AU2001296519A1 - Aluminum alloy products and artificial aging nethod - Google Patents

Description

ALUMINUM ALLOY PRODUCTS AND ARTIFICIAL AGING METHOD

FIELD OF THE INVENTION

[0002] This invention relates to aluminum alloys, particularly 7000 Series (or

7XXX) aluminum ("Al") alloys as designated by the Aluminum Association. More

particularly, the invention relates to Al alloy products in relatively thick gauges, i.e. about

2-12 inches thick. While typically practiced on rolled plate product forms, this invention

may also find use with extrusions or forged product shapes. Through the practice of this

invention, parts made from such thick-sectioned starting materials/products have superior

strength - toughness property combinations making them suitable for structural parts in

various aerospace applications as thick gauge parts or as parts with thinner sections

machined from thick material. Valuable improvements in corrosion resistance

performance have also been imparted by the invention, particularly with respect to stress

corrosion cracking (or "SCC") resistance. Representative structural component parts

made from this alloy include integral spar members and the like which are machined from

thick wrought sections, including rolled plate. Such spar members can be used in the

wingboxes of high capacity aircraft. This invention is particularly suitable for manufacturing high strength extrusions and forged aircraft components, such as, for

example,* main landing gear beams. Such aircraft include commercial passenger jetliners,

cargo planes (as used by overnight mail service providers) and certain military planes. To

a lesser degree, the alloys of this invention are suitable for use in other aircraft including

but not limited to turbo prop planes. In addition, non-aerospace parts like various cast

thick mold plates may be made according to this invention.

[0003] As the size of new jet aircraft get larger, or as current jetliner models grow

to accommodate heavier payloads and/or longer flight ranges to improve performance and

economy, the demand for weight savings of structural components, such as fuselage,

wing and spar parts continues to increase. The aircraft industry is meeting this demand

by specifying higher strength, metal parts to enable reduced section thicknesses as a

weight savings expedient. In addition to strength, the durability and damage tolerance of

materials are also critical to an aircraft's fail-safe structural design. Such consideration of

multiple material attributes for aircraft applications eventually led to today's damage

tolerant designs, which combine the principles of fail-safe design with periodic inspection

techniques.

[0004] A traditional aircraft wing structure comprises a wing box generally

designated by numeral 2 in accompanying Figure 1. It extends outwardly from the

fuselage as the main strength component ofthe wing and runs generally perpendicular to

the plane of Figure 1. That wing box 2 comprises upper and lower wing skins 4 and 6

spaced by vertical structural members or spars 12 and 20 extending between or bridging

upper and lower wing skins. The wing box also includes ribs which can extend generally from one spar to the other. These ribs lie parallel to the plane of Figure 1 whereas the

wing skins and spars run perpendicular to said Figure 1 plane. During flight, the upper

wing structures of a commercial aircraft wing are compressively loaded, calling for high

compressive strengths with an acceptable fracture toughness attribute. The upper wing

skins of today's most large aircraft are typically made from 7XXX series aluminum

alloys such as 7150 (U.S. Reissue Patent No. 34,008) or 7055 aluminum (U.S. Patent No.

5,221,377). Because the lower wing structures of these same aircraft wings are under

tension during flight, they will require a higher damage tolerance than their upper wing

counterparts. Although one might desire to design lower wings using a higher strength

alloy to maximize weight efficiency, the damage tolerance characteristics of such alloys

often fall short of design expectations. As such, most commercial jetliner manufacturers

today specify a more damage-tolerant 2XXX series alloy, such as 2024 or 2324 aluminum

(U.S. Patent No. 4,294,625), for their lower wing applications, both of said 2XXX alloys

being lower in strength than their upper wing, 7XXX series counterparts. The alloy

members and temper designations used throughout are in accordance with the well-

known product standards ofthe Aluminum Association.

[0005] Upper and lower wing skins, 4 and 6 respectively, from accompanying

Figure 1 are typically stiffened by longitudinally extending stringer members 8 and 10.

Such stringer members may assume a variety of shapes, including "J", "I", "L", "T"

and/or "Z" cross sectional configurations. These stringer members are typically fastened

to a wing skin inner surface as shown in Figure 1, the fasteners typically being rivets.

Upper wing stringer member 8 and upper spar caps 14 and 22 are presently manufactured from a 7XXX series alloy, with lower wing stringer 10 and lower spar caps 16 and 24

being made from a 2XXX series alloy for the same structural reasons discussed above

regarding relative strength and damage-tolerance. Vertical spar web members 18 and 26,

also made from 7XXX alloys, fasten to both upper and lower spar caps while running in

the longitudinal direction ofthe wing constituted by member spars 12 and 20. This

traditional spar design is also known as a "built-up" spar, comprising upper spar cap 14 or

22, web 18 or 20, and lower spar cap 16 or 24, with fasteners (not shown). Obviously,

the fasteners and fastener holes at the joints to this spar are structural weak links. In

order to ensure the structural integrity of a built-up spar like 18 or 20, many component

parts like the web and/or spar cap have to be thickened, thereby adding weight to the

overall structure.

[0006] One potential design approach for overcoming the aforementioned spar

weight penalty is to make an upper spar, web and lower spar by machining from a thick

simple section, such as plate, of aluminum alloy product, typically by removing

substantial amounts of metal to make a more complex, less thick section or shape such as

a spar. Sometimes, this machining operation is known as "hogging out" the part from its

plate product. With such a design, one could eliminate the need for making web-to-upper

spar and web-to-lower spar joints. A one-piece spar like that is sometimes known as an

"integral spar" and can be machined from a thick plate, extrusion or forging. Integral

spars should not only weigh less than their built up counterparts; they should also be less

costly to make and assemble by eliminating the need for fasteners. An ideal alloy for

making integral spars should have the strength characteristics of an upper wing alloy combined with the fracture toughness/damage tolerance requirements of a lower wing

alloy. Existing commercial alloys used on aircraft do not satisfy this combination of

preferred property requirements. The lower strengths of lower wing skin alloy 2024-

T351, for example, will not safely carry the load transmittals from a highly loaded, upper

wing unless its section thicknesses are significantly increased. That, in turn, would add

undesirable weight to the overall wing structure. Conversely, designing an upper wing to

2XXX strength capabilities would result in an overall weight penalty.

[0007] Large jet aircrafts require very large wings. Making integral spars for such

wings would require products as thick as 6 to 8 inches or more. Alloy 7050-T74 is often

used for thick sections. The industry standard for 6 inch thick 7050-T7451 plate, as listed

in Aerospace Materials Specification AMS 4050F, specifies a minimum yield strength in

the longitudinal (L) direction of 60 ksi and a plane-strain fracture toughness, or K_lc (L-T),

of 24 ksi in. For that same alloy temper and thickness, specified values in the

transverse direction (LT and T-L) are 60 ksi and 22 ks in , respectively. By comparison,

the more recently developed upper wing alloy, 7055-T7751 aluminum, about 0.375 to

1.5 inches thick, can meet a minimum yield strength of 86 ksi according to MIL-HDBK-

5H. If an integral spar of 7050-T74, with a 60 ksi minimum yield strength is used with

the aforesaid 7055 alloy, overall strength capabilities of that upper wing skin would not

be taken full advantage of for maximum weight efficiencies. Hence, higher strength,

thick aluminum alloys with sufficient fracture toughness are needed for manufacturing

the integral spar configurations now desired for new jetliner designs. This is but one specific example ofthe benefits of an aluminum material with high strength and

toughness in thick sections, but many others exist in modem aircraft, such as the wing

ribs, webs or stringers, wing panels or skins, the fuselage frame, floor beam or

bulkheads, even landing gear beams or various combinations of these aircraft structural

cpmponents.

[0008] The varying tempers that result from different artificial aging treatments are

known to impart different levels of strength and other performance characteristics

including corrosion resistance and fracture toughness. 7XXX series alloys are most often

made and sold in such artificially aged conditions as "peak" strength ("T6-type") or

"over-aged" ("T7-type") tempers. U.S. Patent Nos. 4,863,528, 4,832,758, 4,477,292 and

5,108,520 each describe 7XXX series alloy tempers with a range of strength and

performance property combinations. All ofthe contents of those patents are fully

incorporated by reference herein.

[0009] It is well known to those skilled in the art that for a given 7XXX series

wrought alloy, peak strength or T6-type tempers provide the highest strength values, but

in combination with comparatively low fracture toughness and corrosion resistance

performance. For these same alloys, it is also known that most over-aged tempering, like

a typical T73-type temper, will impart the highest fracture toughness and corrosion

resistance but at a significantly lower relative strength value. When making a given

aerospace part, therefore, part designers must select an appropriate temper somewhere

between the aforesaid two extremes to suit that particular application. A more complete description of tempers, including the "T-XX" suffix, can be found in the Aluminum

Association's Aluminum Standards and Data 2000 publication as is well known in the art.

[0010] Most aerospace alloy processing requires a solution heat treatment (or

"SHT") followed by quenching and subsequent artificial aging to develop strength and

qther properties. However, seeking improved properties in thick sections faces two

natural phenomena. First, as a product shape thickens, the quench rate experienced at the

interior cross section of that product naturally decreases. That decrease, in turn, results in

a loss of strength and fracture toughness for thicker product shapes, especially in inner

regions across the thickness. Those skilled in the art refer to this phenomenon as "quench

sensitivity". Second, there is also a well known, inverse relationship between strength

and fracture toughness such that as component parts are designed for ever greater strength

loads, their relative toughness performance decreases...and vice versa.

[0011] To better understand the present invention, certain demonstrated trends in

the art of commercial aerospace 7XXX series alloys are worth considering. Aluminum

alloy 7050, for example, substitutes Zr for Cr as a dispersoid agent for greater grain

structure control and increases both Cu and Zn contents over the older 7075 alloy. Alloy

7050 provided a significant improvement in (i.e. by decreasing) quench sensitivity over

its 7075 alloy predecessor, thereby establishing 7050 aluminum as the mainstay for thick-

sectioned aerospace applications in plate, extrusion and/or forged shapes. For upper wing

applications with still higher strength-toughness requirements, the compositional

minimums for both Mg and Zn in 7050 aluminum were slightly raised to make an

Aluminum Association-registered 7150 alloy variant of 7050. Compared to its 7050 predecessor, the minimum Zn contents for 7150 increased from 5.7 to 5.9 wt. %, and Mg

level minimums rose from 1.9 to 2.0 wt. %.

[0012] Eventually, a newer upper wing skin alloy was developed. That alloy 7055

exhibited a 10 % improvement in compression yield strength, in part, by employing a

higher range of Zn, from 7.6 to 8.4 wt %, with a similar Cu level and slightly lower Mg

range (1.8 to 2.3 wt %) compared to either alloy 7050 or 7150.

[0013] Past efforts for still higher strengths (by increasing alloying components

and compositional optimizations), had to be offset with metal purity increases and

microstructure control through thermal-mechanical processing ("TMP") to obtain

improvements in toughness and fatigue life among other properties. U.S. Patent No.

5,865,911 reported a significant improvement in toughness, at equivalent strengths, for a

7XXX series alloy plate. However, the quench sensitivity of that alloy, in thicker gauges,

is believed to cause other noticeable property disadvantages.

[0014] Alloy 7040, as registered with the Aluminum Association, calls for the

following ranges of main alloying components: 5.7 - 6.7 wt.% Zn, 1.7 - 2.4 wt.% Mg and

1.5 - 2.3 wt.% Cu. Related literature, namely Shahani et al., "High Strength 7XXX

Alloys For Ultra-Thick Aerospace Plate: Optimization of Alloy Composition," PROC.

ICAA 6, v. 2, pp/ 105-1110 (1998) and U.S. Patent No. 6,027,582, state that 7040

developers pursued an optimization balance between alloying elements for improving

strength and other properties while avoiding excess additions to minimize quench

sensitivity. While thicker gauges of alloy 7040 claimed some property improvements over 7050, those improvements still fall short of newer commercial aircraft designer

needs.

[0015] This invention differs in several key ways from the alloys currently being

supplied on a commercial basis for aerospace-type applications. Main alloying elements

for several current commercial 7XXX aerospace alloys, as listed by the Aluminum

Association, are as follows:

*included in the "0.05% each/0.15% total" for unlisted impurities Note that alloys 7075, 7050, 7010 and 7040 aluminum are supplied to the aerospace

industry in both thick and thin (up to 2 inches) gauges; the others (7150 and 7055) are

generally supplied in thin gauge. By contrast with these commercial alloys, a preferred

alloy in accordance with the invention contains about 6.9 to 8.5 wt.% Zn, 1.2 to 1.7 wt.%

Mg, 1.3 to 2 wt.% Cu, 0.05 to 0.15 wt.% Zr, the balance essentially aluminum, incidental

elements and impurities.

[0016] This invention solves the aforesaid prior art problems with a new 7XXX

series aluminum alloy that, in thicker gauges, exhibits significantly reduced quench

sensitivity so as to provide significantly higher strength and fracture toughness levels than

heretofore possible. The alloy of this invention has a relatively high zinc (Zn) content coupled with lower copper (Cu) and magnesium (Mg) in comparison with the commercial

7XXX aerospace alloys above. For this invention, combined Cu + Mg is usually less than

about 3.5%, and preferably less than about 3.3%. When the aforesaid compositions are

subjected to the preferred 3 -stage aging practice outlined in greater detail below, the

resulting thick wrought product forms (either plate, extrusions or forgings) are shown to

exhibit a highly desirable combination of sfrength, fracture toughness and fatigue

performance, in further combination with superior stress corrosion cracking (SCC)

resistance, particularly when subjected to atmospheric, seacoast type test conditions.

[0017] Prior art examples for aging 7XXX Al alloys in three steps or stages are

known. Representative are U.S. Patent Nos. 3,856,584, 4,477,292, 4,832,758, 4,863,528

and 5,108,520. The first step/stage for many ofthe aforementioned prior art processes

was typically performed at around 250°F. The preferred first step for the alloy

composition of this invention ages between about 150-275°F, preferably between about

200-275°F, and more preferably from about 225 or 230°F to about 250 or 260°F. This

first step or stage can include two temperatures, such as 225°F for about 4 hours, plus

250°F for about 6 hours, both of which count only as the "first stage", i.e. the stage

preceding the second (e.g. about 300°F ) stage described below. Most preferably, the first

aging step of this invention operates at about 250°F, for at least about 2 hours, preferably

for about 6 to 12, and sometimes for as much as 18 hours or more. It should be noted,

however, that shorter holding times can suffice depending on part size (i.e. thickness) and

shape complexity, coupled with the degree to which equipment ramp up temperatures (i.e. relatively slow heat up rates) may be employed in conjunction with short hold times at

temperature for these alloys.

[0018] Preferred second steps in some prior art, 3 step artificial aging practices

normally took place above about 350 or 360°F or higher, followed by a third step age

similar to their first step, at about 250°F. By contrast, the preferred second aging stage

of this invention differs by proceeding at significantly lower temperatures, about 40 to

50°F lower. For preferred embodiments of this 3-stage aging method on the 7XXX alloy

compositions specified herein, the second of three stages or steps should take place from

about 290 or 300°F to about 330 or 335°F. More particularly, that second aging step or

stage should be performed between about 305 and 325°F , with a more preferred second

step aging range occurring between about 310 to 320 or 325°F. Preferred exposure times

for this second step processing depend inversely on the temperature(s) employed. For

instance, if one were to operate substantially at or very near 310°F, a total exposure time

from about 6 to 18 hours would suffice. More preferably, second stage agings should

proceed for about 8 or 10 to 15 total hours at that operating temperature. At a

temperature of about 320°F, total second step times can range between about 6 to 10

hours with about 7 or 8 to 10 or 11 hours being preferred. There is also a preferred target

property aspect to second step aging time and temperature selection. Most notably,

shorter treatment times at a given temperature favor relatively higher strength values

whereas longer exposure times favor better corrosion resistance performance. [0019] The foregoing second stage age is then followed by a third aging stage at a

lower temperature. One preferably should not ramp slowly down from the second step

for performing this third step on thicker workpieces unless extreme care is exercised to

coordinate closely with the second step temperature and total time duration so as to avoid

exposures at higher (second stage type) temperatures for too long. Between the second

and third aging steps, the metal products of this invention can be pu osefully removed

from the heating furnace and rapidly cooled, using fans or the like, to either about 250°F

or less, perhaps even fully back down to room temperature. In any event, the preferred

time/temperature exposures for the third aging stage of this invention closely parallel

those set forth for the first aging step above, at about 150-275°F, preferably between

about 200-275°F, and more preferably from about 225 or 230°F to about 250 or 260°F.

And while the aforementioned method improves particular properties, especially SCC

resistance, for this new family of 7XXX alloys, it is to be understood that similar

combinations of property improvements may be realized by practicing this same 3 -step

aging method on still other 7XXX alloys, including but not limited to 7X50 alloys (either

7050 or 7150 aluminum), 7010 and 7040 aluminum.

[0020] For newer and larger airplanes, manufacturers strongly desire thick

sectioned, aluminum alloy products with compressive yield strengths about 10-15%

higher than those routinely achieved by incumbent alloys 7050, 7010 and/or 7040

aluminum. In response to this need, the present invention 7XXX-type alloy meets the

aforementioned yield strength goals while surprisingly possessing attractive fracture

toughness performance. In addition, this alloy has exhibited excellent stress corrosion cracking resistance when aged by the preferred three stage, artificial aging practices

specified herein. Samples of six inch thick plate made from this alloy passed laboratory

scale, 3.5% salt solution alternate immersion (or "Al") stress corrosion cracking (SCC)

tests. Pursuant to those tests, thick metal samples had to survive at least 30 days without

cracking at a minimum stress of 25 ksi imposed in the short transverse (or "ST") direction

for meeting the T76 tempering conditions currently specified by one major jetliner

manufacturer. These thicker metal samples have also met other static and dynamic

property goals of that jetliner manufacturer.

[0021] While meeting an initial wave of laboratory alternate immersion (Al) SCC

tests at the even higher stress levels of 35 to 45 ksi, the thick alloys samples of this

invention, artificially aged by then known two step tempering practices, exhibited some

unexpected corrosion-related failures, some at even 25 ksi stress levels, when first

exposed to seacoast SCC test conditions. This was even surprising since laboratory-

accelerated, Al SCC tests historically correlated well with atmospheric tests, both

seacoast and industrial. Under these industrial tests, samples of this invention alloy when

aged in 3 stages as described herein for the invention did not fail after 11 months seacoast

exposure to both 25 and 35 ksi stress levels. Even though atmospheric SCC performance

has not been expressly required by aircraft manufacturers' next generation plane

specifications, it nevertheless is considered important for critical aerospace applications

like the spars and ribs of a jetliner's wingbox. Thus while products aged in two stages

may be adequate, the practice of this invention prefers the herein described three stage

artificial aging. [0022] One known "fix" for improving the SCC resistance of some 7XXX alloys

has been to overage the material, but at a typical tradeoff in strength reduction. That sort

of strength tradeoff is undesirable for an integral wing spar because that thick machined

part will still have to meet fairly high compressive yield strength standards. Thus, there is

a clear need for developing an artificial aging practice that won't unduly sacrifice strength

properties while still improving the corrosion resistance of high performance, 7XXX

aluminum alloys. In particular, it is desirable to develop an aging method that will raise

the seacoast SCC performance of these alloys to better levels without compromising

strength and/or other property combinations. The above described three stage aging

method ofthe invention satisfies this need.

[0023] An important aspect of this invention focuses on a newly developed,

aluminum alloy that exhibits significantly reduced quench sensitivity in thick gauges, i.e.,

greater than about 2 inches and, more preferably, in thicknesses ranging from about 4 to 8

inches or greater. A broad compositional breakdown for that alloy consists essentially of:

from about 6% Zn to about 9, 9.5 or 10 wt.% Zn; from about 1.2 or 1.3% Mg to about

1.68, 1.7 or even 1.9 wt. % Mg; from about 1.2, 1.3 or 1.4 wt.% Cu to about 1.9, or even

2.2 wt.% Cu, with %Mg < (%Cu + 0.3 max.); one or more element being present selected

from the group consisting of: up to about 0.3 or 0.4 wt% Zr, up to about 0.4 wt.% Sc, and

up to about 0.3 wt.% Hf, the balance essentially aluminum and incidental elements and

impurities. Except where stated otherwise such as "being present", the expression "up to"

when referring to the amount of an element means that that elemental composition is optional and includes a zero amount of that particular compositional component. Unless

stated otherwise, all compositional percentages are in weight percent (wt.%).

[0024] When used herein, the term "substantially free" means that no purposeful

additions of that alloying element were made to the composition, but that due to

impurities and/or leaching from contact with manufacturing equipment, trace quantities of

such elements may, nevertheless, find their way into the final alloy product. It is to be

understood, however, that the scope of this invention should not/cannot be avoided

through the mere addition of any such element or elements in quantities that would not

otherwise impact on the combinations of properties desired and attained herein.

[0025] When referring to any numerical range of values, such ranges are

understood to include each and every number and/or fraction between the stated range

minimum and maximum. A range of about 6 to 10 wt% zinc, for example, would

expressly include all intermediate values of about 6.1, 6.2, 6.3 and 6.5%, all the way up to

and including 9.5, 9.7 and 9.9% Zn. The same applies to each other numerical property,

thermal treatment practice (i.e. temperature) and/or elemental range set forth herein.

Maximum or "max" refers to a total value up to the stated value for elements, times

and/or other property values, as in a maximum of 0.04 wt.% Cr; and minimum; "min"

refers to all values above the stated minimum value.

[0026] The term "incidental elements" can include relatively small amounts of Ti,

B, and others. For example, titanium with either boron or carbon serves as a casting aid,

for grain size control. The invention herein may accommodate up to about 0.06 wt.% Ti,

or about 0.01 to 0.06 wt.% Ti and optionally up to: about 0.001 or 0.03 wt.% Ca, about 0.03 wt.% Sr and/or about 0.002 wt.% Be as incidental elements. Incidental elements can

also be present in significant amounts and add desirable or other characteristics on their

own without departing from the scope ofthe invention so long as the alloy retains the

desirable characteristics set forth herein, including reduced quench sensitivity and

improved property combinations.

[0027] This alloy can further contain other elements to a lesser extent and on a less

preferred basis. Chromium is preferably avoided, i.e. kept at or below about 0.1 wt.% Cr. •

Nevertheless, it is possible that some very small amounts of Cr may contribute some

value for one or more specific applications of this invention alloy. Presently preferred

embodiments keep Cr below about 0.05 wt.%. Manganese is also kept purposefully low,

below about 0.2 or 0.3 total wt.% Mn, and preferably not over about 0.05 or 0.1 wt.%

Mn. Still, there may be one or more specific applications of this invention alloy where

purposeful Mn additions may make a positive contribution.

[0028] For the alloy, minor amounts of calcium may be incorporated therein,

primarily as a good deoxidizing element at the molten metal stages. Ca additions of up to

about 0.03 wt.%, or more preferably about 0.001-0.008 wt.% (or 10 to 80 ppm) Ca, also

assist in preventing larger ingots cast from the aforesaid composition from cracking

unpredictably. When cracking is less critical, as for round billets for forged parts and or

extrusions, Ca need not be added hereto, or may be added in smaller amounts. Strontium

(Sr) can be used as a substitute for, or in combination with the aforesaid Ca amounts for

the same purposes. Traditionally, beryllium additions has served as a deoxidizer/ingot cracking deterrent. Though for environmental, health and safety reasons, more preferred

embodiments of this invention are substantially Be-free.

[0029] Iron and Silicon contents should be kept significantly low, for example, not

exceeding about 0.04 or 0.05 wt.% Fe and about 0.02 or 0.03 wt. % Si or less. In any

event, it is conceivable that still slightly higher levels of both impurities, up to about 0.08

wt.%) Fe and up to about 0.06 wt.% Si may be tolerated, though on a less preferred basis

herein. Even less preferred, but still tolerable, Fe levels of about 0.15 wt.% and Si levels

as high as about 0.12 wt.% may be present in the alloy of this invention. For the mold

plates embodiments hereof, even higher levels of up to about 0.25 wt.% Fe, and about

0.25 wt.% Si or less, are tolerable.

[0030] As is known in the art of 7XXX Series, aerospace alloys, iron can tie up

copper during solidification. Hence, there are periodic references throughout this

disclosure to an "Effective Cu" content, that is the amount of copper NOT tied up by iron

present, or restated, the amount of Cu actually available for solid solution and alloying.

In some instances, therefore, it can be advantageous to consider the effective amount of

Cu and/or Mg present in the invention, then correspondingly adjust (or raise) the range of

actual Cu and/or Mg measured therein to account for the levels of Fe and/or Si contents

present and possibly interfering with Cu, Mg or both. For example, raising the preferred

amount of Fe content acceptable from about 0.04 or 0.05 wt % to about 0.1 wt.%

maximum can make it advantageous to raise the actual, measurable Cu minimums and

maximums specified by about 0.13 wt.%. Manganese acts in a similar manner to copper

with iron present. Similarly for magnesium, it is known that silicon ties up Mg during the solidification of 7XXX Series alloys. Hence, it can be advantageous to refer to the

amount of Mg present in this disclosure as an "Effective Mg" by which is meant that

amount of Mg not tied up by Si, and thus available for solution at the temperature or

temperatures used for solutionizing 7XXX alloys. Like the aforesaid actual adjusted Cu

ranges, raising the preferred allowable maximum Si content from about 0.02 to about 0.08

or even 0.1 or 0.12 wt.% Si could cause the acceptable/measurable amounts (both max

and min) of Mg present in this invention alloy to be similarly adjusted upwardly, perhaps

on the order of about 0.1 to 0.15 wt.%.

[0031] A narrowly stated composition according to this invention would contain

about 6.4 or 6.9 to 8.5 or 9 wt.% Zn, about 1.2 or 1.3 to 1.65 or 1.68 wt.% Mg, about 1.2

or 1.3 to 1.8 or 1.85 wt.% Cu and about 0.05 to 0.15 wt.% Zr. Optionally, the latter

composition may include up to 0.03, 0.04 or 0.06 wt.% Ti, up to about 0.4 wt.% Sc, and

up to about 0.008 wt.% Ca.

[0032] Still more narrowly defined, the presently preferred compositional ranges of

this invention contain from about 6.9 or 7 to about 8.5 wt.% Zn, from about 1.3 or 1.4 to

about 1.6 or 1.7 wt.% Mg, from about 1.4 to about 1.9 wt.%> Cu and from about 0.08 to

0.15 or 0.16 wt.% Zr. The % Mg does not exceed (% Cu + 0.3), preferably not exceeding

(% Cu + 0.2), or better yet (% Cu + 0.1). For the foregoing preferred embodiments, Fe

and Si contents are kept rather low, at or below about 0.04 or 0.05 wt.% each. A

preferred composition contains: about 7 to 8 wt.% Zn, about 1.3 to 1.68 wt.% Mg and

about 1.4 to 1.8 wt.% Cu, with even more preferably wt.% Mg wt.% Cu, or better yet Mg

< Cu. It is also preferred that the magnesium and copper ranges of this invention, when combined, not exceed about 3.5 wt.% total, with wt.% Mg + wt.% Cu about 3.3 on a

more preferred basis.

[0033] The alloys ofthe present invention can be prepared by more or less

conventional practices including melting and direct chill (DC) casting into ingot form.

Conventional grain refiners such as those containing titanium and boron, or titanium and

carbon, may also be used as is well-known in the art. After conventional scalping (if

needed) and homogenization, these ingots are further processed by, for example, hot

rolling into plate or extrusion or forging into special shaped sections. Generally, the

thick 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 cross section. In the case of plate about 4 to 8 inches

thick, the aforementioned plate is solution heat treated (SHT) and quenched, then

mechanically stress relieved such as by stretching and/or compression up to about 8%, for

example, from about 1 to 3%. A desired structural shape is then machined from these heat

treated plate sections, more often generally after artificial aging, to form the desired shape

for the part, such as, for example, an integral wing spar. Similar SHT, quench, often

stress relief operations and artificial aging are also followed in the manufacture of thick

sections made by extrusion and/or forged processing steps.

[0034] 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 ofthe product also increases. Hence, the alloy ofthe present invention

finds particular utility in thick gauges of, for example, greater than 2 to 3 inches in

thickness up to 12 inches or more. DESCRIPTION OF THE DRAWINGS

[0035] Figure 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;

[0036] Figure 2 is a graph showing two calculated cooling curves to approximate

the mid-plane cooling rates for plant made, 6- and 8-inch thick plates under spray

quenching, over which two experimental cooling curves, simulating the cooling rates of a

6-inch thick and an 8-inch thick plate, are superimposed;

[0037] Figure 3 is a graph showing longitudinal tensile yield strength TYS (L)

versus longitudinal fracture toughness K_q (L-T) relations for selected alloys ofthe present

invention and other alloys including 7150 and 7055 type comparisons or "controls", all

based on simulation of mid-plane (or "T/2") quench rates for a 6-inch thick plate,

extrusion or forging;

[0038] Figure 4 is a graph similar to Figure 3 showing longitudinal tensile yield

strength TYS (L) versus fracture toughness K_q (L-T) relations for selected alloys ofthe

present invention and other alloys including 7150 and 7055 controls, all based on

simulation of mid-plane quench rates for an 8-inch thick plate, extrusion or forging;

[0039] Figure 5 is a graph showing the influence of Zn content on quench

sensitivity as demonstrated by directional arrows for TYS changes in a 6-inch thick plate

quench simulation; [0040] Figure 6 is a graph showing the influence of Zn content on quench

sensitivity as demonstrated by directional arrows for TYS changes in an 8-inch thick plate

quench simulation;

[0041] Figure 7 is a graph showing cross plots of TYS (L) versus plane-strain

fracture^' toughness K_Ic (L-T) values at quarter plane (T/4) of a full-scale production 6-inch

thick plate ofthe invention alloy with the currently extrapolated minimum value line

(M-M) drawn thereon for comparing with literature reported values for 7050 and 7040

aluminum;

[0042] Figure 8 is a graph showing the influence of section thickness on TYS

values, as an index of quench sensitivity property, from a full-scale production, die-

forging study comparing alloys ofthe invention versus 7050 aluminum;

[0043] Figure 9 is a graph comparing longitudinal TYS values (in ksi) versus

electrical conductivity EC (as % IACS) for samples from 6 inch thick plate ofthe

invention alloy after aging by a known 2-step aging method versus the preferred 3 -step

aging practice outlined below. Most notable from this Figure is the surprising and

significant strength increase observed at same EC level, or the significant EC level

increases observed at the same strength value, for 3 -step aged samples as compared to

their 2-step aged counterparts. In each case, the first step age was conducted at 225°F,

250°F or at both temperatures, followed by a second step age at about 310°F; [0044] Figure 10 is a graph depicting the Seacoast SCC performance of 2- versus

3-stage aged for one preferred alloy composition at various short transverse (ST) stress

levels, a visual summary ofthe data found at Table 9 below;

[0045] Figure 11 is a graph depicting the Seacoast SCC performance of 2- versus

3-step aged for a second preferred alloy composition at various short transverse (ST)

stress levels, a visual summary ofthe data found at Table 10 below;

[0046] Figure 12 is a graph plotting open hole fatigue life, in the L-T orientation,

for various sized plate samples ofthe invention, from which a 95% confidence S/N band

(dotted lines) and a currently extrapolated preferred minimum performance (solid line

A-A) were drawn and compared with one jetliner manufacturer's specified values for

7040/7050-T7451 and 7010/7050-T7451 plate product, albeit in a different (T-L)

orientation;

[0047] Figure 13 is a graph plotting open hole fatigue life, in the L-T orientation,

for various sized forgings ofthe invention, from which a mean value line (dotted) and a

currently extrapolated preferred minimum performance (solid line B-B) were drawn; and

[0048] Figure 14 is a graph plotting fatigue crack growth (FCG) rate curves, in the

L-T and T-L orientations, for various sized plate and forgings ofthe invention, from

which a currently extrapolated, FCG preferred maximum curve (solid line C-C) was

drawn and compared with the FCG curves specified by one jetliner manufacturer for the

same size range 7040/7050-T7451 commercial plate of Figure 12 in the same (L-T and T-

L) orientations.

PREFERRED EMBODIMENTS [0049] 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 important are fracture toughness, both plane-strain

and plane-stress, and corrosion resistance performance such as exfoliation and stress

corrosion cracking resistance, and fatigue, both smooth and open-hole fatigue life (S/N)

and fatigue crack growth (FCG) resistance.

[0050] As described above, integral wing spars, ribs, webs, and wing skin panels

with integral stringers, can be machined from thick plates or other extruded or forged

product forms which have been solution heat treated, quenched, mechanically stress

relieved (as needed) and artificially aged. It is not always feasible to solution heat treat

and rapidly quench the finished structural component itself because the rapid cooling

from quenching may induce residual stress and cause dimensional distortions. Such

quench-induced residual stresses can also cause stress corrosion cracking. Likewise,

dimensional distortions due to rapid quenching may necessitate re-working to straighten

parts that have become so distorted as to render standard assembly impracticably difficult.

Other representative aerospace parts/products that can be made from this invention

include, but are not limited to: large frames and fuselage bulkheads for commercial jet

airliners, hog out plates for the upper and lower wing skins of smaller, regional jets,

landing gear and floor beams for various jet aircraft, even the bulkheads, fuselage

components and wing skins of fighter plane models. In addition, the alloy of this invention can be made into miscellaneous small forged parts and other hogged out

structures of aircraft that are currently made from alloy 7050 or 7010 aluminum.

[0051] While it is easier to obtain better mechanical properties in thin cross

sections (because the faster cooling of such parts prevents unwanted precipitation of

alloying elements), rapid quenching can cause excessive quench distortion. To the extent

practical, such parts may be mechanically straightened and/or flattened while residual

stress relief practices are performed thereon after which these parts are artificially aged.

[0052] As indicated above, in solution heat treating and quenching thick sections,

the quench sensitivity ofthe aluminum alloy is of great concern. After solution heat

treating, it is desirable to quickly cool the material for retaining various alloying elements

in solid solution rather than allowing them to precipitate out of solution in coarse form as

otherwise occurs via slow cooling. The latter occurrence produces coarse precipitates and

results in a decline in mechanical properties. In products with thick cross sections, i.e.

over 2 inches thick at its greatest point, and more particularly, about 4 to 8 inches thick or

more, the quenching medium acting on exterior surfaces of such workpieces (either plate,

forging or extrusion) cannot efficiently extract heat from the interior including the center

(or mid-plane (T/2)) or quarter-plane (T/4) regions of that material. This is due to the

physical distance to the surface and the fact that heat extracts through the metal by a

distance dependent conduction. In thin product cross sections, quench rates at the mid-

plane are naturally higher than quench rates for a thicker product cross sections. Hence,

an alloy's overall quench sensitivity property is often not as important in thinner gauges

as it is for thicker gauged parts, at least from the standpoint of strength and toughness. [0053] The present invention is primarily focused on increasing the strength-

toughness properties in a 7XXX series aluminum alloy in thicker gauges, i.e. greater than

about 1.5 inches. The low quench sensitivity ofthe invention alloy is of extreme

importance. In thicker gauges, the less quench sensitivity the better with respect to that

material's ability to retain alloying elements in solid solution (thus avoiding the formation

of adverse precipitates, coarse and others, upon slow cooling from SHT temperatures)

particularly in the more slowly cooling mid- and quarter-plane regions of said thick

workpiece. This invention achieves its desired goal of lowering quench sensitivity by

providing a carefully controlled alloy composition which permits quenching thicker

gauges while still achieving superior combinations of strength-toughness and corrosion

resistance performance.

[0054] To illustrate the invention, twenty-eight, 11 -inch diameter ingots were

direct chill (or DC) cast, homogenized and extruded into 1.25 x 4 inch wide rectangular

bars. Those bars were all solution heat treated before being quenched at different rates to

simulate cooling conditions for thin sections as well as for approximating conditions for

the mid-plane of 6- and 8-inch thick workpiece sections. These rectangular test bars were

then cold stretched by about 1.5% for residual stress relief. The compositions of alloys

studied are set forth in Table 2 below, in which Zn contents ranged from about 6.0 wt. %

to slightly in excess of 11.0 wt.%. For these same test specimens, Cu and Mg contents

were each varied between about 1.5 and 2.3 wt.%.

For all alloys other than the controls: Target Si = 0.03, Fe = 0.05, Zr = 0.12, Ti = 0.025 For 7150 Control (Sample # 27): Target Si = 0.05, Fe = 0.10, Zr = 0.12, Ti = 0.025 For 7055 Control (Sample # 28): Target Si = 0.07, Fe = 0.11, Zr = 0.12, Ti = 0.025

[0055] Different quenching approaches were explored to obtain, at the mid-plane

of a 1.25 inch thick extruded bar, a cooling rate simulating that at the mid-plane of a 6-

inch thick plate spray quenched in 75°F water as would be the case in full-scale production. A second set of data involved simulating, under identical circumstances, a

bar cooling rate corresponding to that of an 8-inch thick plate.

[0056] The aforesaid quenching simulation involved modifying the heat transfer

characteristics of quenching medium, as well as the part surface, by immersion quenching

extruded bars via the simultaneous incorporation of three known quenching practices: (i)

a defined warm water temperature quench; (ii) saturation ofthe water with C0₂ gas; and

(iii) chemically treating the bars to render a bright etch surface finish to lower surface

heat transfer.

[0057] For simulating the 6-inch thick plate cooling condition: the water

temperature for immersion quenching was held at about 180°F; and the solubility level of

C0₂ in the water kept at about 0.20 LAN (a measure of dissolved C0₂ concentration,

LAN = standard volume of C0₂/volume of water). Also, the sample surface was

chemically treated to have a standard, bright etch finish.

[0058] For the 8-inch thick plate cooling simulation, the water temperature was

raised to about 190°F with a C0₂ solubility reading varying between 0.17 and 0.20 LAN.

Like the 6 inch samples above, this thicker plate was chemically treated to have a

standard bright etch surface finish.

[0059] The cooling rates were measured by thermocouples inserted into the

mid-plane of each bar sample. For benchmark reference, the two calculated cooling

curves to approximate the mid-plane cooling rates under spray quenching at plant-made

6- and 8-inch thick plates were plotted per accompanying Figure 2. Superimposed on them were displayed two groups of plots, the lower group (in the temperature scale)

representing simulated cooling rate curves mid-plane of a 6-inch thick plate; and the

upper, simulated mid-plane for an 8-inch thick plate. These simulated cooling rates were

very similar to those of plant production plates in the important temperature range above

about 500°F, although the simulated cooling curves for experimental materials differed

from those for plant plate below 500°F, which was not considered critical.

[0060] After solution heat treating and quenching, artificial aging behaviors were

studied using multiple aging times to. obtain acceptable electrical conductivity ("EC") and

exfoliation corrosion resistance ("EXCO") readings. The first two-step aging practice for

the invention alloy consisted of: a slow heat-up (for about 5 to 6 hours) to about 250°F, a

4 to 6 hour soak at about 250°F, followed by a second step aging at about 320°F for

varying times ranging from about 4 to 36 hours.

[0061] Tensile and compact tension plane-strain fracture toughness test data were

then collected on samples given the different minimum aging times required to obtain a

visual EXCO rating of EB or better (EA or pitting only) for acceptable exfoliation

corrosion resistance performance, and an electrical conductivity EC minimum value of at

or above about 36% IACS (International Annealed Copper Standard), the latter value

being used to indicate degree of necessary over-aging and provide some indication of

corrosion resistance performance enhancement as is known in the art. All tensile tests

were performed according to the ASTM Specification E8, and all plane-strain fracture

toughness per ASTM specification E399, said specifications being well known in the art. [0062] Figure 3 shows the plotted strength-toughness results from Table 2 alloy

samples slowly quenched from their SHT temperatures for simulating a 6-inch thick

product. One family of compositions noticeably stood out from the rest of those plotted,

namely sample numbers 1, 6, 11 and 18 (in the upper portions of Figure 3). All of those

sample -numbers-displayed very high fracture toughness combined with high strength

properties. Surprisingly, all of those sample alloy compositions belonged to the low Cu

and low Mg ends of our choice compositional ranges, namely, at around 1.5 wt.% Mg

together with 1.5 wt.% Cu, while the Zn levels therefor varied from about 6.0 to 9.5

wt.%. Particular Zn levels for these improved alloys were measured at: 6 wt.% Zn for

Sample #1, 7.6 wt.% Zn for Sample #6, 8.7 wt.% Zn for Sample #11 and 9.4 wt.% Zn for

Sample #18.

[0063] Substantial improvements in strength and toughness can also be seen when

the aforementioned alloy performances are compared against two "control" alloys 7150

aluminum (Sample # 27 above) and 7055 aluminum (Sample #28) both of which were

processed in an identical manner (including temper). In Figure 3, a drawn dotted line

connects the latter two control alloy data points to show their "strength-toughness

property trend" whereby higher strength is accompanied by lower toughness

performance. Note how the Figure 3 line for control alloys 7150 and 7055 extends

considerably below the data points discussed for invention alloy Sample Nos. 1, 6, 11 and

18 above.

[0064] Also included in the Figure 3 plots are results for alloys having about 1.9

wt.% Mg and 2.0 wt.% Cu with various Zn levels: 6.8 wt.% (For Sample #5), 8.2 wt.% (for Sample #10), 9.0 wt.% (for Sample #17) and 10.2 wt.% (for Sample #26). Such

results once again graphically illustrate the drop in toughness observed for these alloys

compared to 1.5 wt.% Mg and 1.5 wt.% Cu containing alloys at corresponding levels of

total Zn. And while the thick gauge, strength-toughness properties for higher Mg and Cu

alloy products were similar to or marginally better than those for the 7150 and 7055 ^■

controls (dotted trend line), such results clearly demonstrate a significant degradation in

both strength and toughness properties that occurs with a moderate increase in Cu and

Mg: (1) above the Cu and Mg levels ofthe present invention alloy, and (2) approaching

the Cu/Mg levels of many current commercial alloys.

[0065] A similar set of results are graphically depicted in accompanying Figure 4

for a quench condition even slower than that shown and described for above Figure 3.

The Figure 4 conditions roughly approximate those for an 8-inch thick plate, mid-plane

cooling condition. Similar conclusions as per Figure 3 can be drawn for the data

depicted in Figure 4 for a still slower quench simulation performed to represent a still

thicker plate product.

[0066] Thus, unlike past teachings, some ofthe highest strength-toughness

properties were obtained at some ofthe leanest Cu and Mg levels used thus far for current

commercial aerospace alloys. Concomitantly, the Zn levels at which these properties

were most optimized correspond to levels much higher than those specified for 7050,

7010 or 7040 aluminum plate products.

[0067] It is believed that a good portion ofthe improvement in strength and

toughness properties observed for thick sections ofthe invention alloy are due to the specific combination of alloy ingredients. For instance, the accompanying Figure 5 TYS

strength values increase gradually with increasing Zn content, from Sample #1 to Sample

#6 to Sample #11 and are superior to the prior art "controls". Thus, unlike past teachings,

higher Zn solutes do not necessarily increase quench sensitivity if the alloy is properly

formulated as provided herein. On the contrary, the higher Zn levels of this invention

have actually proven to be beneficial against the slow quench conditions of thick

sectioned workpieces. At still higher Zn levels of 9.4 wt.%, however, the strength can

drop. Hence, the TYS strength of Sample #18 (containing 9.42 wt.% Zn) drops below

those for the other, lower Zn invention alloys in Figure 5.

[0068] In accompanying Figure 6, still further, slower quench conditions for

simulated 8-inch thicknesses are depicted. From that data, it can be seen that quench

sensitivity can increase even at 8.7 wt.% Zn levels, as depicted by the TYS strength

values for Sample #11 displaced below that for Sample #6's total Zn content of 7.6 wt.%.

This high solute effect on quench sensitivity is also evidenced by the relative positions of

control alloys 7150 (Sample #27) and 7055 (Sample #28) on the TYS strength axes ofthe

accompanying figures. Therein, 7055 was stronger than 7150 under slow quench (Figure

5), but the relative scale was reversed under still slower quench conditions (per Figure 6).

[0069] Also noteworthy is the performance of Sample #7 above, which according

to Table 2 contained 1.59 wt.% Cu, 2.30 wt.% Mg and 7.70 wt.% Zn, (so that its Mg

content exceeded Cu content). From Figure 3, that Sample exhibited high TYS strengths

of about 73 ksi but with a relatively low fracture toughness, K_Q(L-T), of about 23 ks in.

By comparison, Sample #6, which contained 7.56% Zn, 1.57% Cu and 1.51% Mg (with Mg < Cu) exhibited a Figure 3 TYS strength greater than 75 ksi and a higher fracture

toughness of about 34 ks in (actually a 48% increase in toughness). This comparative

data shows the importance of: (1) maintaining Mg content at or below about 1.68 or

1.7wt.%, as well as (2) keeping said Mg content less than or equal to the Cu content + 0.3

wt.%, and more preferably below the Cu content, or at a minimum, not above the Cu

content ofthe invention alloy.

[0070] It is desirable to achieve optimum and/or balanced fracture toughness (K_Q)

and strength (TYS) properties in the alloys of this invention. As can be best seen and

appreciated by comparing the compositions of Table 2 with their corresponding fracture

toughness and strength values plotted in Figure 3, those alloy samples falling within the

compositions of this invention achieve such a balance of properties. Particularly, those

Sample Nos. 1, 6, 11 and 18 either possess a fracture toughness value (K_Q) (L-T) in

excess of about 34 ksWin with a TYS greater than about 69 ksi; or they possess a fracture

toughness value greater than about 29 ksWin combined with a higher TYS of about 75 ksi

or greater.

[0071] The upper limit of Zn content appears to be important in achieving the

proper balance between toughness and strength properties. Those samples which

exceeded about 11.0 wt.%, such as Sample Nos. 24 (11.08 wt.% Zn) and 22 (11.38 wt.%

Zn), failed to achieve the minimum combined strength and fracture toughness levels set

forth above for alloys ofthe invention. [0072] The preferred alloy compositions herein thus provide high damage

tolerance in thick aerospace structures resulting from its enhanced, combined fracture

toughness and yield strength properties. With respect to some ofthe property values

reported herein, one should note that K_Q values are the result of plane strain fracture

toughness tests that do not conform to the current validity criteria of ASTM Standard

E399. In the current tests that yield K_Q values, the validity criteria that were not precisely

followed were: (1) P_MAX / P_Q <1.1 primarily, and (2) B (thickness) > 2.5 (K_Q/σ_YS)²

occasionally, where K_Q, σ_γs, P_MAX. ^and P_Q ^{are as} defined in ASTM Standard E399-90.

These differences are a consequence ofthe high fracture toughnesses observed with the

invention alloy. To obtain valid plane-strain K_lc results, a thicker and wider specimen

would have been required than is facilitated with an extruded bar (1.25 inch thick x 4 inch

wide). A valid K_lc is generally considered a material property relatively independent of

specimen size and geometry. K_Q, on the other hand, may not be a true material property

in the strictest academic sense because it can vary with specimen size and geometry.

Typical K_Q values from specimens smaller than needed are conservative with respect to

K_lc, however. In other words, reported fracture toughness (K_Q) values are generally

lower than standard K_lc values obtained when the sample size related, validity criteria of

ASTM Standard E399-90 are satisfied. The K_Q values were obtained herein using

compact tension test specimens per ASTM E399 having a thickness B of 1.25 inch and

width that varied between 2.5 to 3.0 inches for different specimens. Those 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 of ASTM Standard E399 for K_lc were conducted using compact tension specimens with a

thickness, B = 2.0 inch, and width, W = 4.0 inch. Those specimens were fatigue pre-

cracked to a crack length of 2.0 inch (A/W = 0.5). All cases of comparative data between

varying alloy compositions were made using results from specimens ofthe same size and

under similar test conditions.

EXAMPLE 1: PLANT TRIAL - PLATE

[0073] A plant trial was conducted using a standard, full-size ingot cast with the

following invention alloy composition: 7.35 wt.% Zn, 1.46 wt.% Mg, 1.64 wt.% Cu, 0.04

wt.% Fe, 0.02 wt.%) Si and 0.11 wt.% Zr. That ingot was scalped, homogenized at 885°

to 890°f for 24 hours, and hot rolled to 6-inch thick plate. The rolled plate was then

solution heat treated at 885° to 890°F for 140 minutes, spray quenched to ambient

temperature, and cold stretched from about 1.5 to 3% for residual stress relief. Sections

from that plate were subjected to a two-step aging practice that consisting of a 6-

hour/250°f first step aging followed by a second step age at 320°F for 6, 8 and 11 hours,

respectively designated as times TI, T2 and T3 in the table that follows. Results from the

tensile, fracture toughness, alternate immersion SCC, EXCO and electrical conductivity

tests are presented in Table 3 below. Figure 7 shows the cross plot of L-T plane-strain

fracture toughness (K_lc ) versus longitudinal tensile yield strength TYS (L), both samples

having been taken from the quarter-plane (T/4) location ofthe plate. A linear strength-

toughness correlation trend (Line T3-T2-T1) was drawn to define through the data for

these representative, second stage aging times. A preferred minimum performance line

(M-M) was also drawn. Also included in Figure 7 are the 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 plate per AMS D99AA draft specification (ref.

Preliminary Materials Properties Handbook), both specifications being known in the art.

From this preliminary data on two step aged plate, the alloy compositions of this

invention clearly display a much superior strength-toughness combination compared to

either 7050 or 7040 alloy plate. In comparison to 7050-T7451 plate, for example, the two

step aged versions of this invention achieved a TYS increase of about 11% (72 ksi versus

64 ksi), at the equivalent K_lc of 35 ksi/in. Stated differently, significant increases in K_lc

values were obtained with the present invention at equivalent TYS levels. For example,

the two step aged versions of this plate product achieved a 28% K_lc (L-T) toughness

increase (32.3 ksi/in versus 41 ksi/in) as compared to its 7040-T7451 equivalent at the

same TYS (L) level of 66.6 ksi.

EXAMPLE 2; PLANT TRIAL - FORGING

[0074] A die forged evaluation ofthe invention alloy was performed in a plant-trial

using two full-size production sheet/plate ingots, designated COMP1 and COMP2, as

follows:

COMP 1: 7.35 wt.% Zn, 1.46 wt.% Mg, 1.64 wt.% Cu, 0.11 wt.% Zr, 0.038 wt.% Fe, 0.022 wt.% Si, 0.02 wt.% Ti;

COMP 2: 7.39 wt.% Zn, 1.48 wt.% Mg, 1.91 wt.% Cu, 0.11 wt.% Zr, 0.036 wt.% Fe, 0.024 wt.% Si, 0.02 wt.% Ti.

A standard 7050 ingot was also run as a control. All ofthe aforesaid ingots were

homogenized at 885°F for 24 hours and sawed to billets for forging. A closed die, forged

part was produced for evaluating properties at three different thicknesses, 2 inch, 3 inch

and 7 inch. The fabrication steps conducted on these metals included: two pre-forming

operations utilizing hand forging; followed by a blocker die operation and a final finish die operation using a 35,000 ton press. The forging temperatures employed therefor were

between about 725 - 750°F. All the forged pieces were then solution heat treated at 880°

to 890°F for 6 hours, quenched and cold worked 1 to 5% for residual stress relief. The

parts were next given a T74 type aging treatment for enhancing SCC performance. The

aging treatment consisted of 225°F for 8 hours, followed by 250°F for 8 hours, then

350°F for 8 hours. Results from the tensile tests performed in longitudinal, long-

transverse and short-transverse directions are presented in accompanying Figure 8. In all

three orientations, the tensile yield strength (TYS) values for the invention alloy remained

virtually unchanged for thicknesses ranging from 2 to 7 inches. In contrast, the

specification for 7050 allows a drop in TYS values as thickness increased from 2 to 3 to 7

inches consistent with the known performance of 7050 alloy. Thus, Figure 8 results

clearly demonstrate this invention's advantage of low quench sensitivity, or restated, the

ability of forgings made from this alloy to exhibit an insensitivity to strength changes

over a large thickness range in contrast to the comparative strength property dropoff

observed with thicker sections of prior art 7050 alloy forgings.

[0075] The present invention clearly runs counter to conventional 7XXX series

alloy design philosophies which indicate that higher Mg contents are desirable for high

strength. While that may still be true for thin sections of 7XXX aluminum, it is not the

case for thicker product forms because higher Mg actually increases quench sensitivity

and reduces the strength of thick sections.

[0076] Although the primary focus of this invention was on thick cross sectioned

product quenched as rapidly as practical, those skilled in the art will recognize and appreciate that another application hereof would be to take advantage ofthe invention's

low quench sensitivity and use an intentionally slow quench rate on thin sectioned parts

to reduce the quench-induced residual stresses therein, and the amount/degree of

distortion brought on by rapid quenching but without excessively sacrificing strength or

toughness.

[0077] Another potential application arising from the lower quench sensitivities

observed with this invention alloy is for products having both thick and thin sections such

as die forgings and certain extrusions. Such products should suffer less from yield

strength differences between thick and thin cross sectioned areas. That, in turn, should

reduce the chances of bowing or distortion after stretching.

[0078] Generally, for any given 7XXX series alloy, as further artificial aging is

progressively applied to a peak strength, T6-type tempered product (i.e. "overaging"), the

strength of that product has been known to progressively and systematically decrease

while its fracture toughness and corrosion resistance progressively and systematically

increase. Hence, today's part designers have learned to select a specific temper condition

with a compromise combination of strength, fracture toughness and corrosion resistance

for a specific application. Indeed, such is the case for the alloy ofthe invention, as

demonstrated in the cross plot of L-T plane strain fracture toughness K_lc and L tensile

yield strength, in Figure 7, both measured at quarter-plane (T/4) in the longitudinal

direction for 6-inch thick plate product. Figure 7 illustrates how the alloy of this

invention provides a combination of: about 75 ksi yield strength with -about 33 ksWin

fracture toughness, at the TI aging time from Table 3; or about 72 ksi yield sfrength with about 35 ksWin fracture toughness, with Table 3 - aging time T2; or about 67 ksi yield

strength and about 40 ksWin fracture toughness, with Table 3 - aging time T3.

[0079] It is further understood by those skilled in the art that, within limits, for a

specific 7XXX series alloy, the strength- fracture toughness trend line can be interpolated

and, to some extent, extrapolated to combinations of sfrength and fracture toughness

beyond the three examples of invention alloy given above and plotted at Figure 7. The

desired combination of multiple properties can then be accomplished by selecting the

appropriate artificial aging treatment therefor.

[0080] While the invention has been described largely with respect to aerospace

structural applications, it is to be understood that its end use applications are not

necessarily limited to same. On the contrary, the invention alloy and its preferred three

stage aging practice herein are believed to have many other, non-aerospace related end

use applications as relatively thick cast, rolled plate, extruded or forged product forms,

especially in applications that would require relatively high strengths in a slowly

quenched condition from SHT temperatures. An example of one such application is mold

plate, which must be extensively machined into molds of various shapes for the shaping

and/or contouring processes of numerous other manufacturing processes. For such

applications, desired material characteristics are both high strength and low machining

distortion. When using 7XXX alloys as mold plates, a slow quench after solution heat

treatment would be necessary to impart a low residual stress, which might otherwise

cause machining distortions. Slow quenching also results in lowered sfrength and other properties for existing 7XXX series alloys due to their higher quench sensitivity. It is the

unique very low quench sensitivity for this invention alloy that permits a slow quench

following SHT while still retaining relatively high strength capabilities that makes this

alloy an attractive choice for such non-aerospace, non-structural applications as thick

mold plate. For this particular application, though, it is not necessary to perform the

preferred 3 step aging method described hereinbelow. Even a single step, or standard 2

step, aging practice should suffice. The mold plate can even be a cast plate product.

[0081] The instant invention substantially overcomes the problems encountered in

the prior art by providing a family of 7000 Series aluminum alloy products which exhibits

significantly reduced quench sensitivity thus providing significantly higher strength and

fracture toughness levels than heretofore possible in thick gauge aerospace parts or parts

machined from thick products. The aging methods described herein then enhance the

corrosion resistance performance of such new alloys. Tensile yield strength (TYS) and

electrical conductivity EC measurements (as a % IACS) were taken on representative

samples of several new 7XXX alloy compositions and comparative aging processes

practiced on the present invention. The aforesaid EC measurements are believed to

correlate with actual corrosion resistance performance, such that the higher the EC value

measured, the more corrosion resistant that alloy should be. As an illustration,

commercial 7050 alloy is produced in three increasingly corrosion resistant tempers: T76

(with a typical SCC minimum performance, or "guarantee", of about 25 ksi and typical

EC of 39.5% IACS); T74 (with a typical SCC guarantee of about 35 ksi and 40.5%

IACS); and T73 (with it typical SCC guarantee of about 45 ksi and 41.5% IACS). [0082] In aerospace, marine or other structural applications, it is quite customary

for a structural and materials engineer to select materials for a particular component based

on the weakest link failure mode. For example, because the upper wing alloy of an

aircraft is predominantly subjected to compressive stresses, it has relatively lower

requireh ents for SCC resistance involving tensile stresses. As such, upper wing skin

alloys and tempers are usually selected for higher strength albeit with relatively low

short-transverse SCC resistance. Within that same aerospace wing box, the spar members

are subjected to tensile stresses. Although the structural engineer would desire higher

strengths for this application in the interest of component weight reduction, the weakest

link is the requirement of high SCC resistance for those component parts. Today's spar

parts are thus traditionally manufactured from a more corrosion resistant, but lower

strength alloy temper such as T74. Based on the observed EC increase at the same

strength, and the Al SCC test results described above, the preferred, new 3 stage aging

methods of this invention can offer these structural/materials engineers and aerospace part

designers a method of providing the strength levels of 7050/7010/7040-T76 products with

near T74 corrosion resistance levels. Alternatively, this invention can offer the corrosion

resistance of a T76 tempered material in combination with significantly higher strength

levels.

EXAMPLES:

[0083] Three representative compositions ofthe new 7xxx alloy family were cast

to target as large, commercial scale ingots with the following compositions:

Those cast ingot materials, of course after working, i.e. rolling to 6 inch finish gauge

plate, solution heat treating, etc., were subjected to the comparative aging practice

variations set forth in Table 5 below. Actually, two different first stages were compared

in this 3 stage evaluation, one having a single exposure at 250°F with the other broken

into two sub-stages: 4 hours @ 225°F, followed by a second sub-stage of 6 hours @

250°F. This two sub-stage procedure is referred to herein as first a first stage treatment,

i.e., prior to the second stage treatment at about 310°F. In any event, no noticeable

difference in properties was observed between these two "types" of first stages, the lone

treatment at 250°F versus the split freatments at both 225 and 250°F. Hence, referring to

any stage herein embraces such variants.

Specimens from each six inch thick plate were then tested, with the averages for the two-

and three-step aged properties being measured as follows:

[0084] Figure 9 is a graph comparing the tensile yield strengths and EC values that

were used to provide the interpolated data presented in Table 6 above. Significantly, it

was noted that a dramatic increase in EC was observed for the above described, 3-stage

aged Alloys A, B or C at the same yield strength level. From that data, it was also noted

that a surprising and significant strength increase at the same EC level was observed for

the above described, 3-step aged conditions as compared to the 2-step, with the second of

each being performed at about 310°F. For example, the yield strength for the 2-step aged

Alloy A specimen at 39.5% IACS was 72.1 ksi. But, its TYS value increased to 75.4 ksi

when given a 3-step age according to the invention.

[0085] Al SCC studies were performed per ASTM Standard D- 1141 , by alternate

immersion, in a specified synthetic ocean water (or SOW) solution, which is more aggressive than the more typical 3.5% NaCl salt solution required by ASTM Standard

G44. Table 7 shows the results on various Alloy A, B and C samples (all in an ST

direction) with just 2-aging steps, the second step comprising various times (6, 8 and 11

hours) at about 320°F.

TABLE 7 Results of SCC Testbv Alternate Immersion of Plant Processed 6" Plates of Allovs A. B and C Receiving 2-Stage Aging after 121 Days Exposure to Synthetic Ocean Water

6 Hours @ 250°F Stress F/N(l) Days To Stress F/N(l) Days To Failure Stress F/N(l) Days To Failure EC TYS (1^st stage) plus: (ksi) Failure (ksi) (ksi) (% IACS) (ksi) (T/2) (T/2) (T/2) (Surf) (T/4)

Alloy A-T7X 6" Plate 6 Hr/320F 25 1/5 77d 35 4/5 10, 12, 21, 70d 40 5/5 6,7, 7, 27, 9 Id 41.2 74.9 40K@121d 10K@121d

8 Hr/320F 25 0/5 50K@121d 35 2/5 100, lOOd 40 3/5 13, 13, 50d 41.6 72.5 30K@121d 20K@121d

HHr/320F 25 0/5 50K@121d 35 0/5 50K@121d 40 0/5 50K@121d 42.9 67.2

Alloy B-T7X 6" Plate 6 Hr/320F 25 0/5 50K@121d 35 0/5 50K@121d 40 0/5 50K@121d 41.3 74.8

8 Hr/320F 25 0/5 50K@121d 35 0/5 50K@121d 40 0/5 50K@121d 41.7 73.1

HHr/320F 25 0/5 50K@121d 35 0/5 50K@121d 40 0/5 50K@121d 42.2 69.2

Alloy C-T7X 6" Plate 6 Hr/320F 25 1/5 13d 35 0/5 5 OK @ 121d 40 3/5 23, 26, 34d 40.9 75.3 40K@121d 20K@121d

8 Hr/320F 25 0/5 50K@121d 35 0/5 50K@121d 40 3/5 13, 19,35d 41.2 73.9 20K@121d

HHr/320F 25 0/5 50K@121d 35 0/5 5 OK @ 121d 40 0/5 50K@121d 42.2 69.2

Note: F/N(l) = Number of specimens failed over the number exposed

From this data, several SCC failures were observed following exposure for 121 days,

primarily as a function of short transverse (ST) applied stress, aging time and/or alloy.

[0086] Comparative Table 8 lists SCC results for just Alloys A and C (applied

sfress in the same ST direction) after having been aged for an additional 24 hours at

250°F, that is for a total aging practice that comprises: (1) 6 hours at 250°F; (2) 6, 8 or 11

hours at 320°F; and (3) 24 hours at 250°F.

TABLE 8

Results of SCC Test bv Alternate Immersion of Plant Processed 6" Plates of Allovs A and C Receivinε 3-Staεe Aεinε after 93 Days Exposure to Synthetic Ocean Water by Alternate Immersion ASTM D-1141-90

6 Hours @ 250°F Stress F /N(l) Days To Stress F /N(l) Days To Stress F / N(l) Days To Failure EC TYS (1^st stage) plus: (ksi) Failure (ksi) Failure (ksi) (% IACS) (ksi)

(T/2) (T/2) (T/2) (T/10) (T/4)

Alloy A-T7X Plate

6 Hr/320F + 24h 250F

25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 39.7 74.2

8 Hr/320F + 24h/250F 25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 40.4 72.1

H Hr/320F + 24h/250F 25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 41.5 67.4

Alloy C -T7X Plate

6 Hr/320F + 24h/250F

25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 39.5 75.3

8 Hr/320F + 24h/250F 25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 40.0 72.8

H Hr/320F + 24h 250F 25 0/3 3 OK @ 93d 35 0/3 3 OK @ 93d 45 0/3 3 OK @ 93d 41.0 68.8

Note: F/N(l) = Number of specimens failed over the number exposed.

[0087] Quite remarkably, no sample failures were observed under identical test

conditions after the first 93 days of exposure. Thus, the new 3-step aging approach of this

invention is believed to confer unique sfrength/SCC advantages surpassing those

achievable through conventional 2-step aging while promising to develop better property

attributes in new products and confer further property combination improvements in still

other, current aerospace product lines.

[0088] The value of comparing Table 7 data to that in Table 8 is to underscore that

while 2 stage/step aging may be practiced on the alloy according to this invention, the

preferred 3 stage aging method herein described actually imparts a measurable SCC test

performance improvement. Tables 6 and 7 also include SCC performance "indicator"

data, EC values (as a %IACS), along with correspondingly measured TYS (T/4) values.

That data must not be compared, side-by-side, for determining the relative value of a two

versus 3 step aged products, howeve,r as the EC testing was performed at different areas

ofthe product, i.e. Table 7 using surface measured values versus the T/10 meaurements

of Table 8 (it being known that EC indicator values generally decrease when measuring

from the surface going inward on a given test specimen). The TYS values cannot be used

as a true comparison either as lot sizes varied as well as testing location (laboratory

versus plant). Instead, the relative data of Figure 9 (below) should be consulted for

comparing to what extent 3 step aging showed an improved COMBINATION of strength

and corrosion resistance performance using longitudinal TYS values (ksi) versus electrical conductivity EC (% IACS) for side-by-side, commonly tested 6 inch thick plate

samples ofthe invention alloy.

[0089] Seacoast SCC test data confirms the significant improvements in corrosion

resistance realized by imparting a novel three-step aging method to the aforementioned

new family of 7XXX alloys. For the alloy composition identified as Alloy A in above

Table 4, SCC testing extended over a 568 day period for 2-stage aged versus a 328 day

test period for the 3 stage aged, with the comparative 2- versus 3-stage aged SCC

performances mapped per following Table 9 (The latter (3 stage) testing was started after

the former (2 stage) tests had commenced; hence, the longer test times observed for 2

stage aged specimens).

Note: 2 stage aging comprised: 6 hours @ 250°F; and 6 or 8 hours @ 320°F.

3 stage aging comprised: 6 hours @ 250°F; 7 or 9 hours @ 320°F; and 24 hours @ 250°F.

This data is graphically summarized in accompanying Figure 10 with the times in the

upper left key on that Figure always referring to the second step aging times at 320°F,

even for the 3 step aged specimens commonly referred to therein. [0090] A second composition, Alloy C in Table 4 (with its 7.4 wt.% Zn, 1.5 wt.%

Mg, 1.9 wt. % Cu, and 0.11 wt.% Zr), was subjected to the comparative 2- versus 3-step

agings as was Alloy A above. The long term results from those Seacoast SCC tests are

summarized in Table 10 below.

[0091] Graphically, this Table 10 data is shown in accompanying Figure 11 with

the times in the upper left key on that Figure always referring to the second step aging

times at 320°F, even for the 3 step aged specimens commonly referred to therein. From

both the Alloy A and Alloy C data, it is most evident that practicing the preferred 3-step aging process of this invention on its preferred alloy compositions imparts a significant

improvement in SCC Seacoast testing performance therefor, especially when the

specimen days-to-failure rates of 3-step aged materials are compared side-by-side to the

2-step aged counterparts. Prior to this prolonged SCC Seacoast testing, however, the 2-

step aged materials showed some SCC performance enhancements under simulated tests

and may be suitable for some applications ofthe invention alloy even though the

improved 3 step/stage aging is preferred.

[0092] With respect to the 3-stage aging, preferred particulars for the

aforementioned alloy compositions, one must note that: the first stage age should

preferably take place within about 200 to 275°F, more preferably between about 225 or

230 to 260°F, and most preferably at or about 250°F. And while about 6 hours at the

aforesaid temperature or temperatures is quite satisfactory, it must be noted that in any

broad sense, the amount of time spent for first step aging should be a time sufficient for

producing a substantial amount of precipitation hardening. Thus, relatively short hold

times, for instance of about 2 or 3 hours, at a temperature of about 250°F, may be

sufficient (1) depending on part size and shape complexity; and (2) especially when the

aforementioned "shortened" treatment/exposure is coupled with a relatively slow heat up

rate of several hours, for instance 4 to 6 or 7 hours, total.

[0093] The preferred second stage aging practice to be imparted on the preferred

alloy compositions of this invention can be purposefully ramped up directly from the

aforementioned first step heat treatment. Or, there may be a purposeful and distinct

time/temperature interruption between first and second stages. Broadly stated, this second step should take place within about 290 or 300 to 330 or 335°F. Preferably, this

second step age is performed within about 305 and 325°F. Preferably, second step aging

takes place between about 310 to 320 or 325°F. The preferred exposure times for this

critical second step processing depend somewhat inversely on the actual temperature(s)

employed. For instance, if one were to operate substantially at or very near 310°F, a total

exposure time from about 6 to 18 hours, preferably for about 7 to 13, or even 15 hours

would suffice. More preferably, second step agings would proceed for about 10 or 11,

even 13, total hours at that operating temperature. At a second aging stage temperature of

about 320°F, total second step times can range between about 6 to 10 hours with about 7

or 8 to 10 or 11 hours being preferred. There is also a preferred target property aspect to

second step aging time and temperature selection. Most notably, shorter treatment times

at a given temperature favor higher strength values whereas longer exposure times favor

better corrosion resistance performance.

[0094] Finally, with respect to the preferred, third aging practice stage, it is better

to not ramp slowly down from the second step for performing this necessary third step on

such thick workpieces unless extreme care is exercised to coordinate closely with the

second step temperature and total time duration so as to avoid exposures at second aging

stage temperatures for too long a time. Between the second and third aging steps, the

metaljproducts of this invention can be purposefully removed from the heating furnace

and rapidly cooled, using fans or the like, to either about 250°F or less, perhaps even fully

back down to room temperature. In any event, the preferred time/temperature exposures for the third aging step of this invention closely parallel those set forth for the first aging

step above.

[0095] In accordance with the invention, the invention alloy is preferably made

into a product, suitably an ingot derived product, suitable for hot rolling. For instance,

large ingots can be semi-continuously cast ofthe aforesaid composition and then can be

scalped or machined to remove surface imperfections as needed or required to provide a

good rolling surface. The ingot may then be preheated to homogenize and solutionize its

interior structure and a suitable preheat treatment is to heat to a relatively high

temperature for this type of composition, such as 900°F. In doing so, it is preferred to

heat to a first lesser temperature level such as heating above 800°F, for instance about

820°F or above, or 850°F or above, preferably 860°F or more, for instance around 870°F

or more, and hold the ingot at about that temperature or temperatures for a significant

time, for instance, 3 or 4 hours. Next the ingot is heated the rest ofthe way up to a

temperature of around 890°F or 900°F or possibly more for another hold time of a few

hours. Such stepped or staged heat ups for homogenizing have been known in the art for

many years. It is preferred that homogenizing be conducted at cumulative hold times in

the neighborhood of 4 to 20 hours or more, the homogenizing temperatures referring to

temperatures above about 880 to 890°F. That is, the cumulative hold time at

temperatures above about 890°F should be at least 4 hours and preferably more, for

instance 8 to 20 or 24 hours, or more. As is known, larger ingot size and other matters

can suggest longer homogenizing times. It is preferred that the combined total volume percent of insoluble and soluble constituents be kept low, for instance not over 1.5 vol.%,

preferably not over 1 vol.%. Use ofthe herein described relatively high preheat or

homogenization and solution heat treat temperatures aid in this respect, although high

temperatures warrant caution to avoid partial melting. Such cautions can include careful

heat-ups including slow or step-type heating, or both.

[0096] The ingot is then hot rolled and it is desirable to achieve an unrecrystallized

grain structure in the rolled plate product. Hence, the ingot for hot rolling can exit the

furnace at a temperature substantially above about 820°F, for instance around 840 to

850°F or possibly more, and the rolling operation is carried out at initial temperatures

above 775°F, or better yet, above 800°F, for instance around 810 or even 825°F. This

increases the likelihood of reducing recrystallization and it is also preferred in some

situations to conduct the rolling without a reheating operation by using the power ofthe

rolling mill and heat conservation during rolling to maintain the rolling temperature above

a desired minimum, such as 750°F or so. Typically, in practicing the invention, it is

preferred to have a maximum recrystallization of about 50% or less, preferably about

35% or less, and most preferably no more than about 25% recrystallization, it being

understood that the less recrystallization achieved, the better the fracture toughness

properties.

[0097] Hot rolling is continued, normally in a reversing hot rolling mill, until the

desired thickness ofthe plate is achieved. In accordance with the invention, plate product

intending to be machined into aircraft components such as integral spars can range from about 2 to 3 inches to about 9 or 10 inches thick or more. Typically, this plate ranges

from around 4 inches thick for relatively smaller aircraft, to thicker plate of about 6 or 8

inches to about 10 or 12 inches or more. In addition to the preferred embodiments, it is

believed this invention can be used to make the lower wing skins of small, commercial jet

airliners. Still other applications can include forgings and extrusions, especially thick

sectioned versions of same. In making extrusion, the invention alloy is extruded within

around 600° to 750°F, for instance, at around 700°F, and preferably includes a reduction

in cross-sectional area (extrusion ratio) of about 10:1 or more. Forging can also be used

herein.

[0098] The hot rolled plate or other wrought product is solution heat treated (SHT)

by heating within around 840 or 850°F to 880 or 900°F to take into solution substantial

portions, preferably all or substantially all, ofthe zinc, magnesium and copper soluble at

the SHT temperature, it being understood that with physical processes which are not

always perfect, probably every last vestige of these main alloying ingredients may not be

fully dissolved during the SHT (solutionizing). After heating to the elevated temperature

as just described, the product should be quenched to complete the solution heat treating

procedure. Such cooling is typically accomplished either by immersion in a suitably

sized tank of cold water or by water sprays, although air chilling might be usable as

supplementary or substitute cooling means for some cooling. After quenching, certain

products may need to be cold worked, such as by stretching or compression, so as to

relieve internal stresses or straighten the product, even possibly in some cases, to further strengthen the plate product. For instance, the plate may be stretched or compressed 1 or

1 /4 or possibly 2% or 3% or more, or otherwise cold worked a generally equivalent

amount. A solution heat freated (and quenched) product, with or without cold working, is

then considered to be in a precipitation-hardenable condition, or ready for artificial aging

according to preferred artificial aging methods as herein described or other artificial aging

techniques. As used herein, the term "solution heat treat", unless indicated otherwise,

shall be meant to include quenching.

[0099] After quenching, and cold working if desired, the product (which may be a

plate product) is artificially aged by heating to an appropriate temperature to improve

strength and other properties. In one preferred thermal aging treatment, the precipitation

hardenable plate alloy product is subjected to three main aging steps, phases or treatments

as described above, although clear lines of demarcation may not exist between each step

or phase. It is generally known that ramping up to and/or down from a given or target

treatment temperature, in itself, can produce precipitation (aging) effects which can, and

often need to be, taken into account by integrating such ramping conditions and their

precipitation hardening effects into the total aging treatment.

[0100] It is also possible to use aging integration in conjunction with the aging

practices of this invention. For instance, in a programmable air furnace, following

completion of a first stage heat treatment of 250°F for 24 hours, the temperature in that

same furnace can be gradually progressively raised to temperature levels around 310° or

so over a suitable length of time, even with no true hold time, after which the metal can

then be immediately transferred to another furnace already stabilized at 250°F and held for 6 to 24 hours. This more continuous, aging regime does not involve transitioning to

room temperature between first-to-second and second-to-third stage aging treatments.

Such aging integration was described in more detail in U.S. Patent 3,645,804, the entire

content of which is fully incorporated by reference herein. With ramping and its

corresponding integration, two, or on a less preferred basis, possibly three, phases for

artificially aging the plate product may be possible in a single, programmable furnace.

For purposes of convenience and ease of understanding, however, preferred embodiments

of this invention have been described in more detail as if each stage, step or phase was

distinct from the other two artificial aging practices imposed hereon. Generally speaking,

the first of these three steps or stages is believed to precipitation harden the alloy product

in question; the second (higher temperature) stage then exposes the invention alloy to one

or more elevated temperatures for increasing its resistance to corrosion, especially sfress

corrosion cracking (SCC) resistance under both normal, industrial and seacoast-

simulated atmospheric conditions. The third and final stage then further precipitation

hardens the invention alloy to a high strength level while also imparting further improved

corrosion properties thereto.

[0101] The low quench sensitivity ofthe invention alloy can offer yet another

potential application in a class of processes generally described as "press quenching" by

those skilled in the art. One can illustrate the "press quenching" process by considering

the standard manufacturing flow path of an age hardenable extrusion alloy such as one

that belongs to the 2XXX, 6XXX, 7XXX or 8XXX alloy series. The typical flow path

involves: Direct Chill (DC) ingot casting of billets, homogenization, cooling to ambient temperature, reheating to the extrusion temperature by furnaces or induction heaters,

extrusion ofthe heated billet to final shape, cooling the extruded part to ambient

temperature, solution heat treating the part, quenching, stretching and either naturally

aged at room temperature or artificially aged at elevated temperature to the final temper.

The "press quenching" process involves controlling the extrusion temperature and other

extrusion conditions such that upon exiting the extrusion die, the part is at or near the

desired solution heating temperature and the soluble constituents are effectively brought

to solid solution. It is then immediately and directly continuously quenched as the part

exits the extrusion press by either water, pressurized air or other media. The press

quenched part can then go through the usual stretching, followed by either natural or

artificial aging. Hence, as compared to the typical flow path, the costly separate solution

heat treating process is eliminated from this press quenched variation, thereby

significantly lowering overall manufacturing costs, and energy consumption as well.

[0102] For most alloys, especially those belonging to the relatively quench

sensitive 7XXX alloy series, the quench provided by the press quenching process is

generally not as effective as compared to that provided by the separate solution heat

freatment, such that significant degradation of certain material attributes such as sfrength,

fracture toughness, corrosion resistance and other properties can result from press

quenching. Since the invention alloy has very low quench sensitivity, it is expected that

the property degradation during press quenching is either eliminated or significantly

reduced to acceptable levels for many applications. [0103] For the mold plate embodiments of this invention where SCC resistance is

not as critical, known single or two-stage artificial aging treatments may also be

practiced on these compositions instead ofthe preferred three step aging method

described herein.

[0104] When referring to a minimum (for instance, strength or toughness property

value), such can refer to a level at which specifications for purchasing or designating

materials can be written or a level at which a material can be guaranteed or a level that an

airframe builder (subject to safety factor) can rely on in design. In some cases, it can

have a statistical basis wherein 99% ofthe product conforms or is expected to conform

with 95%) confidence using standard statistical methods. Because of an insufficient

amount of data, it is not statistically accurate to refer to certain minimum or maximum

values ofthe invention as true "guaranteed" values. In those instances, calculations have

been made from currently available data for extrapolating values (e.g. maximums and

minimums) therefrom. See, for example, the Currently Extrapolated Minimum S/N

values plotted for plate (solid line A-A in Figure 12) and forgings (solid line B-B in

Figure 13), and the Currently Extrapolated FCG Maximum (solid line C-C in Figure 14).

[0105] Fracture toughness is an important property to airframe designers,

particularly if good toughness can be combined with good sfrength. By way of

comparison, the tensile strength, or ability to sustain load without fracturing, of a

structural component under a tensile load can be defined as the load divided by the area of

the smallest section ofthe component perpendicular to the tensile load (net section stress).

For a simple, straight-sided structure, the strength ofthe section is readily related to the breaking or tensile sfrength of a smooth tensile coupon. This is how tension testing is

done. However, for a structure containing a crack or crack-like defect, the sfrength of a

structural component depends on the length ofthe crack, the geometry ofthe structural

component, and a property ofthe material known as the fracture toughness. Fracture

toughness can be thought of as the resistance of a material to the harmful or even

catastrophic propagation of a crack under a load.

[0106] Fracture toughness can be measured in several ways. One way is to load in

tension a test coupon containing a crack. The load required to fracture the test coupon

divided by its net section area (the cross-sectional area less the area containing the crack)

is known as the residual strength with units of thousands of pounds force per unit area

(ksi). When the strength ofthe material as well as the specimen geometry are constant,

the residual strength is a measure ofthe fracture toughness ofthe material. Because it is

so dependent on strength and specimen geometry, residual strength is usually used as a

measure of fracture toughness when other methods are not as practical as desired because

of some constraint like size or shape ofthe available material.

[0107] When the geometry of a structural component is such that it does not

deform plastically through the thickness when a tension load is applied (plane-strain

deformation), fracture toughness is often measured as plane-strain fracture toughness, K_Ic.

This normally applies to relatively thick products or sections, for instance 0.6 or

preferably 0.8 or 1 inch or more. The ASTM has established a standard test using a

fatigue pre-cracked compact tension specimen to measure K_lc which has the units ksWin.

This test is usually used to measure fracture toughness when the material is thick because it is believed to be independent of specimen geometry as long as appropriate standards for

width, crack length and thickness are met. The symbol K, as used in K_lc, is referred to as

the sfress intensity factor.

[0108] Structural components which deform by plane-strain are relatively thick as

indicated above. Thinner structural components (less than 0.8 to 1 inch thick) usually

deform under plane stress or more usually under a mixed mode condition. Measuring

fracture toughness under this condition can introduce variables because the number which

results from the test depends to some extent on the geometry ofthe test coupon. One test

method is to apply a continuously increasing load to a rectangular test coupon containing

a crack. A plot of stress intensity versus crack extension known as an R-curve (crack

resistance curve) can be obtained this way. The load at a particular amount of crack

extension based on a 25% secant offset in the load vs. crack extension curve and the

effective crack length at that load are used to calculate a measure of fracture toughness

known as K_m5. At a 20% secant, it is known as K^o- It also has the units of ksWin.

Well known ASTM E561 concerns R-curve determination, and such is generally

recognized in the art.

[0109] When the geometry ofthe alloy product or structural component is such that

it permits deformation plastically through its thickness when a tension load is applied,

fracture toughness is often measured as plane-stress fracture toughness which can be

determined from a center cracked tension test. The fracture toughness measure uses the

maximum load generated on a relatively thin, wide pre-cracked specimen. When the

crack length at the maximum load is used to calculate the stress-intensity factor at that load, the stress-intensity factor is referred to as plane-sfress fracture toughness K_c. When

the stress-intensity factor is calculated using the crack length before the load is applied,

however, the result ofthe calculation is known as the apparent fracture toughness, K_app, of

the material. Because the crack length in the calculation of K- is usually longer, values

for K-- are usually higher than K_app for a given material. Both of these measures of

fracture toughness are expressed in the units ksWin. For tough materials, the numerical

values generated by such tests generally increase as the width ofthe specimen increases

or its thickness decreases as is recognized in the art. Unless indicated otherwise herein,

plane sfress (K_c) values referred to herein refer to 16-inch wide test panels. Those skilled

in the art recognize that test results can vary depending on the test panel width, and it is

intended to encompass all such tests in referring to toughness. Hence, toughness

substantially equivalent to or substantially corresponding to a minimum value for K--, or

K_app in characterizing the invention products, while largely referring to a test with a

16-inch panel, is intended to embrace variations in K_c or K_app encountered in using

different width panels as those skilled in the art will appreciate.

[0110] The temperature at which the toughness is measured can be significant. In

high altitude flights, the temperature encountered is quite low, for instance, minus 65 °F,

and for newer commercial jet aircraft projects, toughness at minus 65°F is a significant

factor, it being desired that the lower wing material exhibit a toughness K_lc level of

around 45 ksWin at minus 65°F or, in terms of K^o, a level of 95 ksWin, preferably 100

ksWin or more. Because of such higher toughness values, lower wings made from these alloys may replace today's 2000 (or 2XXX Series) alloy counterparts with their

corresponding property (i.e. strength/toughness) trade-offs. Through the practice of this

invention, it may also be possible to make upper wing skins from same, alone or in

combination with integrally formed components, like stiffeners, ribs and stringers.

[0111] The toughness ofthe improved products according to the invention is very

high and in some cases may allow the aircraft designer's focus for a material's durability

and damage tolerance to emphasize fatigue resistance as well as fracture toughness

measurement. Resistance to cracking by fatigue is a very desirable property. The fatigue

cracking referred to occurs as a result of repeated loading and unloading cycles, or

cycling between a high and a low load such as when a wing moves up and down. This

cycling in load can occur during flight due to gusts or other sudden changes in air

pressure, or on the ground while the aircraft is taxing. Fatigue failures account for a large

percentage of failures in aircraft components. These failures are insidious because they

can occur under normal operating conditions without excessive overloads, and without

warning. Crack evolution is accelerated because material inhomogeneities act as sites for

initiation or facilitate linking of smaller cracks. Therefore, process or compositional

changes which improve metal quality by reducing the severity or number of harmful

inhomogeneities improve fatigue durability.

[0112] Stress-life cycle (S-N or S/N) fatigue tests characterize a material resistance

to fatigue initiation and small crack growth which comprises a major portion of total

fatigue life. Hence, improvements in S-N fatigue properties may enable a component to

operate at higher stresses over its design life or operate at the same sfress with increased lifetime. The former can translate into significant weight savings by downsizing, or

manufacturing cost saving by component or structural simplification, while the latter can

translate into fewer inspections and lower support costs. The loads during fatigue testing

are below the static ultimate or tensile sfrength ofthe material measured in a tensile test

and they are typically below the yield strength ofthe material. The fatigue initiation

fatigue test is an important indicator for a buried or hidden structural member such as a

wing spar which is not readily accessible for visual or other examination to look for

cracks or crack starts.

[0113] If a crack or crack-like defect exists in a structure, repeated cyclic or fatigue

loading can cause the crack to grow. This is referred to as fatigue crack propagation.

Propagation of a crack by fatigue may lead to a crack large enough to propagate

catasfrophically when the combination of crack size and loads are sufficient to exceed the

material's fracture toughness. Thus, performance in the resistance of a material to crack

propagation by fatigue offers substantial benefits to aerostructure longevity. The slower a

crack propagates, the better. A rapidly propagating crack in an airplane structural

member can lead to catastrophic failure without adequate time for detection, whereas a

slowly propagating crack allows time for detection and corrective action or repair.

Hence, a low fatigue crack growth rate is a desirable property.

[0114] The rate at which a crack in a material propagates during cyclic loading is

influenced by the length ofthe crack. Another important factor is the difference between

the maximum and the minimum loads between which the structure is cycled. One

measurement including the effects of crack length and the difference between maximum and minimum loads is called the cyclic stress intensity factor range or ΔK, having units of

ks in, similar to the sfress intensity factor used to measure fracture toughness. The sfress

intensity factor range (ΔK) is the difference between the stress intensity factors at the

maximum and minimum loads. Another measure affecting fatigue crack propagation is

the ratio between the minimum and the maximum loads during cycling, and this is called

the stress ratio and is denoted by R, a ratio of 0.1 meaning that the maximum load is 10

times the minimum load. The sfress, or load, ratio may be positive or negative or zero.

Fatigue crack growth rate testing is typically done in accordance with ASTM E647-88

(and others) well known in the art. As used herein, Kt refers to a theoretical stress

concentration factor as described in ASTM El 823.

[0115] The fatigue crack propagation rate can be measured for a material using a

test coupon containing a crack. One such test specimen or coupon is about 12 inches long

by 4 inches wide having a notch in its center extending in a cross-wise direction (across

the width; normal to the length). The notch is about 0.032 inch wide and about 0.2 inch

long including a 60° bevel at each end ofthe slot. The test coupon is subjected to cyclic

loading and the crack grows at the end(s) ofthe notch. After the crack reaches a

predetermined length, the length ofthe crack is measured periodically. The crack growth

rate can be calculated for a given increment of crack extension by dividing the change in

crack length (called Δa) by the number of loading cycles (ΔN) which resulted in that

amount of crack growth. The crack propagation rate is represented by Δa/ΔN or 'da/dN'

and has units of inches/cycle. The fatigue crack propagation rates of a material can be determined from a center cracked tension panel. In a comparison using R=0.1 tested at a

relative humidity over 90% with ΔK ranging from around 4 to 20 or 30, the invention

material exhibited relatively good resistance to fatigue crack growth. However, the

superior performance in S-N fatigue makes the invention material much better suited for a

buried or hidden member such as a wing spar.

[0116] The invention products exhibit very good corrosion resistance in addition to

the very good sfrength and toughness and damage tolerance performance. The exfoliation

corrosion resistance for products in accordance with the invention can be EB or better

(meaning "EA" or pitting only) in the EXCO test for test specimens taken at either

mid-thickness (T/2) or one-tenth ofthe thickness from the surface (T/10) ("T" being

thickness) or both. EXCO testing is known in the art and is described in well known

ASTM Standard No. G34. An EXCO rating of "EB" is considered good corrosion

resistance in that it is considered acceptable for some commercial aircraft; "EA" is still

better.

[0117] Stress corrosion cracking resistance across the short fransverse direction is

often considered an important property especially in relatively thick members. The stress

corrosion cracking resistance for products in accordance with the invention in the short

transverse direction can be equivalent to that needed to pass a 1/8-inch round bar alternate

immersion test for 20, or alternately 30, days at 25 or 30 ksi or more, using test

procedures in accordance with ASTM G47 (including ASTM G44 and G38 for C-ring

specimens and G49 for 1/8-inch bars), said ASTM G47, G44, G49 and G38, all well

known in the art. [0118] As a general indicator of exfoliation corrosion and stress corrosion

resistance, the plate typically can have an electrical conductivity of at least about 36, or

preferably 38 to 40%> or more ofthe International Annealed Copper Standard (%IACS).

Thus, the good exfoliation corrosion resistance ofthe invention is evidenced by an EXCO

rating of "EB" or better, but in some cases other measures of corrosion resistance may be

specified or required by airframe builders, such as sfress corrosion cracking resistance or

electrical conductivity. Satisfying any one or more of these specifications is considered

good corrosion resistance.

[0119] The invention has been described with some emphasis on wrought plate

which is preferred, but it is believed that other product forms may be able to enjoy the

benefits ofthe invention, including extrusions and forgings. To this point, the emphasis

has been on stiffener-type, fuselage or wing skin stringers which can be J-shaped, Z- or S-

shaped, or even in the shape of a hat-shaped channel. The purpose of such stiffeners is to

reinforce the plane's wing skin or fuselage, or any other shape that can be attached to

same, while not adding a lot of weight thereto. While in some cases it is preferred for

manufacturing economies to separately fasten stringers, such can be machined from a

much thicker plate by the removal ofthe metal between the stiffener geometries, leaving

only the stiffener shapes integral with the main wing skin thickness, thus eliminating all

the rivets. Also the invention has been described in terms of thick plate for machining

wing spar members as explained above, the spar member generally corresponding in

length to the wing skin material. In addition, significant improvements in the properties

of this invention render its use as thickly cast mold plate highly practical. [0120] Because of its reduced quench sensitivity, it is believed that when an alloy

product according to the invention is welded to a second product, it will exhibit in its heat

affected, welding zone an improved retention of its sfrength, fatigue, fracture toughness

and/or corrosion resistance properties. This applies regardless of whether such alloy

products are welded by solid state welding techniques, including friction stir welding, or

by known or subsequently developed fusion techniques including, but not limited to,

electron beam welding and laser welding. Through the practice of this invention, both

welded parts may be made from the same alloy composition.

[0121] For some parts/products made according to the invention, it is likely that

such parts/products may be age formed. Age forming promises a lower manufacturing

cost while allowing more complex wing shapes to be formed, typically on thinner gauge

components. During age forming, the part is mechanically constrained in a die at an

elevated temperature usually about 250°F or higher for several to tens of hours, and

desired contours are accomplished through stress relaxation. Especially during a higher

temperature artificial aging treatment, such as a treatment above about 320°F, the metal

can be formed or deformed into a desired shape. In general, the deformations envisioned

are relatively simple such as including a very mild curvature across the width of a plate

member together with a mild curvature along the length of said plate member. It can be

desirable to achieve the formation of these mild curvature conditions during the artificial

aging treatment, especially during the higher temperature, second stage artificial aging

temperature. In general, the plate material is heated above around 300°F, for instance around 320 or 330°F, and typically can be placed upon a convex form and loaded by

clamping or load application at opposite edges ofthe plate. The plate more or less

assumes the contour ofthe form over a relatively brief period of time but upon cooling

springs back a little when the force or load is removed. The expected springback is

compensated for in designing the curvature or contour ofthe form which is slightly

exaggerated with respect to the desired forming ofthe plate to compensate for springback.

Most preferably, the third artificial aging stage at a low temperature such as around 250°F

follows age forming. Either before or after its age forming freatment, the plate member

can be machined, for instance, such as by tapering the plate such that the portion intended

to be closer to the fuselage is thicker and the portion closest to the wing tip is thinner.

Additional machining or other shaping operations, if desired, can also be performed either

before or after age forming. High capacity aircrafts may require a relatively thicker plate

and a higher level of forming than previously used on a large scale for thinner plate

sections.

[0122] Narious invention alloy product forms, i.e. both thick plate (Figure 12) and

forgings (Figure 13), were made, aged and suitably sized samples taken for performing

fatigue life (S/Ν) tests thereon consistent with known open hole fatigue life testing

procedures. Precise compositions for these product forms were as follows:

For these open hole fatigue life evaluations, in the L-T orientation, specific test

parameters for both plate and forged product forms included: a K_t value of 2.3, Frequency

of 30 Hz, R value = 0.1 and Relative Humidity (RH) greater than 90%. The plate test

results were then graphed in accompanying Figure 12; and the forging results in

accompanying Figure 13. Both plate and forging forms were tested over several product

thicknesses (4, 6 and 8 inches).

[0123] Referring now to Figure 12, a mean S/N performance (solid) line drawn

through both sets of 6 inch thick plate data (alloys D and E above). A 95% confidence

band was then drawn (per the upper and lower dotted lines) around the aforementioned 6

inch "mean" performance line. From that data, a set of points was mapped representing

currently extrapolated minimum open hole fatigue life (S/N) values. Those precise

mapped points were:

Solid line (A-A) was then drawn on Figure 12 to connect the aforementioned currently

extrapolated minimum S/N values of Table 12. Against those preferred minimum S/N values, one jetliner manufacturer's specified S/N value lines for 7040/7050-T7451 plate

(3 to 8.7 inch thick) and 7010/7050-T7451 plate (2 to 8 inch thick) were overlaid. Line

A-A shows this invention's likely relative improvement in fatigue life S/N performance

over known, commercial aerospace 7XXX alloys even though the comparative data for

the latter known alloys was taken in a different (T-L) orientation.

[0124] From the open hole fatigue life (S/N) data for various sized (i.e. 4 inch, 6

inch and 8 inch) forgings, a dotted line was drawn for mathematically representing the

mean values of 6 inch thick comp E and 8 inch thick comp D forgings. Note, several

samples tested did not fracture during these tests; they are grouped together in a circle to

the right of Figure 13. Thereafter, a set of points was mapped representing currently

extrapolated minimum open hole fatigue life (S/N) values. Those precise mapped points

were:

Solid line (B-B) was then drawn on Figure 13 to connect the aforementioned currently

extrapolated minimum S/N forging values of above Table 13.

[0125] In Figure 14, the Fatigue Crack Growth (FCG) rate curves for plate (4 and 6

inch thicknesses, both L-T and T-L orientations) and forged product (6 inch, L-T only) made according to the invention are plotted. The actual compositions tested are listed in

above Table 11. These tests, conducted per the FCG procedures described above,

employed particulars of: Frequency ^■= 25 Hz, an R value = 0.1 and relative humidity (RH)

greater than 95%. From those curves, for the various product forms and thicknesses, one

set of data points was mapped representing currently extrapolated maximum FCG values

for the invention. Those precise points were:

A currently extrapolated maximum FCG value, solid curve line (C-C) for thick plate and

forgings per the invention was drawn, against which one jetliner manufacturer's specified

FCG values for 7040/7050-T7451 (3 to 8.7 in thick) plate was overlaid, said values being

taken in both the L-T and T-L orientation.

[0126] Plate product forms ofthe invention have also been subjected to hole crack

initiation tests, involving the drilling of a preset hole (less than 1 in. diameter) into a test

specimen, inserting into that drilled hole a split sleeve, then pulling a variably oversized

mandrel through said sleeve and pre-drilled hole. Under such testing, the 6 and 8 inch

thick plate product of this invention did not have any cracks initiate from the drilled holes

thereby showing very good performance. [0127] Having described the presently preferred embodiments, it is to be

understood that the invention may be otherwise embodied within the scope ofthe

appended claims.

Claims (1)

1. What is claimed is:

   1. An aluminum alloy product that possesses the ability to achieve:

   (a) in products having a thick section when solution heat freated, quenched and artificially

   aged, and in parts made from said products, an improved combination of at least two

   properties selected from the group consisting of: strength, fracture toughness and

   corrosion resistance; or (b) in thin products that are slowly quenched, and in parts made

   therefrom, less degradation in strength resulting from said slow quench, said alloy

   consisting essentially of:

   about 6 to 10 wt.% Zn; about 1.2 to 1.9 wt.% Mg; about 1.2 to 2.2 wt.% Cu;

   one or more elements present selected from the group consisting of: up to about

   0.4 wt.% Zr, up to about 0.4 wt.% Sc and up to about 0.3 wt.% Hf; said alloy

   optionally containing up to: about 0.06 wt.% Ti, about 0.03 wt.% Ca, about 0.03

   wt.%) Sr, about 0.002 wt.% Be and about 0.3 wt.% Mn, the balance being Al,

   incidental elements and impurities.

   2. The alloy product of claim 1 wherein said alloy contains about

   6.4 to 9.5 wt.% Zn; about 1.3 to 1.7 wt.% Mg; about 1.3 to 1.9 wt.% Cu, with wt % Mg <

   (wt.% Cu + 0.3) and about 0.05 to 0.2 wt.% Zr.

   3. The alloy product of claim 2 which is at least about 2 inches at its

   thickest cross sectional point.

   4. The alloy product of claim 3 which is about 3 to 10 inches at said

   thickest point.

   5. The alloy product of claim 4 which is about 4 to 6 inches at said

   thickest point.

   6. The alloy product of claim 2 wherein wt % Mg < (wt.% Cu + 0.2).

   7. The alloy product of claim 6 wherein wt % Mg < (wt.% Cu + 0.1).

   8. The alloy product of claim 2 wherein wt % Mg wt.% Cu.

   9. The alloy product of claim 2 which further exhibits improved sfress

   corrosion cracking resistance.

   10. The alloy product of claim 2 which is a thick plate, extrusion or

   forged product.

   11. The alloy product of claim 2 which is a thin plate about 2 inches

   thick or less.

   12. The alloy product of claim 11 which further exhibits improved

   exfoliation corrosion resistance.

   13. The alloy product of claim 11 which is age formed to the shape of an

   aerospace structural component.

   14. The alloy product of claim 2 wherein said alloy contains, as

   impurities, about 0.15 wt.% or less Fe and about 0.12 wt.% or less Si.

   15. The alloy product of claim 14 wherein said alloy contains an

   effective Mg content of about 1.3 to 1.65 wt.%, for a total measurable Mg content of

   about 1.47 to 1.82 wt%.

   16. The alloy product of claim 14 wherein said alloy contains an

   effective Cu content of about 1.3 to 1.9 wt.%, for a total measurable Cu content of about

   1.6 to 2.2 t%.

   17. The alloy product of claim 14 wherein said alloy contains about 0.08

   wt.%) or less Fe and about 0.06 wt.% or less Si.

   18. The alloy product of claim 17 wherein said alloy contains about 0.04

   wt.% or less Fe and about 0.03 wt.% or less Si.

   19. The alloy product of claim 2 wherein said alloy contains about 6.9 or

   higher wt% Zn.

   20. The alloy product of claim 2 wherein said alloy contains about 6.9 to

   8.5 wt.% Zn; about 1.3 to 1.68 wt. % Mg; about 1.3 to 1.9 wt.% Cu and about 0.05 to 0.2

   wt.% Zr.

   21. The alloy product of claim 2 wherein said alloy consists essentially

   of: about 6.9 to 8 wt.% Zn; about 1.3 to 1.65 wt. % Mg; about 1.4 to 1.9 wt.% Cu and

   about 0.05 to 0.2 wt.% Zr; with wt.% Mg < wt.% Cu.

   22. The alloy product of claim 2 wherein (wt.%> Mg + wt.%> Cu) 3.5.

   23. The alloy product of claim 22 wherein (wt.% Mg + wt.% Cu) 3.3.

   24. The alloy product of claim 2 which is less than about 50%

   recrystallized.

   25. The alloy product of claim 24 which is about 35% or less

   recrystallized.

   26. The alloy product of claim 25 which is about 25% or less

   recrystallized.

   27. The alloy product of claim 2 which is welded to a second alloy

   product and exhibits in its heat affected, welding zone an improved retention of one or

   more properties selected from the group consisting of: strength, fatigue, fracture

   toughness and corrosion resistance.

   28. The alloy product of claim 27 which is welded by a solid state

   method.

   29. The alloy product of claim 28 which is welded by friction stir

   welding.

   30. The alloy product of claim 27 which is welded by a fusion welding

   method.

   31. The alloy product of claim 30 which is welded by an electron beam

   method.

   32. The alloy product of claim 30 which is welded by a laser method.

   33. The alloy product of claim 27 wherein said second alloy product is

   made ofthe same alloy to which it is welded.

   34. The alloy product of claim 2 which exhibits an improved resistance

   to hole crack initiation.

   35. A wrought aluminum alloy product, said alloy consisting essentially

   of: about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.3 to 1.9 wt.% Cu, with

   wt % Mg < (wt.% Cu + 0.3); at least one element present selected from the group

   consisting of: (up to about 0.3 wt.% Zr; up to about 0.4 wt.%> Sc and up to about 0.3 wt.%)

   Hf); optionally, up to about: 0.06 wt.% Ti and 0.008 wt.% Ca, the balance Al, incidental

   elements and impurities, said alloy product characterized by low quench sensitivity and:

   (a) in products having a thick section when solution heat treated, quenched, and

   artificially aged, and in parts made from said thick products, an improved combination of

   at least two properties selected from the group consisting of: strength, fracture toughness

   and corrosion resistance; or (b) in thin products that are slowly quenched, and in parts

   made from said thin products, less degradation in strength.

   36. The alloy product of claim 35 which is between about 3 to 12 inches

   at its thickest point

   37. The alloy product of claim 36 which is between about 4 to 6 inches

   at said thickest point.

   38. The alloy product of claim 35 wherein wt % Mg does not exceed

   wt.% Cu in said composition.

   39. The alloy product of claim 35 which is a plate, extrusion or forging

   that has been solution heat treated and quenched.

   40. The alloy product of claim 35 wherein said alloy contains, as

   impurities, less than about 0.25 wt.% Fe and wt.% Si each.

   41. The alloy product of claim 35 wherein said alloy contains about 6.9

   to 8 wt.% Zn; about 1.3 to 1.65 wt. % Mg; about 1.3 to 1.9 wt.% Cu; and about 0.05 to

   0.2 wt.%) Zr, with (wt.% Mg + wt.% Cu) 3.5.

   42. The alloy product of claim 41 wherein said alloy contains about 7 to

   8 wt.% Zn; about 1.4 to 1.65 wt.% Mg; about 1.4 to 1.8 wt.% Cu; and about 0.05 to 0.2

   wt.% Zr, with (wt.% Mg + wt.% Cu) 3.3.

   43. A thick aluminum alloy product that when solution heat freated,

   quenched in a thick section, and artificially aged possesses an improved combination of

   strength and toughness along with good corrosion resistance properties, said alloy

   consisting essentially of:

   about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.3 to 2.1 wt.%

   Cu, with wt % Mg < (wt.% Cu + 0.3); about 0.05 to 0.2 wt.% Zr; the balance being

   Al, incidental elements and impurities.

   44. The alloy product of claim 43 wherein wt % Mg wt.% Cu.

   45. The alloy product of claim 43 wherein said alloy contains about 0.15

   wt.% or less Fe and about 0.12 wt.% or less Si.

   46. The alloy product of claim 43 wherein said alloy contains about 7 to

   8 wt.% Zn, about 1.3 to 1.65 wt.% Mg, about 1.4 to 1.8 wt.% Cu and about 0.05 to 0.2

   wt.% Zr, with wt % Mg < (wt.% Cu + 0.1).

   47. The alloy product of claim 43 which has, at a point 2 inches or more

   thick in cross section, a quarter-plane (T/4) tensile yield strength TYS in the longitudinal

   (L) direction and a quarter-plane (T/4) plane-strain fracture toughness (K_lc) in the L-T

   direction at or above (to the right of) line M-M in Figure 7.

   48. The alloy product of claim 43 which is a plate product having a

   minimum open-hole fatigue life (S/N) at one or more ofthe applied maximum stress

   levels set forth in Table 12 equal to or greater than the corresponding cycles to failure

   value in said Table 12.

   49. The alloy product of claim 43 which is a plate product having a

   minimum open hole fatigue life (S/N) at or above (to the right of) line A-A in Figure 12.

   50. The alloy product of claim 43 which is a forging having a minimum

   open hole fatigue life (S/N) at or above (to the right of) line B-B in Figure 13.

   51. The alloy product of claim 43 which has a maximum fatigue crack

   growth (FCG) rate in the L-T test orientation at or below at least one ofthe maximum

   da/dN values set forth in Table 14 for the corresponding K (stress intensity factor) values

   at or greater than 15 ksiin in said Table 14.

   52. The alloy product of claim 43 which has a maximum fatigue crack

   growth (FCG) rate in the L-T test orientation for a K of 15 ksiin or more at or below (to

   the right of) line C-C in Figure 14.

   53. The alloy product of claim 43 which is capable of passing at least 30

   days of alternate immersion, stress corrosion cracking (SCC) testing with a 3.5 % Na

   solution at a short fransverse (ST) sfress level of about 30 ksi or more.

   54. The alloy product of claim 43 which has a minimum life without

   failure against stress corrosion cracking after at least about 100 days of seacoast exposure

   at a short fransverse (ST) stress level of about 30 ksi or more.

   55. The alloy product of claim 54 which has a minimum life without

   failure against stress corrosion cracking after at least about 180 days of said seacoast

   exposure conditions.

   56. The alloy product of claim 43 which has a minimum life without

   failure against sfress corrosion cracking after at least about 180 days of industrial

   exposure at a short fransverse (ST) sfress level of about 30 ksi or more.

   57. The alloy product of claim 43 which has both thick and thin sections

   after one or more machining operations are performed thereon, said thin sections

   exhibiting EXCO corrosion resistance rating of "EB" or better.

   58. The alloy product of claim 43 which exhibits an improved resistance

   to hole crack initiation.

   59. The alloy product of claim 43 which has been artificially aged by a

   method comprising:

   . (i) a first aging stage within about 200 to 275°F;

   (ii) a second aging stage within about 300 to 335°F; and

   (iii) a third aging stage within about 200 to 275°F.

   60. The alloy product of claim 59 wherein first aging stage (i) proceeds

   within about 230 to 260°F.

   61. The alloy product of claim 59 wherein first aging stage (i) proceeds

   for about 2 to 18 hours.

   62. The alloy product of claim 59 wherein second aging stage (ii)

   proceeds within about 300 to 325°F.

   63. The alloy product of claim 59 wherein second aging stage (ii)

   proceeds for about 4 to 18 hours within about 300 to 325°F.

   64. The alloy product of claim 63 wherein second aging stage (ii)

   proceeds for about 6 to 15 hours within about 300 to 315°F.

   65. The alloy product of claim 63 wherein second aging stage (ii)

   proceeds for about 7 to 13 hours within about 310 to 325°F.

   66. The alloy product of claim 59 wherein third aging stage (iii)

   proceeds within about 230 to 260°F.

   67. The alloy product of claim 66 wherein third aging stage (iii)

   proceeds for at least about 6 hours within about 230 to 260°F.

   68. The alloy product of claim 67 wherein third aging stage (iii)

   proceeds for about 18 hours or more within about 240 to 255°F.

   69. The alloy product of claim 59 wherein one or more of said first,

   second and third aging stages includes an integration of multiple temperature aging

   effects.

   70. The alloy product of claim 43 which is a stepped extrusion.

   71. The alloy product of claim 43 which is an extrusion that has been

   press quenched.

   72. The alloy product of claim 43 which is a plate product that can be

   age formed into an aerospace structural component.

   73. The alloy product of claim 43 which has been artificially aged by a

   method comprising:

   (i) a first aging stage within about 200 to 275°F; and

   (ii) a second aging stage within about 300 to 335°F.

   74. An aluminum alloy structural component for a commercial aircraft,

   said structural component made from a thick plate, extrusion or forged product that has

   been solution heat treated, quenched and artificially aged, said structural component

   possessing an improved combination of sfrength, toughness and stress corrosion cracking

   resistance properties, said alloy consisting essentially of:

   about 6.9 to 9.5 wt % Zn; about 1.3 to 1.68 wt.% Mg; about 1.2 to 2.2 wt.%

   Cu, with wt % Mg < (wt.% Cu + 0.3); and about 0.05 to 0.2 wt.% Zr, the balance

   Al, incidental elements and impurities.

   75. The structural component of claim 74 wherein wt % Mg wt.% Cu.

   76. The structural component of claim 74 wherein said plate, extrusion

   or forged product is between about 3 to 12 inches at its thickest cross sectional point.

   77. The structural component of claim 76 wherein said plate, extrusion

   or forged product is between about 4 to 6 inches at said thickest point.

   78. The structural component of claim 74 which exhibits reduced quench

   sensitivity compared to its 7050 aluminum alloy counterpart.

   79. The structural component of claim 74 wherein said alloy contains

   less than about 0.15 wt.% Fe and less than about 0.12 wt.% Si.

   80. The structural component of claim 74 wherein said alloy contains

   about 7 to 8 wt.% Zn, about 1.3 to 1.68 wt.% Mg, about 1.4 to 1.8 wt.% Cu and about

   0.05 to 0.2 wt.% Zr, with (wt.% Mg + wt.% Cu) 3.3.

   81. The structural component of claim 74 which is selected from the

   group consisting of a spar, rib, web, stringer, wing panel or skin, fuselage frame, floor

   beam, bulkhead, landing gear beam or combinations thereof.

   82. The structural component of claim 74 which is integrally formed.

   83. The structural component of claim 74 which has, at a point 2 inches

   or more thick in cross section, a quarter-plane (T/4) tensile yield sfrength TYS in the longitudinal (L) direction and a quarter-plane (T/4) plane-strain fracture toughness (K_lc)

   in the L-T direction at or above (to the right of) line M-M in Figure 7.

   84. The structural component of claim 74 which is a plate product

   having a minimum open hole fatigue life (S/N) at or above (to the right of) line A-A in

   Figure 12.

   85. The structural component of claim 74 which is a forging having a

   minimum open hole fatigue life (S/N) at or above (to the right of) line B-B in Figure 13.

   86. The structural component of claim 74 which has a maximum fatigue

   crack growth (FCG) rate in the L-T test orientation for a K (sfress intensity factor) of 15

   ksiin or more at or below (to the right of) line C-C in Figure 14.

   87. The structural component of claim 74 which is capable of passing at

   least 30 days of alternate immersion, stress corrosion cracking (SCC) testing with a 3.5 %

   Na solution at a short transverse (ST) stress level of about 30 ksi or more.

   88. The structural component of claim 74 which has a minimum life

   without failure against stress corrosion cracking after at least about 100 days of seacoast

   exposure at a short transverse (ST) stress level of about 30 ksi or more.

   89. The structural component of claim 74 which has a minimum life

   without failure against sfress corrosion cracking after at least about 180 days of industrial

   exposure at a short transverse (ST) stress level of about 30 ksi or more.

   90. The structural component of claim 74 which has both thick and thin

   sections, said thin sections exhibiting an EXCO corrosion resistance rating of "EB" or

   better.

   91. The structural component of claim 74 which exhibits an improved

   resistance to hole crack initiation.

   92. The structural component of claim 74 wherein said aircraft is a

   civilian or military jet aircraft.

   93. The structural component of claim 74 wherein said aircraft is a turbo

   prop plane.

   94. The structural component of claim 74 wherein said plate, extrusion

   or forged product is sfretched and/or compressed prior to being artificially aged.

   95. The structural component of claim 74 wherein said plate, extrusion

   or forged product is artificially aged by a method comprising:

   (i) a first aging stage within about 200 to 275°F;

   (ii) a second aging stage within about 300 to 335°F; and

   (iii) a third aging stage within about 200 to 275°F.

   96. The structural component of claim 95 wherein first aging stage (i)

   proceeds within about 230 to 260°F.

   97. The structural component of claim 96 wherein first aging stage (i)

   proceeds for 6 hours or more within about 235 to 255°F.

   98. The structural component of claim 95 wherein first aging stage (i)

   proceeds for about 2 to 12 hours.

   99. The structural component of claim 95 wherein second aging stage

   (ii) proceeds for about 4 to 18 hours within about 300 to 325 °F.

   100. The structural component of claim 99 wherein second aging stage

   (ii) proceeds for about 6 to 15 hours within about 300 to 315°F.

   101. The structural component of claim 99 wherein second aging stage

   (ii) proceeds for about 7 to 13 hours within about 310 to 325°F.

   102. The structural component of claim 95 wherein third aging stage (iii)

   proceeds for at least 6 hours within about 230 to 260°F.

   103. The structural component of claim 102 wherein third aging stage (iii)

   proceeds for 18 hours or more within about 240 to 255°F.

   104. A commercial aircraft structural component selected from the group

   consisting of: a spar, rib, web, stringer, wing panel or skin, fuselage frame, floor beam,

   bulkhead, landing gear beam or combinations thereof, said component having been

   machined from a thick plate, extrusion or forging and having improved strength, fracture

   toughness and corrosion resistance properties, said alloy consisting essentially of:

   about: 6.9 to 8.2 wt.% Zn; 1.3 to 1.68 wt.% Mg; 1.4 to 1.9 wt.% Cu, with

   wt % Mg < (wt.% Cu + 0.3); and about 0.05 to 0.2 wt.% Zr, the balance Al with

   incidental elements and impurities.

   105. The structural component of claim 104 wherein said alloy contains

   about 0.15 wt.% or less Fe and about 0.12 wt.% or less Si.

   106. The structural component of claim 104 which is welded to a second

   structural component and exhibits an improved retention of one or more properties

   selected from the group consisting of: strength, fatigue, fracture toughness and corrosion

   resistance in its heat affected, welding zone.

   107. An aircraft wingbox component made from an aluminum alloy plate,

   extrusion or forged product at least about 2 inches thick, said alloy consisting essentially

   of:

   about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.65 wt.% Mg; about 1.4 to 2 wt.%

   Cu, with (wt.% Mg + wt.% Cu) < 3.5; and about 0.05 to 0.25 wt.% Zr, the balance

   Al, incidental elements and impurities.

   108. The wingbox component of claim 107 wherein said alloy contains

   less than about 0.15 wt.%> Fe and less than about 0.12 wt.% Si.

   109. The wingbox component of claim 107 wherein said alloy contains

   less than about 8 wt.% Zn and less than about 1.9 wt.% Cu.

   110. The wingbox component of claim 107 which is an integral spar.

   111. The wingbox component of claim 110 which has been age formed.

   112. The wingbox component of claim 107 which is a rib, web or stringer.

   113. The wingbox component of claim 107 which is a wing panel or skin.

   114. The wingbox component of claim 113 which has been age formed.

   115. The wingbox component of claim 107 which is made from a stepped

   extrusion.

   116. The wingbox component of claim 107 which is a press quenched

   extrusion.

   117. The wingbox component of claim 107 which is welded to a second

   wingbox component and exhibits in its heat affected, welding zone an improved retention

   of one or more properties selected from the group consisting of: strength, fatigue, fracture

   toughness and stress corrosion cracking resistance.

   118. The wingbox component of claim 107 wherein said plate, extrusion

   or forged product was solution heat treated and intentionally quenched slowly for

   reducing quench distortion.

   119. The wingbox component of claim 107 which has, at a point 2 inches

   or more thick in cross section, a quarter-plane (T/4) tensile yield strength TYS in the

   longitudinal (L) direction and a quarter-plane (T/4) fracture toughness (K_lc) in the L-T

   direction at or above (to the right of) line M-M in Figure 7.

   120. The wingbox component of claim 107 which is plate -derived and

   has a minimum open hole fatigue life (S/N) at or above (to the right of) line A-A in

   Figure 12.

   121. The wingbox component of claim 107 which is forging-derived and

   has a minimum open hole fatigue life (S/N) at or above (to the right of) line B-B in Figure

   13.

   122. The wingbox component of claim 107 which has a maximum fatigue

   crack growth (FCG) rate in the L-T test orientation for a K (stress intensity factor) of 15

   ksiin or more at or below (to the right of) line C-C in Figure 14.

   123. The wingbox component of claim 107 which is capable of passing at

   least 30 days of alternate immersion, stress corrosion cracking (SCC) testing with a 3.5 %

   Na solution at a short transverse (ST) stress level of about 30 ksi or more.

   124. The wingbox component of claim 107 which has a minimum life

   without failure against sfress corrosion cracking after at least about 100 days of seacoast

   exposure at a short fransverse (ST) stress level of about 30 ksi or more.

   125. The wingbox component of claim 124 which has a minimum life

   without failure against stress corrosion cracking after at least about 180 days of said

   seacoast exposure conditions.

   126. The wingbox component of claim 107 which has a minimum life

   without failure against stress corrosion cracking after at least about 180 days of industrial

   exposure at a short fransverse (ST) sfress level of about 30 ksi or more.

   127. The wingbox component of claim 107 which has both thick and thin

   sections, said thin sections exhibiting an EXCO corrosion resistance rating of "EB" or

   better.

   128. The wingbox component of claim 107 which exhibits an improved

   resistance to hole crack initiation.

   129. A mold plate made from a thick aluminum alloy product consisting

   essentially of: about 6 to 10 wt.% Zn; about 1.2 to 1.9 wt.% Mg; and about 1.2 to 2.2

   wt.% Cu; optionally up to about 0.4 wt.% Zr, the balance Al, incidental elements and

   impurities.

   130. The mold plate of claim 129 wherein said alloy contains about 0.25

   wt.% or less Fe and about 0.25 wt.% or less Si.

   131. The mold plate of claim 129 wherein said alloy contains about 6.5 to

   8.5 wt.% Zn, about 1.3 to 1.65 wt.% Mg and about 1.4 to 1.9 wt.% Cu.

   132. The mold plate of claim 129 wherein said product is a rolled plate or

   forging and said alloy contains about 0.05 to 0.2 wt.% Zr.

   133. The mold plate of claim 129 wherein said product is a casting.

   134. A method for making a structural component that possesses an

   improved combination of at least two properties selected from the group consisting of:

   sfrength, fatigue, fracture toughness and corrosion resistance, said method comprising:

   (a) providing an alloy that consists essentially of: about 6.9 to 9 wt.%

   Zn; about 1.3 to 1.68 wt.% Mg; about 1.2 to 1.9 wt.% Cu, with wt.% Mg < (wt.%

   Cu +0.3); and about 0.05 to 0.3 wt.% Zr, the balance Al, incidental elements and

   impurities;

   (b) homogenizing and hot forming said alloy into a workpiece by one or

   more methods selected from the group consisting of: rolling, extruding and

   forging;

   (c) solution heat treating said workpiece;

   (d) quenching said solution heat treated workpiece; and

   (e) artificially aging said quenched workpiece.

   135. The method of claim 134 which further includes: (f) machining said

   structural component from the artificially aged workpiece.

   136. The method of claim 134 which optionally includes: sfress relieving

   the workpiece after quenching step (d) by stretching, compressing and/or cold working.

   137. The method of claim 134 which optionally includes: age forming the

   workpiece into a structural component shape.

   138. The method of claim 134 wherein said quenched workpiece is about

   3 to 12 inches at its thickest cross sectional point.

   139. The method of claim 134 wherein quenching step (d) includes spray

   or immersion in water or other media.

   140. The method of claim 134 wherein the workpiece is intentionally

   quenched slowly after solution heat treating step (c).

   141. The method of claim 134 wherein said alloy contains less than about

   8 wt.% Zn and less than about 1.8 wt.% Cu.

   142. The method of claim 134 wherein wt.% Mg wt.% Cu.

   143. The method of claim 134 wherein said alloy contains, as impurities,

   less than about 0.15 wt.% Fe and less than about 0.12 wt.% Si.

   144. The method of claim 134 wherein said workpiece is a plate product.

   145. The method of claim 134 wherein said workpiece is an extrusion.

   146. The method of claim 134 wherein said workpiece is a forged

   product.

   147. The method of claim 134 wherein artificial aging step (e) comprises:

   (i) a first aging stage within about 200 to 275°F; and

   (ii) a second aging stage within about 300 to 335°F.

   148. The method of claim 134 wherein artificial aging step (e) comprises:

   (i) a first aging stage within about 200 to 275 °F;

   (ii) a second aging stage within about 300 to 335°F; and

   (iii) a third aging stage within about 200 to 275°F.

   149. The method of claim 148 wherein said first aging stage (i) proceeds

   within about 230 to 260°F.

   150. The method of claim 148 wherein said first aging stage (i) proceeds

   for about 2 to 12 hours.

   151. The method of claim 148 wherein said first aging stage (i) proceeds

   for 6 or more hours within about 235 to 255°F.

   152. The method of claim 148 wherein said second aging stage (ii)

   proceeds for about 4 to 18 hours within about 310 to 325°F.

   153. The method of claim 152 wherein said second aging stage (ii)

   proceeds for about 6 to 15 hours within about 300 to 315°F.

   154. The method of claim 152 wherein said second aging stage (ii)

   proceeds for about 7 to 13 hours within about 310 to 325°F.

   155. The method of claim 148 wherein said third aging stage (iii)

   proceeds within about 230 to 260°F.

   156. The method of claim 148 wherein one or more of said first, second

   and third aging stages includes an integration of multiple temperature aging effects.

   157. The method of claim 134 wherein said structural component is for a

   commercial jet aircraft.

   158. The method of claim 157 wherein said structural component is

   selected from the group consisting of: a spar, rib, web, stringer, wing panel or skin,

   fuselage frame, floor beam, bulkhead, landing gear beam or combinations thereof.

   159. The method of claim 134 wherein said structural component has, at a

   point 2 inches or more thick in cross section, a quarter-plane (T/4) tensile yield sfrength

   TYS in the longitudinal (L) direction and a quarter-plane (T/4) plane- strain fracture

   toughness (K_lc) in the L-T direction at or above (to the right of) line M-M in Figure 7.

   160. The method of claim 134 wherein said structural component is a

   plate product having a minimum open hole fatigue life (S/N) at one or more ofthe applied

   maximum stress levels set forth in Table 12 equal to or greater than the corresponding

   cycles to failure value in said Table 12.

   161. The method of claim 134 wherein said structural component is a

   plate product having a minimum open hole fatigue life (S/N) at or above (to the right of)

   line A-A in Figure 12.

   162. The method of claim 134 wherein said structural component is a

   forging having a minimum open hole fatigue life (S/N) at or above (to the right of) line B-

   B in Figure 13.

   163. The method of claim 134 wherein said structural component has a

   maximum fatigue crack growth (FCG) rate in the L-T test orientation at or below at least

   one ofthe maximum da/dN values set forth in Table 14 for the corresponding K values at

   or greater than 15 ksiin in said Table 14.

   164. The method of claim 134 wherein said structural component has a

   maximum fatigue crack growth (FCG) rate in the L-T test orientation for a K (stress

   intensity factor) of 15 ksiin or more at or below (to the right of) line C-C in Figure 14.

   165. The method of claim 134 wherein said structural component is

   capable of passing at least 30 days of alternate immersion, sfress corrosion cracking

   (SCC) testing with a 3.5 % Na solution at a short fransverse (ST) sfress level of about 30

   ksi or more.

   166. The method of claim 134 wherein said structural component has a

   minimum life without failure against sfress corrosion cracking after at least about 100

   days of seacoast exposure at a short transverse (ST) stress level of about 30 ksi or more.

   167. The method of claim 166 wherein said structural component has a

   minimum life without failure against sfress corrosion cracking after at least about 180

   days of said seacoast exposure conditions.

   168. The method of claim 134 wherein said structural component has a

   minimum life without failure against stress corrosion cracking after at least about 180

   days of industrial exposure at a short fransverse (ST) stress level of about 30 ksi or more.

   169. The method of claim 134 wherein said structural component has both

   thick and thin sections, said thin sections exhibiting an EXCO corrosion resistance rating

   of "EB" or better.

   170. A method for making a jet aircraft structural component selected

   from the group consisting of: a spar, rib, web, stringer, wing panel or skin, fuselage

   frame, floor beam, bulkhead, landing gear beam or combinations thereof, said component

   having improved combinations of two or more properties selected from the group

   consisting of: strength, fatigue, fracture toughness and stress corrosion cracking

   resistance, said method comprising:

   (a) providing a wrought alloy consisting essentially of: about 6.9 to 9

   wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.2 to 1.9 wt.% Cu, with wt.% Mg <

   (wt.% Cu +0.3); and about 0.05 to 0.3 wt.% Zr, the balance Al, incidental elements

   and impurities;

   (b) homogenizing and hot forming said alloy into a workpiece by one or

   more methods selected from the group consisting of: rolling, extruding and

   forging;

   (c) solution heat treating said hot formed workpiece;

   (d) quenching said solution heat treated workpiece; and

   (e) artificially aging said quenched workpiece by a method comprising:

   (i) a first aging stage within about 200 to 275°F;

   (ii) a second aging stage within about 300 to 335°F; and

   (iii) a third aging stage within about 200 to 275°F.

   171. The method of claim 170 which optionally includes stress relieving

   the workpiece after quenching step (d) by stretching, compressing and/or cold working.

   172. The method of claim 170 which optionally includes age forming the

   workpiece into a near structural component shape.

   173. The method of claim 170 which further includes :

   (f) machining said structural component from the artificially aged

   workpiece.

   174. The method of claim 170 wherein first aging stage (i) proceeds for

   within about 230 to 260°F .

   175. The method of claim 174 wherein first aging stage (i) proceeds for

   about 2 to 12 hours within about 230 to 260°F .

   176. The method of claim 170 wherein second aging step (ii) proceeds

   within about 300 to 325°F.

   177. The method of claim 176 wherein second aging step (ii) proceeds for

   about 4 to 18 hours within about 300 to 325°F.

   178. The method of claim 177 wherein second aging stage (ii) proceeds

   for about 6 to 15 hours within about 300 to 315°F.

   179. The method of claim 177 wherein second aging stage (ii) proceeds

   for about 7 to 13 hours within about 310 to 325°F.

   180. The method of claim 170 wherein third aging stage (iii) proceeds

   within about 230 to 260°F.

   181. The method of claim 180 wherein third aging stage (iii) proceeds for

   at least about 6 hours within about 235 to 255°F.

   182. The method of claim 180 wherein third aging stage (iii) proceeds for

   about 18 hours or more at about 240 to 255°F.

   183. The method of claim 170 wherein one or more of said first, second

   and third aging stages includes an integration of multiple temperature aging effects.

   184. In a method for making a structural component from an aluminum

   plate, extrusion or forged product, the alloy of said product being substantially Cr-free

   and consisting essentially of: about 5.7 to 9.5 wt.% Zn; about 1.2 to 2.7 wt.% Mg; about

   1.3 to 2.7 wt.% Cu, and about 0.05 to 0.3 wt.% Zr, the balance Al, incidental elements

   and impurities, said method comprising the steps of: (a) solution heat treating said

   product; (b) quenching said solution heat freated product; and (c) artificially aging said

   quenched product, the improvement that imparts an improved combination of sfrength

   and toughness to said structural component, along with good corrosion resistance, said

   improvement comprising artificially aging said product by a method comprising:

   (i) a first aging stage within about 200 to 275°F;

   (ii) a second aging stage within about 300 to 335°F; and

   (iii) a third aging stage within about 200 to 275°F.

   185. The improvement of claim 184 wherein said alloy is selected from the

   group consisting of: 7050, 7040, 7150 and 7010 aluminum (Aluminum Association

   designations).

   186. The improvement of claim 184 wherein first aging stage (i) proceeds

   within about 230 to 260°F.

   187. The improvement of claim 186 wherein first aging stage (i) proceeds

   for about 2 to 12 hours within about 230 to 260°F.

   188. The improvement of claim 184 wherein first aging stage (i) proceeds

   for about 6 hours or more.

   189. The improvement of claim 184 wherein second aging step (ii)

   proceeds within about 300 to 325°F.

   190. The improvement of claim 184 wherein second aging step (ii)

   proceeds for about 6 to 30 hours within about 300 to 330°F.

   191. The improvement of claim 190 wherein second aging stage (ii)

   proceeds for about 10 to 30 hours within about 300 to 325°F.

   192. The improvement of claim 184 wherein third aging stage (iii)

   proceeds within about 230 to 260°F.

   193. The improvement of claim 192 wherein third aging stage (iii)

   proceeds for at least 6 hours within about 230 to 260°F.

   194. The improvement of claim 193 wherein third aging stage (iii)

   proceeds for about 18 hours or more within about 240 to 255°F.

   Il l

   195. The improvement of claim 184 wherein one or more of said first,

   second and third aging stages includes an integration of multiple temperature aging

   effects.

   196. The improvement of claim 184 wherein said product is at least about

   2 inches at its thickest cross sectional point.

   197. The improvement of claim 196 wherein said product is about 4 to 8

   inches at said thickest point.

   198. The improvement of claim 184 wherein said structural component is

   selected from the group consisting of: a spar, rib, web, stringer, wing panel or skin,

   fuselage frame, floor beam, bulkhead and/or landing gear beam for a commercial aircraft.

   199. A wing for a large aircraft, said wing including a wingbox comprised

   of upper and lower wing skins, at least one of said skins including a plurality of stringer

   reinforcements, said wingbox further including spar members spacing said wing skins, at

   least one of said spar members being an integral spar made by removing substantial

   quantities of metal from a thick aluminum product made from an alloy consisting

   essentially of:

   about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.3 to 2.1 wt.%

   Cu, with wt % Mg ≤ (wt.% Cu + 0.3); and about 0.05 to 0.2 wt.% Zr, the balance being

   Al, incidental elements and impurities.

   200. A wing for a large aircraft, said wing including a wingbox comprised

   of upper and lower wing skins, at least one of said skins including a plurality of stringer

   reinforcements, said wingbox further including upper and lower wing skins, at least one

   of said skins having an integral stringer reinforcement made by machining substantial

   quantities of metal from a thick wrought product, the alloy of which consists essentially

   of:

   about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.3 to 2.1 wt.%

   Cu, with wt % Mg < (wt.% Cu + 0.1); and about 0.05 to 0.2 wt.% Zr, the balance Al,

   incidental elements and impurities.

   201. A large aircraft having several large structural components, said

   components being made by removing substantial quantities of metal from thick aluminum

   workpieces, the alloy of which consists essentially of:

   about 6.9 to 8.5 wt.% Zn; about 1.3 to 1.68 wt.% Mg; about 1.3 to 2.1 wt.%

   Cu, with wt % Mg < (wt.% Cu + 0.3); and about 0.05 to 0.2 wt.% Zr, the balance Al,

   incidental elements and impurities.

   202. The large aircraft of claim 201 wherein at least one of said

   components is a bulkhead member.

   203. The large aircraft of claim 201 wherein two or more of said

   components are wing spars.