U.S. patent number RE34,008 [Application Number 07/075,838] was granted by the patent office on 1992-07-28 for method of producing an aluminum alloy product.
This patent grant is currently assigned to The Boeing Company. Invention is credited to Michael V. Hyatt, William E. Quist.
United States Patent |
RE34,008 |
Quist , et al. |
July 28, 1992 |
Method of producing an aluminum alloy product
Abstract
A 7000 series aluminum alloy characterized by high strength,
high fatigue resistance and high fracture toughness consists
essentially of 5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5%
copper, 0.08 to 0.15% zirconium, a maximum of 0.15% iron, maximum
of 0.12% silicon, a maximum of 0.06% titanium, a maximum of 0.04%
chromium, a maximum of 0.05% for each of any other trace elements
present in the alloy, the total of the other trace elements in the
allow being a maximum of 0.15%, the balance of the alloy being
aluminum. The foregoing alloy is hot worked to provide a wrought
product, such as an extruded or plate product, in which
recrystallization is held to a minimum. The wrought product is
subjected to a solution treatment, quench, and elevated temperature
aging cycle, normally until the product is at or near its maximum
strength.
Inventors: |
Quist; William E. (Redmond,
WA), Hyatt; Michael V. (Bellevue, WA) |
Assignee: |
The Boeing Company (Seattle,
WA)
|
Family
ID: |
25485502 |
Appl.
No.: |
07/075,838 |
Filed: |
July 20, 1987 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
Reissue of: |
947089 |
Sep 29, 1978 |
04305763 |
Dec 15, 1981 |
|
|
Current U.S.
Class: |
148/417; 148/439;
148/552; 148/694 |
Current CPC
Class: |
C22C
21/10 (20130101); C22F 1/053 (20130101) |
Current International
Class: |
C22C
21/10 (20060101); C22F 1/053 (20060101); C22F
001/04 () |
Field of
Search: |
;148/12.7A,11.5A,2,417,439 |
References Cited
[Referenced By]
U.S. Patent Documents
Other References
Staley et al., "Further Development of Aluminum Alloy X7050, Final
Report," Naval Air Systems Command Contract N00019-71-C0131, May 8,
1972 (Navy Report III). .
Lin, Fu-Shiong, "Low Cicyle Corrosion Fatigue and Corrosion Fatigue
Crack Propagation of High Strength 7000-Type Aluminum Alloys",
Georgia Institute of Technology, May, 1978. .
Starke, E. A., "Aluminum Alloys of the 70's: Scientific Solutions
to Engineering Problems. An Invited Review", 29 Materials Science
and Engineering 99-115 (1977). .
Sanders, Jr. et al, The Effect of Intermediate Thermomechanical
Treatments on the Fatigue Properties of a 7050 Aluminum Alloy,
Metallurgical Transactions A, Vol. 9A, Aug. 1978, pp. 1087-1100.
.
"Synthesis of High-Strength Aluminum Alloys," Technical Report
AFML-TR-74-129, Part 1, dated Dec. 1974, and prepared for the Air
Force Materials Laboratory, Wright Patterson Air Force Base, by
Battelle Columbus Laboratories and the Reynolds Metal Company.
.
Thompson, David S., "Metallurgical Factors Affecting High Strength
Aluminum Alloy Production", Apr. 1975. .
Staley, J. T., "Microstructure and Toughness of High-Strength
Aluminum Alloys", Properties Related To Fracture Toughness. ASTM
STP 605. American Society For Testing and Materials, 1976, pp.
71-103..
|
Primary Examiner: Dean; R.
Assistant Examiner: Phipps; Margery S.
Attorney, Agent or Firm: Finnegan, Henderson, Farabow,
Garrett & Dunner
Claims
The embodiments of the invention in which an exclusive property or
privilege is claimed are defined as follows:
1. .[.A method of producing.]. .Iadd.An upper wing skin of a
commercial jet aircraft comprising .Iaddend.an improved alloy
product .Iadd.having high strength, high fracture toughness and
high fatigue resistance made in accordance with the process
.Iaddend.comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body .Iadd.to a thickness of less than about 1.5
inches .Iaddend.to provide a wrought product, said .[.alloy.].
.Iadd.body .Iaddend.being hot worked .[.so as to prevent
substantial recrystallization of said alloy.]. .Iadd.by rolling at
a temperature that is intentionally maintained sufficiently high so
that less than about 50 volume % of said alloy is recrystallized
product.Iaddend.,
(c) subjecting said product to solution treatment and quenching,
.Iadd.and .Iaddend.
(d) subjecting said product to an artificial aging treatment at an
elevated temperature .Iadd.wherein said artificial aging treatment
is continued only until said alloy reaches its peak
strength.Iaddend..
2. .[.The method of claim 1.]. .Iadd.An upper wing skin of a
commercial jet aircraft comprising an improved alloy product having
high strength, high fracture toughness and high fatigue resistance
made in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 30 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature .Iaddend.wherein said artificial aging
treatment is continued only until said alloy reaches its peak
strength.
3. .[.The method of claim 1.]. .Iadd.An upper wing skin of a
commercial jet aircraft comprising an improved alloy product having
high strength, high fracture toughness and high fatigue resistance
made in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 50 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature .Iaddend.wherein said artificial aging
treatment comprises:
first aging said product at an intermediate temperature above room
temperature and below said elevated temperature, and
thereafter aging said product at said elevated temperature until
said alloy reaches it peak strength.
4. .[.The method of claim 3.]. .Iadd.An upper wing skin of a
commercial jet aircraft comprising an improved alloy product having
high strength, high fracture toughness and high fatigue resistance
made in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 50 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said project to an artificial aging treatment at an
elevated temperature wherein said artificial aging treatment
comprises:
first aging said product at an intermediate temperature above room
temperature and below said elevated temperature, and
thereafter aging said product at said elevated temperature until
said alloy reaches its peak strength .Iaddend.wherein the second
aging step comprises:
aging said product at said elevated temperature T of from
300.degree. F. to 340.degree. F. for about the period of time
(t.sub.T) indicated by the following formula: ##EQU3## wherein Y is
a factor read from the graph of FIG. 1 for a desired aging
temperature T, wherein t.sub.325 can range from about 3 to 5 hours,
and wherein t.sub.T can be varied up to about .+-.3 hours from the
value calculated by the formula.
5. The method of claim 1.]. .Iadd.An upper wing skin of a
commercial jet aircraft comprising an improved alloy product having
high strength, high fracture toughness and high fatigue resistance
made in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 50 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature .Iaddend.wherein said artificial aging step
comprises:
initially aging said product for a period of from 4 to 48 hours at
a temperature of from 225.degree. F. to 275.degree. F., and
thereafter aging said product for a period of from 3 to 12 hours at
an elevated temperature of from 310.degree. F. to 325.degree. F.
.[.
6. The method of claims 1, 2, 3, 4 or 5 wherein said alloy is hot
worked at a temperature sufficiently high so that less than about
50% of said alloy is recrystallized..]. .[.7. The method of claims
1, 2, 3, 4 or 5 wherein said alloy is hot worked at a temperature
sufficiently high so that less than about 30% of said alloy is
recrystallized..]. .[.8. The method of claim 1 wherein said
artificial aging treatment is continued after said alloy reaches
peak strength to enhance the corrosion resistance properties of
said alloy..]. .[.9. The product produced by the method of claim
1..]. .[.10. The product produced by the method of claim 2..].
.[.11. The product produced by the method of claim 3..]. .[.12. The
product produced by the method of claim 4..]. .[.13. The product
produced by the method of claim 5..]. .[.14. The product produced
by the method of claim 6..]. .[.15. The product produced by the
method of claim 7..].
.[.16. The product produced by the method of claim 8..]. .Iadd.17.
An upper wing skin of a commercial jet aircraft comprising an
improved alloy product having high strength, high fracture
toughness and high fatigue resistance made in accordance with the
process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 50 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature, to thereby produce a product having a minimum
compression yield strength of 76 ksi, an average fracture toughness
of 70 ksi.sqroot.in., and a fatigue crack growth rate of 7.3
microinches per cycle at an average cyclic stress intensity
.DELTA.K of 11 ksi.sqroot.in.
at R=0.06. .Iaddend. .Iadd.18. An upper wing skin of a commercial
jet aircraft comprising an improved alloy product having high
strength, high fracture toughness and high fatigue resistance made
in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 30 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated
temperature. .Iaddend. .Iadd.19. An upper wing skin of a commercial
jet aircraft comprising an improved alloy product having high
strength, high fracture toughness and high fatigue resistance made
in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 30 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature wherein said artificial aging treatment
comprises:
first aging said product at an intermediate temperature above room
temperature and below said elevated temperature, and
thereafter aging said product at said elevated temperature until
said alloy
reaches its peak strength. .Iaddend. .Iadd.20. An upper wing skin
of a commercial jet aircraft comprising an improved alloy product
having high strength, high fracture toughness and high fatigue
resistance made in accordance with the process comprising the steps
of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 30 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature wherein said artificial aging treatment
comprises:
first aging said product at an intermediate temperature above room
temperature and below said elevated temperature, and
thereafter aging said product at said elevated temperature until
said alloy reaches its peak strength wherein the second aging step
comprises:
aging said product at said elevated temperature T of from
300.degree. F. to 340.degree. F. for about the period of time
(t.sub.T) indicated by the following formula: ##EQU4## .Iaddend.
wherein Y is a factor read from the graph of FIG. 1 for a desired
aging temperature T, wherein t.sub.325 can range from about 3 to 5
hours, and wherein t.sub.T can be varied up to about .+-.3 hours
from the value
calculated by the formula. .Iadd.21. An upper wing skin of a
commercial jet aircraft comprising an improved alloy product having
high strength, high fracture toughness and high fatigue resistance
made in accordance with the process comprising the steps of:
(a) providing a body composed of an alloy consisting essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, a maximum of 0.15% iron, a maximum of 0.12%
silicon, a maximum of 0.06% titanium, a maximum of 0.04% chromium,
a maximum of 0.05% for each of any other trace elements present in
the alloy, the maximum total of said other trace elements being
0.15%, the balance being aluminum, all percentages herein by weight
based on the total alloy,
(b) working said body to a thickness of less than about 1.5 inches
to provide a wrought product, said body being hot worked by rolling
at a temperature that is intentionally maintained sufficiently high
so that less than about 30 volume % of said alloy is recrystallized
product,
(c) subjecting said product to solution treatment and quenching,
and
(d) subjecting said product to an artificial aging treatment at an
elevated temperature wherein said artificial aging step
comprises:
initially aging said product for a period of from 4 to 48 hours at
a temperature of from 225.degree. F. to 275.degree. F., and
thereafter aging said product for a period of from 3 to 12 hours at
an elevated temperature of from 310.degree. F. to 325.degree. F.
.Iaddend.
Description
BACKGROUND OF THE INVENTION
The present invention relates to aluminum alloys, and more
particularly to a 7000 series alloy of the
aluminum-zinc-magnesium-copper type characterized by high strength,
high fatigue properties and high fracture toughness.
A significant economic factor in operating aircraft today is the
cost of fuel. As a consequence, aircraft designers and
manufacturers are constantly striving to improve the overall fuel
efficiency. One way to increase this fuel efficiency, as well as
overall performance, is to reduce structural weight. Since aluminum
alloys are used in a large proportion of the structural components
of most aircraft, significant efforts have been expended to develop
aluminum alloys that have higher strength to weight ratios than the
alloys in current use, while maintaining the same or higher
fracture toughness, fatigue resistance and corrosion
resistance.
For example, one alloy currently used on the upper wing skins of
some commercial jet aircraft is alloy 7075 in the T651 temper.
Alloy 7075-T651 has a high strength to weight ratio, while
exhibiting good fracture toughness, good fatigue properties, and
adequate corrosion resistance. Another currently available alloy
sometimes used on commercial jet aircraft, alloy 7178-T651, is
stronger than 7075-T651; however, alloy 7178-T651 is inferior to
alloy 7075-T651 in fracture toughness and fatigue resistance. Thus
there are more restrictions to taking advantage of the higher
strength to weight ratio of alloy 7178-T651 without sacrificing
fracture toughness and/or fatigue performance of the component on
which it is desired to use the alloy. Other currently available
alloys and tempers, although sometimes exhibiting good toughness
properties and high resistance to stress-corrosion cracking and
exfoliation corrosion, offer no strength advantage over alloy
7075-T651. Examples of such alloys are 7475-T651, T7651 and T7351
and 7050-T7651 and T73651. Thus, with currently available alloys
and tempers, it is impossible to achieve a weight saving in
aircraft structural components while maintaining fracture
toughness, fatigue resistance and corrosion resistance at or above
the level currently available with alloy 7075-T651.
It is therefore an object of the present invention to provide an
aluminum alloy for use in structural components of aircraft that
has a higher strength to weight ratio than the currently available
alloy 7075-T651. It is a further object of the present invention to
provide such an alloy that exhibits improved fatigue and fracture
toughness properties while maintaining stress-corrosion resistance
and exfoliation corrosion resistance at a level approximately
equivalent to that of alloy 7075-T651.
SUMMARY OF THE INVENTION
The 7000 series alloy of the present invention fulfills the
foregoing objects by providing a strength increase of from 10 to
15% over alloy 7075 in T6 tempers. Indeed, the alloy of the present
invention is stronger than any other commercially available
aluminum alloy. At the same time, the fracture toughness and
fatigue resistance of the aluminum alloy of the present invention
are higher than that achievable in alloys having strengths
approaching that of the alloy of the present invention, such as
7075 and 7178 in the T6 tempers. Additionally, the corrosion
resistance of the alloy of the present invention is approximately
equivalent to that exhibited by alloy 7075 in the T6 tempers.
The desired combination of properties of the aluminum alloy of the
present invention has been achieved in a 7000 series alloy by
precisely controlling the chemical composition ranges of the
alloying and trace elements, by heat treating the alloy to increase
its strength to high levels, and by maintaining a substantially
unrecrystallized microstructure. The alloy of the present invention
consists essentially of 5.9 to 6.9% zinc, 2.0 to 2.7% magnesium,
1.9 to 2.5% copper, 0.08 to 0.15% zirconium, a maximum of 0.15%
iron, a maximum of 0.12% silicon, a maximum of 0.06% titanium, a
maximum of 0.04% chromium, and a maximum of 0.05% for other trace
elements present in the alloy, the total of the other trace
elements being a maximum of 0.15%, the balance of the alloy being
aluminum. Once the alloy is cast, it is hot worked to provide a
wrought product, such as extrusions or plate. The product is then
solution treated, quenched and subjected to an artificial aging
treatment at an elevated temperature. To achieve the high strength
requirements, the invention alloy is aged at elevated temperatures
until it reaches its peak strength condition. The resulting product
exhibits a strength increase of 10% to 15% over that exhibited by
commercially available alloys such as 7075-T651 and 7050-T7651.
Also, by hot working the alloy when forming the product so as to
prevent any substantial recrystallization in the final product, the
fracture toughness of the alloy of the present invention can be
maintained at a level approximately 10% higher than that of alloy
7075-T651 and substantially above that of alloy 7178-T651.
BRIEF DESCRIPTION OF THE DRAWINGS
A better understanding of the present invention can be derived by
reading the ensuing description in conjunction with the
accompanying drawings wherein:
FIG. 1 is a graph of a correction factor (Y) versus aging
temperature used to determine equivalent heat treatment times for
the invention alloy;
FIG. 2 shows bar graphs comparing the properties of the alloy of
the present invention with prior art 7000 series aluminum
alloys;
FIG. 3 shows graphs of strength versus aging time for the invention
alloy and other 7000 series aluminum alloys;
FIG. 4 shows graphs of the fracture toughness parameter (K.sub.app)
versus thickness comparing the invention alloy with prior art 7000
series aluminum alloys;
FIG. 5 shows graphs of fatigue crack growth rate (da/dN) versus
cyclic stress intensity factor (.DELTA.K) comparing the invention
alloy with prior art 7000 series alloys; and
FIG. 6 shows graphs of fatigue crack length versus stress cycles
comparing the invention alloy with prior art 7000 series
alloys.
DETAILED DESCRIPTION OF THE INVENTION
The high strength, high fatigue resistance, high fracture toughness
and corrosion resistance properties of the alloy of the present
invention are dependent upon a chemical composition that is closely
controlled within specific limits as set forth below, a carefully
controlled heat treatment of products made from the alloy, and a
microstructure that is substantially unrecrystallized. If the
composition, fabrication, and heat treatment parameters of the
invention alloy stray from the limits set forth below, the desired
combination of strength increase, fracture toughness increase and
fatigue improvement objectives will not be achieved.
The aluminum alloy of the present invention consists essentially of
5.9 to 6.9% zinc, 2.0 to 2.7% magnesium, 1.9 to 2.5% copper, 0.08
to 0.15% zirconium, the balance being aluminum and trace elements.
Of the trace elements present, the maximum percentage of iron
allowable is 0.15%, of silicon allowable is 0.12%, of manganese
allowable is 0.10%, of chromium allowable is 0.04%, and of titanium
allowable is 0.06%. Any other remaining trace elements have maximum
limits of 0.05%, with a maximum total for the remaining trace
elements being 0.15%. (The foregoing percentages are weight
percentages based on the total alloy.) The most critical of the
trace elements present are normally iron and silicon. If iron and
silicon are present in the alloy in excess of the amounts stated
above, the undesirable intermetallic compounds formed by iron and
silicon during solidification, fabrication, and heat treatment will
lower the fracture toughness properties of the alloy of the present
invention to unacceptable levels.
The high zinc, magnesium and copper contents of the alloy of the
present invention are major contributors to the high strength
characteristics of the present alloy. If the zinc, magnesium and
copper contents are below the limits set forth above, the strength
of the alloy will fall below the strength objectives of a 10% to
15% increase over that of the base line standard, alloy
7075-T651.
Conventional melting and casting procedures are employed to
formulate the alloy. Care must be taken, as pointed out above, to
maintain high purity in the aluminum and the alloying constituents
so that the trace elements, and especially iron and silicon, are
maintained below the requisite maximums. Ingots are produced from
the alloy using conventional procedures such as continuous direct
chill casting. Once the ingot is formed, it can be homogenized by
conventional techniques, for example, subjecting the ingot to
elevated temperatures of about 900.degree. F. for a period of time
sufficient to homogenize the internal structure of the ingot and to
provide an essentially uniform distribution of the alloying
elements. The ingot can then be subjected to hot working procedures
to produce a desired product such as plate or extrusions. When
fabricating products from the alloy of the present invention, no
unusual metallurgical procedures are required. However, in order to
maintain the combination of mechanical and fracture properties of
the alloy of the present invention, it is important to hot roll,
extrude, or otherwise work produces of the alloy in a manner that
avoids excessive recrystallization of the microstructure of the
final product. Avoiding hot working (or cold working) practices
which lead to significant amounts of recrystallization is critical,
particularly for thinner plate and extrusions, for which there is
an increased tendency for recrystallization to occur during
solution treatment. Therefore, the product formed from an alloy of
the present invention must be substantially unrecrystallized. By
"substantially unrecrystallized" it is meant that less than about
50 volume percent of the alloy microstructure in a given product is
in a recrystallized form, excepting surface layers which often show
a much higher degree of recrystallization. (The surface layers of
plate and extrusion products are usually removed during fabrication
into final part configurations). Most preferably, it is desired to
maintain the volume percent of recrystallized microstructure less
than about 30%. Recrystallization can be minimized by maintaining
the temperature during hot working at levels that cause annealing
out of internal strains produced by the working operation such that
recrystallization will be minimized during the working operation
itself, or during subsequent solution treatment. For example, hot
rolling a plate product produced from the alloy of the present
invention to a thickness on the order of 1 inch at a metal
temperature of about 800.degree. F. will ordinarily prevent
substantial recrystallization. Under a given set of conditions in a
production rolling operation, it may be possible to roll at lower
temperatures and still prevent substantial recrystallization. It
has been found, for example, that the fracture toughness of an
alloy having a microstructure that is greater than about 50%
recrystallized deteriorates drastically, and in fact can fall
considerably below the fracture toughness of prior art alloys such
as 7075-T651.
After the alloy is hot worked into a product, the product is
typically solution heat treated at a temperature on the order of
890.degree. F., and preferably between 890.degree. F. and
900.degree. F. for a time sufficient for solution effects to
approach equilibrium. Once the solution effects have approached
equilibrium, the product is quenched, normally by spraying the
product with, or immersing the product in, room temperature water.
Thereafter the product is stretched 1% to 3% in the rolling or
extrusion direction to eliminate residual quenching stresses.
It should be noted at this point that the tensile strength of the
alloy of the present invention is relatively insensitive to quench
rate. Thus its superior strength levels are maintained in both
plate and extrusions of substantial thickness. This property of the
alloy of the present invention results from the use of zirconium
instead of chromium as the grain refining element. Chromium is used
for most other 7000 series alloys and results in substantial
decreases in strength for section thicknesses over about 3 inches,
whereas the alloy of the present invention decreases only
moderately in strength even when produced in section thicknesses
well over 3 inches.
Although the high zinc, magnesium and copper content of the alloy
of the present invention is required to obtain its superior
strength characteristics, it is also necessary to artificially age
the product formed from the alloy at an elevated temperature until
the superior strength characteristics are achieved. In accordance
with the present invention, the presently preferred method to
artificially age the product produced from the alloy of the present
invention is to use a two step aging procedure. The alloy is
preferably first aged at an intermediate temperature on the order
of 250.degree. F. for a period of from about 4 to about 48 hours.
It should be noted that the first aging step can be modified or
even possibly eliminated. For example, data accumulated to date
indicates that the alloy can be aged during the first stage at
temperatures ranging from 225.degree. F. to 275.degree. F.
The second stage aging treatment is conducted at a temperature that
is above the aging temperature employed during the first stage. The
second stage aging is preferably conducted in the range of from
310.degree. F. to 325.degree. F. until the alloy reaches peak
strength. By peak strength it is meant a strength at or near the
maximum strength of the alloy. For example, if the second stage
aging is conducted at 325.degree. F., the aging time will range
from about 3 to about 5 hours. If the second stage aging is
conducted at 310.degree. F., the aging time will range from about 6
to about 12 hours.
If desired, the second stage aging can also be conducted at
temperatures in an expanded range of from 300.degree. F. to
340.degree. F. until peak strength is achieved. However, for
temperatures at the lower end of the foregoing range, the aging
time must be adjusted upwardly and for temperatures toward the
upper end of the foregoing range, the aging time must be adjusted
downwardly. The formula below may be used to determine the
preferred second stage aging time (t.sub.T) for aging temperatures
other than 325.degree. F. This formula will provide an aging time
for a given temperature within the range of 300.degree. F. to
340.degree. F. that is equivalent to the second stage aging time
for the aging temperature of 325.degree. F. as set forth in the
preceding paragraph. The formula is: ##EQU1## wherein t.sub.T is
the time for which the product of the present invention is aged
during the second stage aging at a temperature T other than
325.degree. F. to achieve peak strength,
wherein t.sub.325 can range from about 3 to about 5 hours for
various products as set forth in the preceding paragraph, and
wherein Y is a factor for converting the 325.degree. F. aging time
(t.sub.325) to the aging time t.sub.T at the temperature T.
The factor Y is derived from the graph of FIG. 1 which is a
log-linear graph of the Y factor versus aging temperature. For
example, if it were desired to conduct the second stage aging at a
temperature of 312.degree. F., the factor Y would be about 0.5; and
if it were desired to age at a temperature of 338.degree. F., the
factor Y would be about 2. It should also be realized that the
aging time (t.sub.T) calculated from the above formula can be
varied up to about 3 hours and still achieve the peak strength
properties in accordance with the present invention. For example,
for second stage aging temperatures near the upper limit of the
expanded range, the variation from t.sub.T is preferably no more
than about .+-.1/2 hour; however, at the lower end of the expanded
range, t.sub.T can be varied up to about .+-.3 hours.
EXAMPLES
The following Examples are intended to be illustrative of the
present invention and are intended to teach one of ordinary skill
how to make and use the invention. They are not intended in any way
to delimit or otherwise narrow the scope of protection afforded by
the grant of Letters Patent hereon.
EXAMPLE I
More than fifty ingots of the alloy of the present invention were
formulated in accordance with conventional procedures. These ingots
had a nominal composition of 6.4% zinc, 2.35% magnesium, 2.2%
copper, 0.11% zirconium, 0.07% iron, 0.05% silicon, <0.01%
manganese, 0.01% chromium, 0.02% titanium, and a total of <0.03%
of other trace elements, the balance of the alloy being aluminum.
The ingots were rectangular in shape and had thicknesses between
sixteen and twenty-four inches. The ingots were scalped,
homogenized at about 880.degree. F., and hot rolled to plate
thicknesses varying from 0.375 to about 1.5 inches. These plates
were then solution heat treated at about 890.degree. F. for 1 to 2
hours, depending on thickness, and spray quenched in room
temperature water. The plates were then stretched 11/2 to 3% in the
rolling direction to eliminate residual quenching stresses and were
artificially aged for 24 hours at 250.degree. F., followed by a
second stage aging at about 310.degree. F. for about 11 to 12
hours. Compression yield strength, fracture toughness and fatigue
crack growth rate tests were then run on specimens taken from the
plate products. The data from these tests were analyzed to provide
minimum and mean values for each of the tests.
Similar data from conventional commercially available 7075-T651
alloy, 7178-T651 alloy and 7050-T7651 alloy plate were also
analyzed for comparison. The 7075 alloy had a nominal composition
of 5.6% zinc, 2.5% magnesium, 1.6% copper, 0.2% chromium, 0.05%
manganese, 0.2% iron and 0.15% silicon, the balance of the alloy
being aluminum and small amounts of other extraneous elements. The
7178 alloy had a nominal composition of 6.8% zinc, 2.7% magnesium,
2.0% copper, 0.2% chromium, 0.05% manganese, 0.2% iron and 0.15%
silicon, the balance of the alloy being aluminum and small amounts
of other extraneous elements. The 7050 alloy had a nominal
composition of 6.2% zinc, 2.25% magnesium, 2.3% copper, 0.12%
zirconium, 0.09% iron, 0.07% silicon, 0.01% chromium, 0.02%
titanium, the balance of the alloy being aluminum and small amounts
of other extraneous elements.
Compression yield strength (F.sub.cy) tests were run in a
conventional manner. The fracture toughness tests were also run in
a conventional manner at room temperature using center cracked
panels, with the data being represented in terms of the apparent
critical stress intensity factor K.sub.app at panel fracture. The
fracture toughness parameter (K.sub.app) is related to the stress
required to fracture a flat panel containing a crack oriented
normal to the stressing direction and is determined from the
following formula: ##EQU2## wherein .sigma..sub.g is the gross
stress required to fracture the panel;
a.sub.o is one-half the initial crack length for a center cracked
panel, and
.alpha. is a finite width correction factor (for the panels tested,
.alpha. was slightly greater than 1).
For the present tests, 16 inch wide to 20 inch wide panels
containing center cracks approximately one-third the panel width
were used to obtain the K.sub.app values.
The data for the fatigue crack growth rate comparisons were taken
from data developed from precracked, single edge notched panels.
The panels were cylically stressed in laboratory air in a direction
normal to the orientation of the fatigue crack. The minimum to
maximum stress ratio (R) for these tests was 0.06. Fatigue crack
growth rates (da/dN) were determined as a function of the cyclic
stress intensity parameter (.DELTA.K) applied to the precracked
specimens. The parameter .DELTA.K (ksi.sqroot.in) is a function of
the cyclic fatigue stress (.DELTA..sigma.) applied to the panel,
the stress ratio (R), the crack length and the panel dimensions.
Fatigue comparisons were made by noting the cyclic stress intensity
(.DELTA.K) required to propagate the fatigue crack at a rate of 7.3
microinches/cycle for each of the alloys.
The results of the strength, fracture toughness and fatigue crack
growth rate tests are set forth in the bar graphs of FIG. 2 as
percentage changes from the base line alloy 7075-T651, which was
chosen for comparison as it is currently used for many aircraft
applications including upper wing surfaces. The values for the
minimum compression yield strength (99% of the test specimens meet
or exceed the value shown with a 95% confidence level), and the
average K.sub.app are set forth at the top of the appropriate bar
in FIG. 2. Fatigue crack growth rate behavior is expressed as a
percentage difference between the average cyclic stress intensity
(.DELTA.K) required for a crack growth rate of 7.3
microinches/cycle for a given alloy and the .DELTA.K required for a
crack growth rate of 7.3 microinches/cycle in 7075-T651. As can be
seen from FIG. 2, the .DELTA.K level required to provide a crack
growth rate of 7.3 microinches/cycle for the 7075-T651 alloy was
about 10 ksi.sqroot.in.; for the alloy of the present invention, 11
ksi.sqroot.in.; for the 7178 alloy 8.2 ksi.sqroot.in.; and for the
7050 alloy, 11 ksi.sqroot.in.
The bar graphs in FIG. 2 show that the alloy of the present
invention has strength, fracture toughness and fatigue properties
that are 10 to 15% better than the 7075-T651 base line alloy. As
can be seen, the 7050-T7651 alloy has fracture toughness and
fatigue properties similar to that of the invention alloy, however,
the compression yield strength of the 7050-T7651 alloy is not only
below that of the alloy of the present invention but is also
slightly below that of the base line alloy 7075-T651. As is readily
observed, the fracture toughness and fatigue crack growth rate
properties of the invention alloy are substantially improved over
those of the 7178-T651 alloy. Thus it is observed that only by
staying within the compositional limits of the alloy of the present
invention, by carefully hot working the alloy of the present
invention to prevent substantial recrystallization, and by aging
the alloy of the present invention to its peak strength can all
three of the strength, fracture toughness and fatigue properties be
improved over that of the base line alloy 7075-T651. Although not
noted in the above comparisons or in the data of FIG. 2, it should
also be emphasized that comparisons for extruded products show
similar relative improvements for the invention alloy over prior
art alloys.
EXAMPLE II
The procedures of Example I were employed to produce a plate and
extrusion product from typical ingots of the alloy of the present
invention. After initially artificially aging the products for
about 24 hours at about 250.degree. F., the products produced from
the alloy of the present invention were subjected to a second stage
aging step at 325.degree. F. for varying amounts of time ranging
from 0 to 24 hours. The alloys had the same nominal composition as
the alloys of the present invention shown in Example I. Specimens
taken from the products were then tested for longitudinal yield
strength using conventional procedures. The resulting typical yield
strengths versus aging time are plotted in graphs A and B of FIG.
3. Graph A indicates the strength values obtained from the extruded
product and graph B indicates the strength values obtained from the
plate product. Additionally, typical yield strengths from plate
products of conventional 7178-T651 and 7075-T651 alloys subjected
to a second stage aging at 325.degree. F. for various times ranging
from 0 to 24 hours are shown. The strength values for the 7178
plate were shown in graph C, and the strength values for the 7075
plate are shown in graph D of FIG. 3.
It will be noted from FIG. 3 that the invention alloy achieves and
maintains peak strength after additional aging at 325.degree. F.
for about 3 to 5 hours. To the contrary, as the 7075 and 7178
plates are exposed to the 325.degree. F. second stage aging
treatment, their strength immediately begins to decrease. It is
also observed that when the alloy of the present invention is
overaged significantly, on the order of 15 to 25 hours, its
strength falls below its peak or maximum strength. At these
significantly overaged tempers, however, the alloy of the present
invention shows significant improvements in short transverse
stress-corrosion resistance and exfoliation corrosion
resistance.
EXAMPLE III
Conventional fracture toughness tests were conducted on center
cracked test panels from the alloy of the present invention
produced in accordance with the procedure set forth in Example I,
and also from alloys 7075-T651 and 7178-T651. The test panels had
varying thicknesses and were machined from 0.5 inch and 1.0 inch
thick plate produced from the alloys. The nominal composition of
the alloy of the present invention, and of 7075 and 7178, were the
same as those shown in Example I. The fracture toughness data
(K.sub.app) from several tests at room temperature were averaged
and are plotted versus panel thickness in FIG. 4. The fracture
toughness for the product produced from the alloy of the present
invention is shown by graph E of FIG. 4, the fracture toughness for
the 7075-T651 alloy by graph F, and the fracture toughness of the
7178-T651 alloy of graph C. As will be observed, the alloy of the
present invention exhibits better fracture toughness than alloy
7075-T651 and much improved toughness compared to alloy
7178-T651.
Additionally, an alloy having the composition of the alloy of the
present invention were formed into plate products of varying
thickness in accordance with the procedure set forth in Example I,
with the exception that the hot working temperatures were not
sufficiently high to prevent excessive recrystallization in the
plate products. It was determined that approximately 75 volume
percent of the alloy was recrystallized. The room temperature
fracture toughness data for these substantially recrystallized
plates of the alloy are plotted versus plate thickness in graph H
of FIG. 4. As will be observed, the fracture toughness properties
of the invention alloy, when substantially recrystallized, fall to
approximately the levels of the 7178-T651 alloy. As a consequence,
it is important that the alloy of the present invention be hot
worked in a manner that will prevent substantial recrystallization.
The volume percent recrystallized was determined for this Example
by the point count method on photomicrographs (100.times.
magnification) of a full thickness sample. For purposes of
comparison, the alloy of the present invention for which fracture
toughness data is presented in graph E of FIG. 4 was only about 17%
recrystallized, while the alloy for which fracture toughness data
is presented in graph H was about 75% recrystallized. From this, it
is apparent that an alloy of the present invention must be
substantially unrecrystallized in order to provide fracture
toughness properties that are better than the prior art alloys.
EXAMPLE IV
The fatigue crack growth rate (da/dN) properties of the alloy of
the present invention are improved over other commercial alloys
having similar strength characteristics, namely the 7075-T651 and
7178-T651 alloys. Four production lots of plate material of the
alloy of the present invention were prepared in accordance with the
general procedure set forth in Example I. In addition, nine
production lots of 7075-T651 alloy plate and two production lots of
7178-T651 alloy plate were procured. Using the general procedures
outlined in Example I, fatigue crack growth rate tests were
conducted on precracked single edge notched panels produced from
the production lots on each of the alloys. For the alloy of the
present invention, eight da/dN tests were run; for the 7075-T651
alloy, nine da/dN tests were run; and for the 7178-T651 alloy,
eight da/dN tests were run. The da/dN values for the various alloys
were then averaged and plotted. FIG. 5 is a plot of the mean values
of the crack growth rates (da/dN) in microinches per cycle versus
the cyclic stress intensity parameter (.DELTA. K) for each of the
alloys. Curve I represents the crack growth rates for 7178-T651
alloy, curve J for 7075-T651 alloy, and curve K for the alloy of
the present invention. As is readily observed from the graphs of
FIG. 5, the alloy of the present invention has superior fatigue
crack growth rate properties at each stress intensity level
examined when compared with the 7178-T651 and 7075-T651 alloys.
The data from FIG. 5 were utilized to plot the graphs of FIG. 6
wherein crack length is plotted versus the number of stress cycles,
wherein the maximum stress applied was selected to be 10,000 psi
and wherein the minimum to maximum stress ratio (R) was equal to
0.06. The initial crack length in the panels was selected to be
0.45 inches. Curve L is the graph of the data for the 7178-T651
alloy, curve M for the 7075-T651 alloy and curve N for the alloy of
the present invention. Again, the graphs of FIG. 6 clearly
illustrate that the alloy of the present invention outperforms
alloys 7178-T651 and 7075-T651 in crack growth rate properties by
substantial margins.
As can be readily observed by reference to the foregoing Examples,
the alloy of the present invention has a superior combination of
strength, fracture toughness and fatigue resistance when compared
to the prior art alloys typified by 7075-T651, 7178-T651 and
7050-T7651. Other tests conducted on the alloy of the present
invention and comparable 7075-T651 and 7178-T651 alloys also
indicate that the stress corrosion resistance and exfoliation
corrosion resistance of the alloy of the present invention are
approximately equivalent to the corrosion resistance properties of
alloy 7075-T651, and thus can be employed for the same
applications, such as wing panels and the like.
Accordingly, one of ordinary skill, after reading the foregoing
specification, will be able to effect various changes,
substitutions of equivalents and other alterations to the
compositions and procedures set forth without varying from the
general concepts disclosed. It is therefore intended that a grant
of Letters Patent hereon be limited only by the definition
contained in the appended claims and equivalents thereof.
* * * * *