Aluminum-copper-magnesium-manganese alloy useful for aircraft applications

A product comprising an aluminum base alloy including about 3.8 wt. % copper, about 1.2 wt. % magnesium, about 0.3 to 0.6 wt. % manganese, not more than about 0.15 wt. % silicon, not more than about 0.12 wt. % iron, not more than about 0.1 wt. % titanium, the remainder substantially aluminum, incidental elements and impurities, the product having at least 5% improvement over 2024 alloy in fracture toughness, fatigue crack growth rate, corrosion resistance, and formability properties.

BACKGROUND OF THE INVENTION 
1. Field of Invention 
This invention relates to aluminum alloys suitable for use in aircraft 
applications. More specifically, it relates to a method of making an 
improved aluminum product having improved damage tolerant characteristics, 
including improved fracture toughness, fatigue resistance, corrosion 
resistance, formability and surface roughness properties. 
2. Description of the Related Art 
The design of commercial aircraft requires different sets of properties for 
different types of structures. Depending on the design criteria for a 
particular airplane component, improvements in fracture toughness and 
fatigue resistance result in weight savings, which translate to fuel 
economy over the lifetime of the aircraft, and/or a greater level of 
safety. For example, a slower fatigue crack growth rate will require a 
longer time for a crack or flaw to grow to a size where it becomes 
"critical" leading to catastrophic failure; and higher fracture toughness 
means that a crack can grow to a longer length before it is critical. 
Corrosion damage has been a perennial problem in today's aircraft, and the 
fuselage is the prime location for corrosion to occur. Improvements in 
corrosion resistance, therefore, are often sought with or without weight 
savings. 
The issues of toughness, fatigue and corrosion all relate to structural 
integrity of the airplane. Formability, on the other hand, is a 
manufacturing concern. Sheets that are easier to bend, stretch and draw 
into various shapes reduce the cost of building an airplane because the 
scrap rate is reduced, and less labor is required. 
For some time, heat treatable aluminum base alloy sheet and plate 
containing copper, magnesium and manganese has found considerable 
acceptance for various structural members. Such alloys generally contain 
3.8 to 4.9 wt. % copper, 1.2 to 1.8 wt. % magnesium and 0.3 to 0.9 wt. % 
manganese and carries the Aluminum Association designation of 2024 alloy. 
This alloy is noted for its superior strength to weight ratio, its good 
toughness and tear resistance, and adequate resistance to general and 
stress corrosion effects. 
Workers in the field have generally adapted the 2024 alloy for use in the 
construction of commercial aircraft. For example, one alloy used on the 
lower wing skins of some commercial jet aircraft is alloy 2024 in the T351 
temper. Alloy 2024-T351 has a relatively high strength-to-density ratio 
and exhibits reasonably good fracture toughness, good fatigue properties, 
and adequate corrosion resistance. U.S. Pat. Nos. 4,336,075 to Quist et 
al. and 4,294,625 to Hyatt et al. disclose an alloy which has a higher 
strength to density ratio, improved fatigue and fracture toughness 
characteristics over alloy 2024 while maintaining corrosion resistance 
levels approximately equal to or slightly better than 2024. Quist et al. 
and Hyatt et al. achieve their improvements by homogenizing the alloy at a 
moderate temperature, carefully controlling the hot-rolling and extrusion 
parameters and then natural age-hardening to produce a highly elongated, 
substantially unrecrystallized microstructure. Similarly, U.S. Pat. No. 
5,213,639 to Colvin et al. discloses an alloy which has at least a 5.0% 
improvement over 2024 alloy in T-L fracture toughness or fatigue crack 
growth rate by re-heating the alloy prior to hot rolling. 
There still remains a need, however, for an improved alloy that has 
increased fracture toughness, fatigue resistance, corrosion resistance and 
formability over alloy 2024. Accordingly, it is the primary object of this 
invention to provide such an alloy. 
SUMMARY OF THE INVENTION 
The present invention provides a product comprising an aluminum base alloy 
including about 3.8 to 4.5 wt. % copper, about 1.2 to 1.6 wt. % magnesium, 
about 0.3 to 0.6 wt. % manganese, not more than about 0.15 wt. % silicon, 
not more than about 0.12 wt. % iron, not more than about 0.1 wt. % 
titanium, the remainder substantially aluminum, incidental elements and 
impurities, the product having at least 5% improvement over 2024 alloy in 
fracture toughness, fatigue crack growth rate, corrosion resistance, 
formability properties. 
In an alternative embodiment, the invention provides a method of producing 
an aluminum product comprising providing stock including an aluminum alloy 
comprising about 3.8 to 4.9 wt. % copper, about 1.2 to 1.8 wt. % 
magnesium, about 0.3 to 0.9 wt. % manganese, not more than 0.30 wt. % 
silicon, not more than 0.30 wt. % iron, not more than 0.15 wt. % titanium, 
the remainder substantially aluminum, incidental elements and impurities; 
hot working the stock; annealing; cold rolling; solution heat treating; 
and cooling thereby producing an alloy having improved fracture toughness, 
fatigue resistance, corrosion resistance, and formability properties. 
The foregoing and other objects, features, and advantages of the invention 
will become more readily apparent from the following detailed description 
of preferred embodiment which proceeds with reference to the drawings.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
The fracture toughness, fatigue resistance, corrosion resistance, and 
formability properties of the present invention are dependent upon a 
chemical composition that is closely controlled within specific limits as 
set forth below and upon carefully controlled and sequenced process steps. 
If the composition limits or process parameters stray from the limits set 
forth below, the desired combination of fracture toughness, fatigue 
resistance, corrosion resistance, formability, and surface smoothness 
objectives will not be achieved. 
The aluminum alloy of the present invention comprises about 3.8 to 4.5 wt. 
% copper, about 1.2 to 1.6 wt. % magnesium, about 0.3 to 0.6 wt. % 
manganese, not more than about 0.15 wt. % silicon, not more than about 
0.12 wt. % iron, and not more than about 0.10 wt. % titanium, the balance 
being aluminum and impurity elements. For any remaining trace elements, 
each has a maximum limit 0.05 wt. %, with a total maximum of 0.15 wt. %. A 
preferred alloy would comprise about 4.0 to 4.4 wt. % copper, about 1.25 
to 1.5 wt. % magnesium, about 0.35 to 0.50 wt. % manganese, not more than 
about 0.12 wt. % silicon, not more than about 0.08 wt. % iron, and not 
more than about 0.06 wt. % titanium, the balance being aluminum and 
impurity elements. 
The chemical composition of the alloy of the present invention is similar 
to that of alloy 2024, but is distinctive in several important aspects. 
The alloying elements contained in the allowed range of variation for 
alloying elements contained in the invention alloy is less than for 2024. 
This is important because many mechanical and physical properties change 
as composition changes. To maintain the desired close balance of 
properties of the invention it is therefore necessary to restrict 
composition changes to a greater degree than is normally done. In 
addition, to the restricted ranges of copper, magnesium, and manganese, 
the silicon, iron, and titanium concentrations are reduced to the lowest 
levels commercially feasible for aluminum alloys of the present type in 
order to improve the fracture toughness. 
Improved Fracture Toughness 
The "damage tolerant" design philosophy being used today for commercial and 
military aircraft assumes that all structures contain flaws (cracks). The 
stress at the tip of a sharp crack is characterized by a stress intensity 
factor K, given by 
EQU K=Y.sigma..sqroot. c. 
where .sigma. is the average applied stress on the structure (pounds per 
square inch), Y is a dimensionless parameter dependant on the geometry of 
the structural member, and c is the crack length. The stress intensity 
factor at which the crack begins to extend, generally resulting in 
catastrophic failure, is known as the fracture toughness of the material. 
I now consider the factors that affect the fracture toughness of heat 
treated aluminum alloys, i.e., those alloys that derive their strength 
from thermal treatments such as a "solutionizing" operation that dissolves 
the alloying elements, followed by rapidly cooling to room temperature 
(quenching), and then "aging" at room temperature or higher (250 to 
375.degree. F.) to precipitate the alloying elements as small discrete 
particles within the aluminum matrix. Two major controls must be 
maintained on the chemical composition of these alloys. First, the iron 
impurity level must be kept to a minimum, because it is insoluble in 
aluminum and forms coarse, brittle intermetallic particles that contribute 
to crack extension and fracture. Second, the amounts of the major alloying 
elements, such as copper and magnesium, should be controlled to ensure 
that they are dissolved into the aluminum matrix during the solutionizing 
operation. Any intermetallic particles that are left undissolved reduce 
fracture toughness as do those resulting from high iron levels. 
FIG. 1 graphically illustrates an equilibrium phase diagram for the 
aluminum (Al)--copper (Cu)--magnesium (Mg) system at 930.degree. F. 
Specifically, FIG. 1 defines the copper and magnesium concentrations that 
can be dissolved. If the limits defined by the alpha aluminum region are 
exceeded, undissolved particles of Al.sub.2 CuMg (commonly designated as 
"S" phase) and Al.sub.2 Cu (commonly designated as ".theta." phase), 
remain after solution heat treatment. This situation is complicated by the 
presence of iron, which can combine with copper to from an insoluble 
Al.sub.7 Cu.sub.2 Fe intermetallic constituent. The copper level in FIG. 1 
therefore must be adjusted upwards by an amount equal to approximately 
twice the iron concentration because the Al.sub.7 Cu.sub.2 Fe constituent 
contains about two times as much copper as iron. 
A third compositional factor is the role of sparingly soluble alloying 
elements such as chromium, manganese and zirconium. One or more of these 
alloying elements are intentionally added to aluminum to form 
"dispersoids," which are small intermetallic particles that are useful in 
controlling the crystallite, or "grain" structure of aluminum alloys. All 
metallic products are comprised of numerous crystallites, or grains, which 
should not be allowed to grow to a large size during any of the thermal 
processing operations, because strength and good fracture toughness are 
favored by small grains. The dispersoid particles act to "pin" the grains 
and prevent their growth. 
The dispersoid forming element in Al-Cu-Mg alloy 2024 is manganese in the 
range of 0.3 to 0.9%. Unexpectedly, I have discovered a significant effect 
of manganese on fracture toughness as measured by two test methods. In one 
method, I tested to failure 16-inch wide by 36-inch long panels with a 
4-inch long through--thickness sharp crack in the orientation transverse 
to the rolling direction (T-L). Using an equation similar to 
K=Y.sigma..sqroot. c, above, I calculated values of K (apparent) or 
K.sub.app. It is noteworthy that K.sub.app determined in this manner is 
only an indicator of the true fracture toughness, because the stress 
required to cause failure exceeded the elastic limit of the material. A 
much wider panel would be required to obtain the actual fracture 
toughness, and is beyond the capability of most test laboratories. In the 
second method, a 1.5-inch wide by 2.25-inch long panel with a sharp notch 
on one side was pulled to failure and the tear strength (maximum load 
divided by the cross sectional area) was measured. The tear strength 
divided by the yield strength as determined in a standard tensile test, 
commonly called the tear-yield ratio (TYR), is known to correlate with 
fracture toughness. Table 1 illustrates a number of production lots of 
2024 alloy sheets having various iron and manganese contents which I 
tested for toughness by the aforementioned methods. Table 2 illustrates 
the results of these tests. 
TABLE 1 
______________________________________ 
CHEMICAL COMPOSITIONS OF PRODUCTION LOTS OF 2024-T3 
SHEETS 
(Core alloy with cladding removed) 
% by wt.sup.a 
Alloy 
Si Fe Cu Mn Mg Ti Zn 
______________________________________ 
1 &lt;0.1 0.035 4.31 0.33 1.37 0.02 0.02 
2 &lt;0.1 0.04 4.21 0.46 1.28 0.02 0.02 
3 &lt;0.1 0.07 3.99 0.32 1.37 0.02 0.06 
4 &lt;0.1 0.07 3.99 0.44 1.28 0.04 0.22 
5 &lt;0.1 0.17 4.21 0.39 1.44 0.03 0.03 
6 &lt;0.1 0.16 4.17 0.77 1.21 0.03 0.07 
7 &lt;0.1 0.19 4.43 0.54 1.48 0.01 0.01 
______________________________________ 
.sup.a By inductively coupied plasma spectroscopy. 
TABLE 2 
______________________________________ 
EFFECT OF IRON AND MANGANESE CONTENTS ON TOUGHNESS 
OF ALCLAD 2024-T3 SHEET.sup.a 
T-L K.sub.app YS.sup.c 
Alloy No. 
% Fe % Mn (ksi.check mark.in) 
T-L TS/YS.sup.b 
(ksi) 
______________________________________ 
1 0.035 0.33 89 1.76 45.5 
2 0.04 0.46 87 1.70 46.3 
3 0.07 0.32 85 1.65 45.4 
4 0.07 0.44 82 1.59 44.4 
5 0.17 0.39 83.5 1.60 43.4 
6 0.16 0.77 77.5 1.52 44.5 
7 0.19 0.54 79.5 1.53 44.2 
______________________________________ 
.sup.a Also: 4.0-4.5% Cu and 1.2-1.5% Mg; all 0.063" thick. 
.sup.b Tear strengthyield strength ratio (TYR). 
.sup.c Transverse tensile yield strength. 
I used data from Table 2 to plot the toughness measurements as a function 
of wt. % iron in FIGS. 2 and 3. As expected, FIGS. 2 and 3 demonstrate a 
correlation of fracture toughness with decreasing concentrations of iron. 
Surprisingly, however, the lots with relatively low manganese levels 
exhibit higher toughness values for a given iron content. Table 3, which 
compares the toughness levels at two manganese levels for a number of iron 
concentrations, also demonstrates this phenomenon. Table 3 also lists 
copper contents for each alloy, because high levels of copper can reduce 
toughness by the presence of undissolved Al.sub.2 Cu and Al.sub.2 CuMg 
phases. Notably, the copper levels of the alloys being compared in each 
case are almost equivalent. 
TABLE 3 
______________________________________ 
EFFECT OF MANGANESE AT VARIOUS IRON LEVELS ON 
TOUGHNESS OF ALCLAD 2024-T3 SHEET 
ALLOY % MN % CU T-L K.sub.app 
T-L TS/YS.sup.b 
______________________________________ 
0.035-0.04% Fe 
1 0.33 4.31 89 1.76 
2 0.46 4.21 87 1.70 
.DELTA.(%) 2.3 3.50 
0.07% Fe 
3 0.32 3.99 85 1.65 
4 0.44 3.99 82 1.59 
.DELTA.(%) 3.5 3.80 
0.16-0.17% Fe 
5 0.39 4.17 83.5 1.6 
6 0.77 4.21 77.5 1.52 
.DELTA.(%) 7.7 5.3 
Ave. .DELTA./0.1% 2.2 2.4 
Mn 
______________________________________ 
.sup.a Tear strengthyield strength ratio (TYR). 
A linear regression analysis of the K.sub.app data showed that manganese 
has approximately half the detrimental effect that iron has. This 
discovery is particularly important because of the relatively high levels 
of manganese in alloys such as 2024. Specifically, FIG. 4 demonstrates 
toughness, K.sub.app, as a function of iron and manganese concentrations, 
producing the correlation: 
EQU K.sub.app =93.2-29.2 (% Fe+0.50 % Mn) 
Similarly, toughness, expressed as tear-yield ratio or TYR, is represented 
by the correlation: 
EQU TYR=1.81-0.27(2% Fe+Mn) 
This correlation is illustrated in FIG. 5. 
Improved Fatigue Resistance 
As noted in the previous section, the "damage tolerant" design philosophy 
assumes that flaws (cracks) are present in all structural materials. If 
these cracks are permitted to grow to a "critical" size such that the 
stress intensity factor at the crack tip exceeds the fracture toughness of 
the material, catastrophic failure occurs. Cracks can grow as a result of 
cyclic loads (fatigue) caused by takeoff and landing or cabin 
pressurization and depressurization. Fatigue crack growth rates for the 
projected cyclic loading stresses are therefore desirably low. 
Experimentally, I have found that the velocity of cracks growing under 
fatigue conditions, i.e., the fatigue crack growth rate, is dependent on 
the stress intensity factor difference (.DELTA.K) associated with the 
minimum and maximum load. The stress intensity factor increment (.DELTA.K) 
must therefore, be specified when comparing fatigue crack growth rates for 
different materials 
In addition to improved toughness, I also discovered that higher purity 
alloys with relatively low manganese levels also had low fatigue crack 
growth rates. I determined this by running tests at a stress intensity 
increment (.DELTA.K) of 30 ksi.sqroot.in. and a load ratio (maximum load 
divided by minimum load) of 0.1 on several of the alloys listed in Tables 
1-3. For example, alloys 1 and 2 had average crack growth rates of 
7.0.times.10-5 and 7.5.times.10-5 inches/cycle, compared to a nominal 
value of 20.times.10-5 inch/cycle for standard 2024 alloy typified by 
alloy 7. Thus, the alloy of my invention has about a 50% decrease in crack 
growth rate over standard 2024 alloy at a .DELTA.K of 30 ksi.sqroot.in. 
Similarly, I discovered fatigue benefits at lower values of .DELTA.K. For 
example, at a .DELTA.K of 5 ksi.sqroot.in., alloys 1,2,3, and 4 had crack 
growth rates of 1.5 to 2.2.times.10-7 inches/cycle compared to 1.7 to 
4.0.times.10-7 inches/cycle for standard 2024 alloy. Or stated another 
way, my new alloy had about a 25% decrease in crack growth rate in the low 
.DELTA.K regime. 
Improved Corrision Resistance 
Yet another benefit of the new alloy of my invention is improved corrosion 
resistance. As noted earlier, good corrosion resistance is of prime 
concern in aircraft fuselage structures. Corrosion of aluminum alloys is 
usually aggravated by salt (sodium chloride) containing environments such 
as can be present near oceans. Sheet samples from alloys 3 and 7 (of 
Tables 1-3) were therefore exposed to a marine atmosphere at Daytona 
Beach, Fla. for one year. The protective cladding was removed from one 
surface so that the inherent corrosion resistance of the core alloy could 
be assessed. This also simulates the practical situation where one side of 
a fuselage panel is chemically milled to a thinner section size. After the 
one-year exposure period, tensile specimens were machined from the 
samples, and as recommended in the Corrosion Handbook (edited by H. H. 
Uhlig, John Wiley & Sons, p. 956), the corrosion damage was quantified by 
loss in ductility. This method is particularly suited to materials that 
are susceptible to pitting and intergranular corrosion. Table 4 summarizes 
tensile elongation measurements before and after the exposure to the 
marine atmosphere. Metallographic examination revealed that ductility loss 
corresponded with the depth of pitting corrosion attack on the exposed and 
corroded alloys. It is apparent that alloy 3, which has lower iron and 
manganese contents, is superior in corrosion resistance. 
TABLE 4 
______________________________________ 
EFFECT OF MARINE EXPOSURE ON DUCTILITY LOSS 
Elongation, 
% in 1 inch 
Alloy Before After % Loss in Ductility 
______________________________________ 
3 23.5 19.1 19 
7 22.5 14.5 36 
______________________________________ 
Improved Formability 
Another advantage of my invention is improved formability. Good formability 
is important to the aircraft manufacturers because of lower costs because 
of reduced scrap rates and manpower requirements. Two indicators of 
formability are (1) ball punch depth as determined by indenting the sheet 
with a 1-inch diameter steel ball until it cracks (also known as Olsen cup 
depth), a measure of a material's capability of being stretched in more 
than one direction, and (2) minimum bend radius, a measure of a material's 
ability to be bent without cracking. Note that there is some uncertainty 
in minimum bend radius measurements because the determination of surface 
cracking is somewhat subjective, and the method involves bending sheet 
samples around dies of incremental (not continuously varying) radii. Table 
5 lists bend radius and ball punch depth of alloys 1, 2, 4, 6 and 7. As 
FIG. 6 illustates, both of these indicators correlate with % Fe+1/2% Mn, 
i.e., alloys with about 0.1 % Fe and less than about 0.5% Mn have superior 
formability. 
TABLE 5 
______________________________________ 
FORMABILITY OF 2024-T3 SHEET 
Olsen Cup 
180.degree. Min. Bend 
Alloy % Fe % Me Depth, in. 
Radius, in. 
______________________________________ 
1 0.035 0.33 0.336 0.025-0.032 
2 0.04 0.46 0.319 0.025-0.032 
4 0.07 0.44 0.333 0.032-0.064 
6 0.16 0.77 0.287 0.080-0.100 
7 0.19 0.54 0.309 0.080-0.100 
______________________________________ 
Improved Surface Roughness 
Three lots each of standard 2024 and the invention composition were 
chemically milled to half thickness in a buffered 14% NaOH solution. The 
roughness of the milled surfaces was measured in a direction perpendicular 
to the rolling direction using a profilometer with a 2 .mu.m 
(2.times.10.sup.-6 meters) diamond stylus. The results listed in Table 6 
show a 10 to 45% improvement for the invention product. 
TABLE 6 
______________________________________ 
SURFACE ROUGHNESS OF CHEMICALLY MILLED SHEET 
Roughness 
Gage, in. Alloy (.times. 10.sup.-6 in.) 
______________________________________ 
0.125 IP.sup.a 
58 
2024 107 
0.160 IP 107 
2024 119 
0.190 IP 139 
2024 186 
______________________________________ 
.sup.a Invention Product 
Improved Physical Properties by Intermediate Annealing Step 
I have also discovered that I can further improve the properties that I 
have discussed above by an intermediate thermal treatment. Specifically, I 
introduce an intermediate annealing step after hot rolling but before cold 
rolling to the final gage to produce an improved alloy. 
For purposes of the present invention, I prefer a method which includes 
providing stock comprising an aluminum alloy having about 3.8 to 4.5 wt. % 
copper, about 1.2 to 1.6 wt. % magnesium, about 0.3 to 0.6 wt. % 
manganese, not more than 15 wt. % silicon, not more than 12 wt. % iron, 
not more than 0.1 wt. % titanium, the remainder substantially aluminum, 
incidental elements and impurities; hot working the stock; annealing; cold 
rolling; solution heat treating; and cooling. 
Optionally, before the hot working step, I homogenize the stock to produce 
a substantially uniform distribution of alloying elements. In general, I 
homogenize by heating the stock to a temperature ranging from about 900 to 
975.degree. F. for a period of at least 1.0 hour to dissolve soluble 
elements and to homogenize the internal structure of the metal. I caution, 
however, that temperatures above 935OF are likely to damage the metal and 
thus I avoid these increased temperatures if possible. Generally, I 
homogenize for at least 4.0 hours in the homogenization temperature range. 
Most preferably, I homogenize for about 6.0 to 12.0 hours at about 
920.degree. F. 
As discussed above, my preferred aluminum alloy comprises about 4.0 to 4.4 
wt. % copper, about 1.25 to 1.5 wt. % magnesium, about 0.35 to 0.5 wt. % 
manganese, not more than 0.12 wt. % silicon, not more than 0.08 wt. % 
iron, not more than 0.06 wt. % titanium, the remainder substantially 
aluminum, incidental elements and impurities. 
For hot working, I prefer a hot rolling step where the stock is heated to a 
temperature ranging from about 750 to 925.degree. F. for about 1.0 to 12.0 
hours. Most preferably, I heat the stock to a temperature ranging from 
about 825 to 900.degree. F. for about 1.0 to 2.0 hours to obtain a gage 
thickness ranging from about 0.1 to 0.25 inches. I generally perform hot 
rolling at a starting temperature ranging from about 600 to 900.degree. 
F., or even higher as long as no melting or other ingot damage occurs. 
When the alloy is to be used for fuselage skins, for example, I typically 
perform hot rolling on ingot or starting stock 12 to 16 or more inches 
thick to provide an intermediate product having a thickness ranging from 
about 0.1 to 0.25 inches. 
After hot rolling, I next anneal the stock. Preferably, I anneal at a 
temperature ranging from about 725 to 875.degree. F. for about 1.0 to 12.0 
hours. Most preferably, I anneal the stock at a temperature ranging from 
about 750 to 850.degree. F. for about 4.0 to 6.0 hours at heating rate 
ranging from about 25 to 100.degree. F. per hour, with the optimum being 
about 50.degree. F. per hour. 
After annealing, I next cold roll the intermediate gage stock. Preferably, 
I allow the annealed stock to cool to less than 100.degree. F. and most 
preferably to room temperature before I begin cold rolling. Preferably, I 
cold roll to obtain at least a 40% reduction in sheet thickness, most 
preferably I cold roll to a thickness ranging from about 50 to 70% of the 
hot rolled gage. 
After cold rolling, I next solution heat treat the stock. Preferably, I 
solution heat treat at a temperature ranging from about 900 to about 
940.degree. F. for about 10 to 30 minutes. It is important to rapidly heat 
the stock, preferably at a heating rate of about 100 to 2000.degree. F. 
per minute. Most preferably, I solution heat treat at about 920 to 
930.degree. F. for about 15 minutes at a heating rate of about 
1000.degree. F. per minute. 
If the temperature is substantially below 920.degree. F., then the soluble 
elements, copper and magnesium are not taken into solid solution. This 
circumstance can be illustrated by reference to FIG. 1. As the temperature 
is decreased, the lines encompassing the aluminum solid solution region 
shift to the left as depicted by the arrows. When copper and magnesium are 
not taken into solution, two undesirable consequences result: (1) there 
are insufficient alloying elements to provide adequate strength upon 
subsequent age hardening; and (2) the copper and magnesium-containing 
intermetallic compounds (Al.sub.2 Cu and Al.sub.2 CuMg) that remain 
undissolved detract from fracture toughness and fatigue resistance. 
Similarly, if the time at the solution heat treatment temperature is too 
short, these intermetallic compounds do not have time to dissolve. The 
heating rate to the solutionizing temperature is important because 
relatively fast rates generate a fine grain (crystallite) size, which is 
desirable for good fracture toughness and high strength. 
After solution heat treatment, I rapidly cool the stock to minimize 
uncontrolled precipitation of secondary phases, such as Al.sub.2 CuMg and 
Al.sub.2 Cu. Preferably, I quench at a rate of about 1000.degree. F./sec. 
over the temperature range 750 to 550.degree. from the solution 
temperature to a temperature of 100.degree. F. or lower. Most preferably, 
I quench using a high pressure water spray at room temperature or by 
immersion into a water bath at room temperature, generally ranging from 
about 60 to 80.degree. F. 
EXAMPLE 
To demonstrate the present invention, I first homogenized two 3".times.9" 
ingots having the composition listed in Table 7 at a temperature of about 
910.degree. F. for about 15 hours. 
TABLE 7 
______________________________________ 
CHEMICAL COMPOSITIONS OF LABORATORY INGOTS 
% by wt. 
Alloy Si Fe Cu Mn Mg Ti 
______________________________________ 
A 0.07 0.07 3.84 0.54 1.24 0.02 
B 0.07 0.09 3.83 0.98 1.22 0.02 
______________________________________ 
I then reheated the ingots to a temperature of about 800.degree. F. and hot 
rolled them to an intermediate gage thickness of about 0.200" having a 
final temperature of about 550.degree. F. I then divided each hot rolled 
sheet into two sheets. I annealed one of the sheets at a temperature of 
about 835.degree. F. for about 2.0 hours using heating and cooling rates 
of about 50.degree. F./hr. The other control sheet was not annealed. Then, 
I cold rolled all four sheets to a gage of about 0.063" and solution 
heated treated them at about 920.degree. F. for about 30 minutes. Finally, 
I quenched all four sheets in room temperature water. I then tested all 
four sheets after naturally aging them at room temperature for greater 
than one month (T4 temper) for tensile properties and Kahn tear energy, 
which are listed in Table 8. 
TABLE 8 
______________________________________ 
PROPERTIES WITH AND WITHOUT INTERMEDIATE ANNEAL 
UPE, in- 
Alloy 
Anneal UTS, ksi YS, ksi 
Elong, % 
lb/in.sup.2 
______________________________________ 
A Yes 68.3 42.6 23.5 845 
A No 66.4 40.8 24.5 755 
B Yes 68.9 42.3 22 705 
B No 68.4 41.2 21 650 
______________________________________ 
I used the data from Table 8 to plot yield strength versus unit propagation 
energy in FIG. 7. Surprisingly, the intermediate-annealed variants of both 
alloys were not only somewhat stronger, they were also significantly 
tougher than their un-annealed counterparts. The lower manganese Alloy A 
also had higher toughness values than the high manganese Alloy B, as 
expected based on FIGS. 5-7 and my previous discussion, above. 
In addition to improvements in mechanical properties, the sheets produced 
with intermediate annealed had improved formability as evidenced by deeper 
ball punch depths shown in Table 9. 
TABLE 9 
______________________________________ 
BALL PUNCH DEPTHS 
WITH AND WITHOUT INTERMEDIATE ANNEAL 
Olsen Cup 
Alloy Anneal Depth, in. 
______________________________________ 
A Yes 0.330 
A No 0.304 
B Yes 0.295 
B No 0.265 
______________________________________ 
Notably, the lower manganese alloy also had superior forming behavior as 
would be expected based on my previous discussion. 
In addition, when I examined the grain structures of Alloys A and B, I 
discovered considerably finer grain sizes in the intermediate-annealed 
alloys. FIGS. 8a and 9a compared to FIGS. 8b and 9b, respectively, 
illustrate the phenomenon of finer grain size that I observed. 
Having illustrated and described the principles of my invention in a 
preferred embodiment thereof, it should be readily apparent to those 
skilled in the art that the invention can be modified in arrangement and 
detail without departing from such principles. I claim all modifications 
coming within the spirit and scope of the accompanying claims.