughness properties were obtained at some of the leanest
Cu and Mg levels not considered thus far in the current commercial
alloys. Concomitantly, the Zn levels at which the properties were
most optimized corresponded to levels much higher than those cited
in type 7040, 7050 or 7010 aluminum thick product alloys.
[0057] The substantial improvement in strength-toughness properties
in quench sensitive thick sections in the alloys of the invention
conceivably resulted from a significant reduction in the quench
sensitivity of the alloys owing to the lean Cu and Mg compositions,
while higher Zn additions facilitated this through expanding the
stability of the matrix phase field. This conclusion is supported
in FIG. 5 where the strength (TYS) increased gradually with
increasing Zn content for alloys #57 to #62 to #67. Thus, unlike
the past teachings in the literature, including that for 7040
alloy, higher solute in the case of Zn did not hurt quench
sensitivity and was, in fact, proven to be beneficial against slow
quench conditions, i.e., thick sections. However, at a still higher
Zn level of 9.4 wt. % the resultant enlarged phase field could not
further compensate for the increased solute level, and the strength
(TYS) of alloy #74 dropped below the other lower Zn alloys above,
see FIG. 5. Also, with further, still slower quench conditions in
simulated 8-inch thicknesses as depicted in FIG. 6, the quench
sensitivity increased even at 8.7 wt. % Zn as depicted by the
strength (TYS) of alloy #67 displaced below that of alloy #62 with
7.6 wt. % Zn. The high solute effect on quench sensitivity is also
depicted in the relative positioning on the strength axis of the
two control alloys, 7150 (alloy #83) and 7055 (alloy #84). 7055 was
stronger than 7150 under slow quench (FIG. 5), but the relative
scale was reversed under still slower quench conditions as in FIG.
6.
[0058] It is also of interest to note alloy #63, which, according
to Table 1, contained 1.59 wt. % Cu, 2.30 wt. % Mg and 7.70 wt. %
Zn, (i.e., the Mg>Cu). Alloy #63 exhibited in FIG. 3 a high
strength (TYS) of about 73 ksi but a relatively low fracture
toughness K.sub.q (L-T) of about 23 ksi{square root}in. By way of
comparison, alloy #62, which contained 1.57 wt. % Cu, 1.51 wt. % Mg
and 7.56 wt. % Zn (i.e., Mg<Cu), exhibited in FIG. 3 a strength
(TYS) of greater than 75 ksi and a high fracture toughness of about
34 ksi{square root}in (a 48% increase in toughness). The data show
the criticality of maintaining the Mg content below 1.7 wt. % as
well as the importance of keeping the Mg content generally and
preferably no greater than the Cu content or no grater than Cu+0.3
wt. %.
[0059] It is desirable to achieve optimum and/or balanced fracture
toughness (K.sub.q) and strength (TYS) properties in the alloys of
the present invention. As can be seen and appreciated by comparing
the compositions set forth in Table 1 with the corresponding
fracture toughness and strength values plotted in FIG. 3, those
sample alloys falling within the compositions of the present
invention achieve this balance of properties. It will be seen that
the alloy samples having compositions according to the present
invention, namely, alloys #57, #62, #67, and #74, either possess a
fracture toughness value (K.sub.q) in excess of about 34 ksi{square
root}in with a tensile yield strength (TYS) of greater than 69 ksi
or they possess a fracture toughness value of greater than about 29
ksi{square root}in combined with a higher tensile yield strength of
greater than about 75 ksi.
[0060] It will also be seen that the upper limit of the Zn content
is important in achieving balanced toughness and strength
properties. Those alloy samples which exceeded the upper limit on
Zn of 11.0 wt. %, such as alloy #80 (11.08 wt. % Zn), or alloy #78
(11.38 wt. % Zn), failed to possess the minimum combined strength
and fracture toughness levels set forth above for alloys of the
invention.
[0061] The alloy of the present invention, thus, provides high
damage tolerance in thick aerospace structures resulting from its
enhanced, combined fracture toughness and yield strength
properties.
[0062] K.sub.q values are results of plane strain fracture
toughness tests that do not conform to validity criteria of ASTM
E399. In the current tests yielding K.sub.q values, the validity
criteria that were violated were: (1) Pmax/Pq.ltoreq.1.1 primarily,
and (2) B (thickness).gtoreq.2.5 (K.sub.q/.sigma..sub.ys).sup.2
occasionally, where K.sub.q, .sigma..sub.ys, P.sub.max, and P.sub.q
are as defined in ASTM E399-90. These invalidities are a
consequence of the high fracture toughness of the invention alloy.
A thicker and wider specimen than that permitted by the extruded
bar (1.25 inch thick.times.4 inch wide) would have been required to
obtain valid plane-strain K.sub.Ic results. A valid K.sub.Ic is
generally considered a material property relatively independent of
specimen size and geometry. K.sub.q on the other hand is not a
material property in the strict sense because it can vary with
specimen size and geometry. However, typically, K.sub.q values from
specimens smaller than that needed to obtain a valid result are
conservative with respect to K.sub.Ic, in other words, the reported
fracture toughness (K.sub.q) values are, in fact, generally lower
than the standard K.sub.Ic values obtained when sample size related
validity criteria of ASTM E399-90 are satisfied. The values of
K.sub.q were obtained using compact tension test specimen per ASTM
E399 with a thickness, B, of 1.25 inch and width that varied
between 2.5 to 3.0 inches for different specimens. The specimens
were fatigue pre-cracked to a crack length, a, of 1.2 to 1.5 inch
(a/W=0.45 to 0.5). The tests on plant trial material, discussed
below, which did satisfy the validity criterion in ASTM E399 for
K.sub.Ic, were conducted using compact tension specimens with a
thickness, B=2.0 inch and width, W=4.0 inch. These specimens were
fatigue pre-cracked to a crack length of 2.0 inch (a/W=0.5). All
cases of comparison of data between alloys were made using results
from specimens of the same size and under similar test
conditions.
Example 1
Plant Trial--Plate
[0063] A plant trial of the alloy of the present invention,
designated alloy C80A, was conducted using a standard, full-size
ingot cast with the composition (in wt. %): 7.35 Zn, 1.46 Mg, 1.64
Cu, 0.04 Fe, 0.02 Si and 0.11 Zr. The ingot was scalped,
homogenized at 885.degree. to 890.degree. F. for 24 hours, and hot
rolled to 6-inch thick plate. The rolled plate was solution heat
treated at 885.degree. to 890.degree. F. for 140 minutes, spray
quenched to ambient temperature, and cold stretched 1.5% for
residual stress relieval. Sections from the plate were given a
two-step aging comprising a 6-hour 250.degree. F. first step aging
followed by second step aging for 6, 8 and 11 hours, respectively,
at 320.degree. F. designated as times (t1), (t2) and (t3),
respectively. Results from the tensile, fracture toughness,
alternate immersion SCC, EXCO and electrical conductivity tests are
presented in Table 2 below. FIG. 7 shows the cross plot of L-T
plane-strain fracture toughness (K.sub.Ic) versus longitudinal
tensile yield strength TYS(L). A linear strength-toughness
correlation trend was possible to define through the use of data
from three different second-step aging times. Included in FIG. 7
are typical properties from 6-inch thick 7050-T7451 plates produced
by industry specification, BMS 7-323C, and the 7040-T7451 typical
values for 6-inch thick plates per AMS D99AA (Draft) specification
(ref. Preliminary Materials Properties Handbook). The C80A alloy
plate of the present invention clearly displays a much superior
strength-toughness combination compared to either prior art 7050 or
7040 alloy plate. For example, with respect to 7050-T7451 plate, a
TYS increase from about 64 ksi to 72 ksi (11% increase) is obtained
at the equivalent K.sub.Ic of about 35 ksi=29 in. Alternatively,
significant increase in KIc values is obtained at equivalent TYS.
For example, with respect to 7040-T7451 a K.sub.Ic (L-T) toughness
increase from 32.3 to about 41 ksi{square root}in (28% increase) is
obtained at the equivalent TYS (L) of 66.6 ksi.
2TABLE 2 Properties of Plant Processed 6-inch Thick C80A Plates SCC
- Stress Aging Time L-UTS L-TYS (ASTM G44) at 320.degree. F. (T/4)
(T/4) EL L-CYS L-TK.sub.Ic EXCO EC(T/4) (20d-Pass) (Hrs.) (ksi)
(ksi) (%) (ksi) (ksi{square root}in) (T/4) (% IACS) (ksi) 6 (t1)
77.1 74.9 6.8 73.2 33.6 EB 40.5 35 8 (t2) 75.6 72.5 7.3 71.0 35.2
EB 41.3 40 11 (t3) 71.9 67.2 8.6 65.6 40.5 EA 42.7 45
Example 2
Plant Trial--Forging
[0064] Die forging evaluation of the alloy of the present invention
was performed in a plant-trial using two full-size production sheet
ingots, designated C8XA-1 and C8XA-2, with the following
compositions: Ingot C8XA-1: 7.35 Zn, 1.46 Mg, 1.64 Cu, 0.038 Fe,
0.022 Si and 0.11 Zr; Ingot C8XA-2: 7.39 Zn, 1.48 Mg, 1.91 Cu,
0.036 Fe, 0.024 Si and 0.11 Zr. A standard 7050 ingot was also run
as a control. The ingots were homogenized at 885.degree. F. for 24
hours and sawed to billets for forging. A closed die forged part
was produced for evaluation of properties at three different
thicknesses, namely, 2 inch, 3 inch and 7 inch thicknesses. The
fabrication steps involved two pre-forming operations utilizing
hand forging. This was followed by a blocker die operation and a
final finish die operation both carried out in a 35,000 ton press.
Forging temperatures used were about 725.degree. to 750.degree. F.
The forged pieces were solution heat treated at 880.degree. to
890.degree. F. for 6 hours, quenched and cold worked 1 to 5% for
residual stress relief. The parts were next given a 3-step aging
treatment to develop a "T74" equivalent temper. The 7050 control
pieces were aged to "T74" temper using the standard aging of
225.degree. F. for 8 hours followed by 250.degree. F. for 8 hours
followed by 350.degree. F. for 8 hours. Mechanical property tests
were performed on samples taken from the same locations as used for
the 7050 parts using standard procedures. Results from the tensile
tests performed in longitudinal, long-transverse and
short-transverse directions are presented in FIG. 8. In all three
orientations, the tensile yield strength (TYS) for the alloy of the
invention remained virtually unchanged for thicknesses ranging from
2 to 7 inches. In contrast, the properties for 7050 showed a drop
in TYS values as the thickness increased from 2 to 3 to 7 inches
consistent with the required minimum properties for this alloy
(also shown in FIG. 8). Thus, the results in FIG. 8 demonstrate the
advantages of the low quench sensitivity properties of the alloy
forgings of the present invention in being insensitive to strength
changes over a large thickness range, in contrast to the properties
of the prior 7050 alloy.
[0065] Hence, the present invention runs counter to conventional
7XXX series aluminum alloy design which indicates that a high Mg
content is desirable for high strength. This is because a high Mg
content is good for thin sections but not in thick product. It has
been shown that more solute increases strength, but also
concurrently increases the quench sensitivity of the material. The
present invention recognizes that by lowering the Mg content to at
or less than 1.7 wt. % and by controlling the Cu and Mg
relationship to Mg.ltoreq.Cu+0.3 or Mg.ltoreq.Cu+0.3 or
Mg.ltoreq.Cu or Mg.ltoreq.Cu (contrary to the teachings of U.S.
Pat. No. 6,027,582), and by controlling the Zn content, the quench
sensitivity of the alloy is decreased by keeping the Zn, Mg and Cu
in solution at the slow cooling rate inherently present in thick
gauges, viz., 4 to 8 inches or more. Thus, the alloy of the current
invention seemingly runs counter to the historical alloy precedent
in that, instead of increasing solute levels, the overall solute
levels of Mg and Cu in the invention are actually decreased
compared to current thick section alloys.
[0066] Although the primary focus of the present invention is on
product with thick cross sections quenched as rapidly as practical,
those skilled in the art will recognize and appreciate that another
application of the invention would be use of an intentionally
slowed quench rate in parts with thin sections in order to reduce
quench induced residual stresses, and distortion brought on by
rapid quenching. Other, more quench sensitive alloys would suffer
unacceptable losses in strength, toughness and other properties
when treated this way (slowed quench rate) and would, thus, be
unsuitable.
[0067] Another variant of potential application arising from lower
quench sensitivity of the present alloy is with parts that have
both thick and thin sections. These parts will suffer less from
differences in yield strength between the thick and thin cross
sections caused by the quench rate differences. This would, in
turn, reduce the incidence of bowing or distortion in the parts
after stretching.
[0068] It is known that, for a given 7XXX series alloy, as further
artificial aging is progressively applied to the peak strength
T6-type temper condition, the strength of the alloy progressively
and systematically decreases while its fracture toughness and
corrosion resistance progressively and systematically increase.
Hence, those skilled in the art can select a specific temper
condition with an optimum combination of strength, fracture
toughness and corrosion resistance for a specific application.
Indeed, such is the case for the alloy of the invention, as
demonstrated in the cross plot of L-T plane strain fracture
toughness K.sub.Ic and L-yield strength, in FIG. 7, both measured
at quarter plane (T/4) in the longitudinal direction for a 6-inch
thick plate product. In FIG. 7, the alloy of the invention is shown
to provide the combination of about 75 ksi yield strength with
about 33 ksi{square root}in fracture toughness (at aging time t1
from Table 2), or about 72 ksi yield strength with about 38
ksi{square root}in fracture toughness (at aging time t2), or about
67 ksi yield strength with about 40 ksi{square root}in fracture
toughness (at aging time t3). It is further understood by those
skilled in the art that for a specific 7XXX series alloy, the
strength-fracture toughness trend line can be interpolated and, to
some extent, extrapolated to combinations of strength and fracture
toughness beyond the three examples of alloy C80A given above and
plotted in FIG. 7. The desired combination then can be accomplished
by an appropriately selected artificial aging treatment.
[0069] While the invention has been described largely in connection
with aerospace structural applications, such is not necessarily
intended to limit the invention which, it is believed, may find
applications as a relatively thick cast, plate, extruded or forged
product requiring a relatively high strength in a slowly quenched
condition from the solution heat treating temperature. An example
of one such application is a mold plate which is extensively
machined into molds of various shapes to facilitate shaping or
contouring processes in manufacturing processes. For this
application, the desired material characteristics are high strength
and low machining distortion. For alloys used as mold plates, a
slow quench after solution heat treatment would be necessary to
impart a low residual stress, which causes machining distortion.
However, a slow quench also results in lower strength and other
properties for existing 7XXX series alloys as a result of their
higher quench sensitivity as previously described. It is the unique
lack of quench sensitivity of the present alloy which permits a
slow quench following solution heat treatment while still retaining
the capability to attain a relatively high strength that makes it
also an attractive choice for such non-structural applications as a
thick mold plate.
[0070] Having thus described the presently preferred embodiments of
our invention, it will be understood by those skilled in the art
that the invention may be otherwise embodied or modified without
departing from the spirit and scope of the following claims.
* * * * *