Products from aluminium alloy, which have improved combinations of properties

FIELD: metallurgy.

SUBSTANCE: plate from aluminium alloy consists of 6.8-8.5 wt % Zn, 1.75-2.3 wt % Cu, 1.5-1.84 wt % Mg and up to 0.25 wt % at least of one of Zr, Hf, Sc, Mn and V, and if necessary, additives crushing the grain, and the rest includes aluminium and inevitable impurities; at that, plate has the thickness of not more than 2.00 inches. Plate has the ratio between yield strength and failure viscosity, which meets the following equation: FT_LT ≥ -4.0* (TYS_L)+453. Plate has TYS_L that is at least 80 ksi and FT_LT that is at least 100 ksi√ inch, where TYS_L - tensile yield strength of the plate in direction L, in ksi, which is measured in compliance with ASTM E8 and ASTM B557, FT_LT - failure viscosity (Kapp) of the plate in flat stressed state in direction L-T, in ksi√inch, which is measured in compliance with ASTM E561 and B646 on specimen of aluminium alloy with central crack in position T/2 of the plate; at that, specimen has the width of 16 inches, thickness of 0.25 inches and initial length of preliminary fatigue crack of 4 inches.

EFFECT: plates are made from aluminium alloys having high failure viscosity at maintaining the acceptable strength level.

32 cl, 10 dwg, 10 tbl, 4 ex

 

The technical field to which the invention relates.

The present invention relates to aluminum alloys, in particular aluminum alloys ("A1") 7000 series (or 7XXX), according to the designations of the Association of the aluminium industry, more specifically, to products made of aluminum alloy, suitable for the manufacture of structural elements of commercial aircraft, having a thickness not more than 4 inches.

The level of technology

Industry requirements for aluminum alloys are becoming more and more stringent with each new series of aircraft produced by the aerospace industry. With increasing size of new jet aircraft or as escalated modern model jet liners to accommodate greater payload and/or to fit more range, to improve operational quality and efficiency continue to increase demands to reduce the weight of such elements as wing components.

The design of the conventional airplane wing shown in figure 1 and includes the torsion box, indicated generally by the number 2. The caisson 2 wing passes outward from the fuselage as the main force component of the wing and runs essentially perpendicular to the plane of Figure 1. In the caisson 2 wing upper and lower skin of the wing 4 and 6 separated by a vertical design is klonnie elements, or spars, 12 and 20 passing between the upper and lower skins of the wing or overlapping them. The caisson 2 wing also contains ribs, which usually runs from one side member to the other. These ribs are parallel to the plane of Figure 1, while covering the wing and the side members are perpendicular to the plane of Figure 1.

The top floor of the wing usually consists of a casing 4 and stiffeners or stringers 8. These stiffeners can join individually by attachment or be made as one piece with the casing, to avoid the need for a separate stringers and rivets. During the flight, the upper structure of the wing commercial aircraft experiences a compressive loads that require alloys with high compressive strength. This requirement led to the creation of alloys with more and more high compressive strength while maintaining the nominal level of fracture toughness. The upper structural members of the wing of a modern large aircraft are usually made of high strength aluminum alloys 7XXX series, such as aluminum 7150 (reissued U.S. patent 34,008), 7449 (U.S. patent 5 560 789) or 7055 (U.S. patent 5 221 377). More recent U.S. patent 7 097 719 describes the improved aluminum alloy 7055.

However, the development of aircraft ultrahigh capacity has led to new design requirements. Because Bo is the larger and heavier wing and high total takeoff weight of the aircraft these aircraft have high bending down load when landing, causing high tensile forces in the upper structural members of the wing. Although the tensile strength in the modern alloys for the upper wing is more than sufficient to withstand this bending load, the fracture toughness of them becomes the limiting settlement criterion for located aboard the areas of the upper surface. This has led to the search for alloys for the upper structural elements superliner with very high fracture toughness, closer to the viscosity of alloys for the lower skin of the wing, such as 2324 (U.S. patent 4 294 625), even if you had to sacrifice to some extent high strength. That is, there has been a shift of the optimum combination of strength and fracture toughness necessary for the maximum reduction of weight of the upper structural members of the wing superliner, in the direction substantially higher fracture toughness and lower strength.

New technologies of welding, such as friction welding with stirring, also opened up many new possibilities for designs and products from alloys for use in the wing spars and components of the ribs in order to reduce weight and/or cost reduction. For maximum performance spar part of the spar, which connects with the upper wing skin must have properties which a, close to the properties of the upper casing, and the portion of the spar, which connects with the lower wing skin, should have properties similar to the properties of the lower skin of the wing. This has led to the use of composite spars, containing the upper zone of the spar, 14 or 22, the wall 18 or 20 and the lower zone of the spar, 16, or 24, United fastening means (not shown). This "composite" design allows you to use products made of alloys, the optimum for each component. However, the introduction of a number of fastening means requires increasing the cost of Assembly. The mounting means and the hole may also be weak structural links, and you may need to increase the thickness of the parts, which reduces the gain in the properties from multiple alloys.

One approach, used in order to avoid increasing the cost of installation associated with composite spars, consists in machining the entire spar of a thick sheet, stamping or forging from a single alloy. Sometimes this machining is known as the "cut" part. With this design is the need to create a connection wall/upper spar and sidewall/lower spar disappears. One-piece spar, made in this way are sometimes called "solid spar". The ideal alloy for the manufacture of the part spars must have the strength characteristics of the alloy for the upper wing in combination with the fracture toughness and other characteristics resistance to damage appropriate alloys for the bottom wing. Usually, to achieve both characteristics simultaneously is difficult, and requires a compromise between the requirements of the properties of the upper cladding and the lower cladding. One drawback to be overcome by one-piece spar, is that the characteristics of strength and fracture toughness of thick products, used as the starting material, usually lower than thinner products commonly used in composite spars, even if the whole spar made of the same alloy and at the same vacation. Thus, a compromise in properties and application of thick products for solid spar can lead to poor weight gain. One alloy for thick products, which sufficiently meets the requirements to properties of both the upper and lower belts of the spar and retains good properties even in thick products due to its low sensitivity to training, is the alloy 7085 described in U.S. patent 6 972 110. Another disadvantage of the one-piece side members, regardless of the alloy is high purchase weight (i.e. material that is procured) to the flight weight (i.e. the weight of the material, flying an airplane), known as the "purchase/flight". This is at least partially reduces the gain in the value of a solid spar compared to the composition of the second side member, attainable due to the lowered cost of the Assembly.

However, new technologies such as friction welding with stirring, made further enhancements to weight and cost. Multicomponent spar, assembled by friction welding with stirring or other modern methods of welding or connection, combines the advantages of composite and solid spars. The use of such methods allows the use of products with a smaller thickness, and use a few of alloys, shapes of products and/or types of leave that are optimized for each component of the spar. This expands the options of the product/vacation alloy and improves the ratio of "purchase/flight" for the material as a composite spar, while maintaining a significant portion of the benefits from saving on Assembly, as in the whole spar.

U.S. patent 5 865 911 describes the alloy of the 7000 series, intended for use in structural elements of the lower skin of the wing and in the elements of the wing spar planes ultra-high capacity. This alloy exhibits improved strength, fracture toughness and fatigue strength in thin sheet form, compared with the mainstream alloys for the lower part of the wing, such as 2024 and 2324 (U.S. patent 4 294 625). Similar characteristics of strength and fracture toughness were obtained in the alloy 708 (U.S. patent 6 972 110) in sheet form, as shown in Table 1. Any of these alloys in thin form would be suitable for structural elements of the lower cover of the wing and the lower zone of the spar and the walls of the multi-spar, United mechanical fastening or welding. These alloys are also suitable for use in rib as for the composite, and solid design. However, levels of strength achieved in these alloys, it is generally not sufficient for use in the upper structural members of the wing of large commercial aircraft. The higher strength is advantageous for the upper zone of the side member, the walls of the spar and ribs, provided that there is adequate fracture toughness.

Klc, Kq (ksi√inch)
room temperature
Table 1
The properties of the alloy Miyasato (U.S. patent 5 865 911) and alloy 7085 (U.S. patent 6 972 110) in sheet form
PropertyDirectionMiyasato(1)7085(2)
UTS (ksi)L
LT
82,1
81,4
82,6
82,2
TYS (ksi)L
LT
76,2
75,4
78,0
77,2
L-T
T-L
47,5
40,7
44,0
35,9
Klc, Kq (ksi√inch)
-65F
L-T
T-L
42,0
no data
40,5
34,3
Kapp (ksi√inch)
room temperature
L-T
T-L
120,8
94,3
128,7
104,4
Kapp (ksi√inch)
-65F
L-T
T-L
115,5
74,7
106,8
79,0
Kc (ksi√inch)
room temperature
L-T
T-L
172,9
123,9
165,7
129,1
Kc (ksi√inch)
-65F
L-T
T-L
166,4
79,8
140,1
84,8
(1)U.S. patent 5 865 911: rolled sheet with a thickness of 1.2 inches, a width of 86 inches
(2)7085, U.S. patent 6 972 110; rolled sheet with a thickness of 1.5 inches, a width of 102 inches.

Thus, for aircraft ultrahigh capacity there is a need in the alloy, which has a mean is correctly higher fracture toughness, than modern alloys used in the upper structural members of the wing, while maintaining an acceptable level of strength. Such an alloy would also be useful for application in the upper zone and the wall of the multi-spar, United mechanical fastening or welding, as well as for wing ribs integral or one-piece construction. Although the particular discussed the needs of the planes ultra-high capacity and wings, this alloy may also be advantageous for use in the fuselage and on smaller aircraft with composite and solid structures. In addition, this alloy can also be details that are not relevant to the aerospace industry, such as armor for military vehicles.

Disclosure of invention

Proposed articles of a new aluminum alloy that is particularly well suited for structural components in the aerospace industry. In one aspect of new aluminum alloys (sometimes called here "the alloys of the present invention") contains from about 6,80 to about 8.5 wt.% Zn, from about 1.5 or 1.55 to about 2.00 wt.% Mg, from about 1.75 to about 2,30% wt. Cu; from about 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn, less than about 0.05 wt.% Cr, the rest mainly Al, minor elements and impurities. Products from alloy to have a thickness of about 4 inches or less is the more, sometimes a thickness of about 2.5 or 2.0 inches or less, having a significantly higher fracture toughness than the alloys of the prior art used in these applications, while maintaining acceptable levels of strength, and Vice versa.

In one embodiment provides an article of aluminum alloy. Aluminum alloy product consists essentially of: from about 6,80 to about 8.5 wt.% Zn, from about 1.5 or 1.55 to about 2.00 wt.% Mg, from about 1.75 to about 2,30% wt. Cu; from about 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn, less than about 0.05 wt.% Cr, and the rest are aluminum, non-essential elements and impurities. Aluminum alloy may be being subjected to a heat treatment in solid solution, conventional quenching and artificial aging, and parts made from these products, the improved combination of strength and fracture toughness. In one embodiment, the alloy contains low amounts of iron and silicon impurities. In one embodiment, the alloy comprises not more than about 0.15 wt.% Fe and not more than about 0.12 wt.% impurities Si. In one embodiment, the alloy comprises not more than about 0.08 wt.% Fe and not more than about 0.06 wt.% impurities Si. In one embodiment, the alloy comprises not more than about 0.04 wt.% Fe and not more than about 0.03 wt.% impurities Si. The aluminum can be in the form of a rolled t is nkiye sheets, rolled sheets, stampings or forgings. In some embodiments of the product of the alloy has a thickness of less than 2.5, or 2.0 inches at its thickest point. In some embodiments of the product of the alloy has a thickness of from about 2.5 inches to 4 inches at its thickest point.

In one embodiment, the aluminum alloy is in the form of a rolled sheet having a thickness of less than 2.5 inches, for example, the thickness of 2.00 inches. In one embodiment, the aluminum alloy sheet contains 6,8-8,5% wt. Zn, 1.5 to 2.0 wt% Mg, about 1.75 to 2.3 wt.% Cu, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,95 wt.% aluminum. In one embodiment, the aluminum alloy contains 7.5 to 8.5 wt.% Zn, 1.9 to 2.3 wt.% Cu, 1.5 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and to about of 89.1 wt.% aluminum. In one embodiment, the aluminum alloy contains about 7.8-8.5 wt.% Zn, 1,95-2.25 wt.% Cu, 1.7 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,55 wt.% aluminum. In one embodiment, the aluminum alloy contains a 7.9-8.2 wt.% Zn, 2,05-2,15% wt. Cu, 1,75-of 1.85 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88.3% wt. aluminum. In one embodiment, the aluminum alloy contains between 7.4 to 8.0 wt.% Zn, 1,95-2.25 wt.% Cu, 1.7 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,95% wt. aluminum. In one embodiment, the Alu is injuy alloy contains 7.5 to 7.9 wt.% Zn, 2,05-2,20% wt. Cu, 1.8 to 1.9 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,65% wt. aluminum. In various of these embodiments, the aluminum alloy may contain from 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn and less than about 0.05 wt.% Cr. In any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities. In any of these embodiments, the product of the alloy may have a thickness of less than about 2.5, or 2.0 inches at its thickest point.

In one embodiment, the aluminum alloy is used in the form of a sheet having a thickness of 2.5 or 3.0 inch, or of 2.51 inches to about 3.5 inches, 3.75 inch or 4 inch. In one embodiment, the aluminum alloy sheet contains 6,8-8,5% wt. Zn, 1.5 to 2.0 wt% Mg, about 1.75 to 2.3 wt.% Cu, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,95 wt.% aluminum. In one embodiment, the aluminum alloy contains between 7.4 to 8.0 wt.% Zn, 1.9 to 2.3 wt.% Cu, of 1.55 to 2.0 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,15% wt. aluminum. In one embodiment, the aluminum alloy contains 7.5 to 7.9 wt.% Zn, 2,05-2,20% wt. Cu, 1.6 to about 1.75 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,55 wt.% aluminum. In various of these embodiments and miniawy alloy may contain from 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn and less than about 0.05 wt.% Cr. In any of these embodiments, the aluminum alloy aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities.

The product of the alloy can realize improved characteristics of strength and fracture toughness. In one embodiment, the product of the alloy contains a plot of thickness not more than about 2.5 inches or 2.00 inches, with a minimum yield strength in tension in the longitudinal direction and the fracture toughness under plane deformation in the direction L-T, lying on or above and to the right of line A-A in FIGU or Figv (i.e. in the shaded area). In one embodiment, the alloy contains a plot of thickness not more than about 2.5 inches or 2.00 inches, having a yield strength in tension and apparent fracture toughness in a plane stress state in the direction L-T, lying on or above and to the right of line B-B in figure 4 (i.e., shaded area), measured on the plate of a width of 16 inches with a Central crack with an initial crack length (2ao) about 4 inches and a thickness cracks about 0.25 inches.

In one embodiment, the product of the alloy contains a plot of thickness from about 2.00 or 2.5 inches to 3.0, or 3,125, or 3.25 inches, with a limit of those is uchesti under tension in the direction of LT (long transverse) and fracture toughness under plane deformation in the direction T-L, lying on or above and to the right of the line C-C 7 (i.e., shaded area). In one embodiment, the product of the alloy contains a plot of thickness from about 2.00 or 2.5 inches to 3.0, or 3,125, or 3.25 inches (for example, at the thickest point) and has a yield strength in tension in the direction of the ST (short transverse) and fracture toughness under plane deformation in the direction of S-L, lying on or above and to the right of the line E-E in Figure 9 (i.e., shaded area).

In one embodiment, the product of the alloy contains a plot of thickness from about 2,75, 3,0, 3,125 or 3.25 inches to about 3.5, 3.75, or 4 inches (for example, at the thickest point) and has a minimum yield strength in tension in the direction of LT and fracture toughness under plane deformation in the direction T-L, lying on or above and to the right from the line D-D in Fig (i.e. in the shaded area). In one embodiment, the product of the alloy contains a plot of thickness from about 2,75, 3,0, 3,125 or 3.25 inches to about 3.5, 3.75, or 4 inches and has a minimum yield strength in tension in the direction of ST and fracture toughness under plane deformation in the direction of S-L, lying on or above and to the right from the line F-F in Figure 10 (i.e. in the shaded area).

The product of the alloy may also provide excellent corrosion resistance. In one embodiment, the product of the Plava is the evaluation of the corrosion resistance by EXCO "EB" or better. In one embodiment, the product of the alloy consistently withstands the test on resistance to stress corrosion cracking under tension with alternative dive when the voltage level of 35 ksi for a vacation T74, when the voltage level of 25 ksi for a vacation T76 and when the voltage level of 15 ksi for a vacation T79. In one embodiment, the product of the alloy consistently withstands the test on resistance to stress corrosion cracking under stress, in terms of the sea coast when the voltage level of 35 ksi for a vacation T74, when the voltage level of 25 ksi for a vacation T76 and when the voltage level of 15 ksi for a vacation T79. In one embodiment, the product of the alloy invariably reaches evaluate the corrosion resistance by EXCO "EB" or better and always stand the test of resistance to stress corrosion cracking under tension with alternate immersion tests on resistance to stress corrosion cracking under stress, in terms of the sea coast when the voltage level of 35 ksi for a vacation T74, when the voltage level of 25 ksi for a vacation T76 and when the voltage level of 15 ksi for a vacation T79. In one embodiment, the product of the alloy invariably reaches evaluate the corrosion resistance by EXCO "EB" or better and consistently withstands both tests on resistance to stress corrosion cracking under the strain the m alternative immersion and tested for resistance to stress corrosion cracking under stress, in terms of the sea coast when the voltage level of 35 ksi for a vacation T74, when the voltage level of 25 ksi for a vacation T76 and when the voltage level of 15 ksi for a vacation T79, and reaches the above-described characteristics of the yield strength tensile and fracture toughness. The product of the alloy can withstand other tests on resistance to stress corrosion cracking under the strain.

The product of the alloy can be used in many applications. In one embodiment, the product of the alloy is a structural element for the aerospace industry. The element of the aircraft structure may be any of the top wing panels (cladding), the upper stringers of the wing, the upper surface of the wing with integral stringers, belt spar, the wall of the spar, rib, leg rib or wall ribs, stiffeners, and their combination. In one embodiment, the product of the alloy is a part of the fuselage (for example, covering of the fuselage). In one embodiment, the product of the alloy is a component of armor (for example, motorized vehicles). In one embodiment, the product of the alloy used in the oil and gas industry (such as pipes, structural elements).

Products of alloy can be obtained in a variety of ways. For example, of the products of the alloy can be made a component that, to get it, weld by fusion welding methods or welding in the solid phase with one or more products of aluminum alloy, is made essentially of the same alloy with the same or another vacation. In one embodiment, the product of the alloy is connected with one or more products from aluminum alloy of different composition, to obtain a component of several alloys. In one embodiment, the product connect mechanical fastening. In one embodiment, the product of the alloy are connected by fusion welding methods or welding in the solid phase. In one embodiment, the product of the alloy sostarivayut separately or after connection with other products from the alloy in the manufacturing process of the component. In one embodiment, the product of the alloy reinforce the layered materials of metal fibers or other reinforcing materials.

Are also ways of obtaining aluminum alloys and goods made of aluminum alloy. In one embodiment, the method includes the steps of forming or molding of aluminum alloy structural member of the aircraft. The method can include obtaining or purchasing aluminum alloy, t is anyone as aluminum alloy, have any of the above structures, homogenization and hot processing of the alloy to one or more method selected from the group consisting of rolling, stamping and forging, methods of heat treatment on the solid solution of the alloy, quenching and removal of internal stresses in the alloy. The element of design in conditions of artificial aging can be improved combination of strength and fracture toughness. In one embodiment, the alloy has a thickness of less than about 4 inches during hardening. In one embodiment, the method includes aging component, one or after connection to other components.

In one embodiment, the step of forming or molding of the structural element includes mechanical processing. In one embodiment, the mechanical treatment is carried out after artificial aging, or between one of the stages of aging. In one embodiment, machining before heat treatment is carried out in solid solution.

In one embodiment, the step of forming or molding of the structural element includes forming ageing before or after the connection with other components. In one embodiment, at least part of the step of forming or molding of the structural element is performed before or during the about at least part of the artificial aging.

In one embodiment, the alloy is artificially age the way, including: (i) the first stage of aging in the temperature range from about 150 to about 275°F, and (ii) the second stage of aging in the temperature range from about 290 to about 335°F. In one embodiment, the first stage of aging (i) occurs in the temperature range from about 200 to about 260°F. In one embodiment, the first stage of aging (i) occurs in a period of from about 2 to about 18 hours. In one embodiment, the second stage aging occurs within from about 4 to about 30 hours in the temperature range from about 290 to about 325°F. In one embodiment, the second stage of aging (ii) occurs in a period of from about 6 to about 30 hours in the temperature range from about 290 to about 315°F. In one embodiment, the second stage of aging (ii) occurs in a period of from about 7 to about 26 hours in the temperature range from about 300 to about 325°F. In one embodiment, one or both stages of aging include introduction multiple effects of thermal ageing. In one embodiment, one or both stages of aging interrupted, to weld the item to another component of the same or different alloy or other vacation.

In another embodiment, the alloy is artificially age the way, including: (i) the first stage of starane is in the temperature range from about 290 to about 335°F and (ii) the second stage of aging in the temperature range from about 200 to about 275°F. In one embodiment, the first stage of aging (i) occurs in a period of from about 4 to about 30 hours in the temperature range from about 290 to about 325°F. In one embodiment, the first stage of aging (ii) occurs in a period of from about 6 to about 30 hours in the temperature range from about 290 to about 315°F. In one embodiment, the first stage of aging (i) occurs in a period of from about 7 to about 26 hours in the temperature range from about 300 to about 325°F. In one embodiment, one or both the aging stage include the introduction of multiple effects of thermal ageing. In one embodiment, one or both stages of aging interrupted, to weld the item to another component of the same or different alloy or alloy with another vacation.

In another embodiment, the alloy is artificially age the way, including: (i) the first stage of aging in the temperature range from about 150 to about 275°F, (ii) the second stage of aging in the temperature range from about 290 to about 335°F and (iii) the third stage of aging in the temperature range from about 200 to about 275°F. In one embodiment, the first stage of aging (i) occurs in the temperature range from about 200 to about 260°F. In one embodiment, the first stage of aging (i) occurs within about 2 to about 18 hours. In one embodiment, the implementation of the of the second stage of aging (ii) occurs in a period of from about 4 to about 30 hours in the temperature range from about 290 to about 325°F. In one embodiment, the second stage of aging (ii) occurs in a period of from about 6 to about 30 hours in the temperature range from about 290 to about 315°F. In one embodiment, the second stage of aging (ii) occurs in a period of from about 7 to about 26 hours in the temperature range from about 300 to about 325°F. In one embodiment, the third stage of ageing (iii) occurs for at least about 2 hours in the temperature range from about 230 to about 260°F. In one embodiment, the third stage of ageing (iii) occurs within about 18 hours or more in the temperature range from about 240 to about 255°F. In one embodiment, one, two or all stages of aging include the introduction of multiple effects of thermal ageing. In one embodiment, one, two or all stages of aging interrupt to weld the item to another component of the same or different alloy or alloy with another vacation.

The method or methods may enable the connection of the components of the alloy. In one embodiment, one or more components are joined by mechanical fastening. In one embodiment, one or more components are joined by welding. In one embodiment, the components are welded by electron beam welding. In one embodiment, the components are welded on what redstem friction welding with stirring. In one embodiment, the component is attached or welded to the other aluminum product to get a component of several alloys and/or one alloy with multiple vacations.

As you can appreciate, different from the above-mentioned aspects, approaches and/or embodiments can be combined to give a variety of useful products and components from aluminum alloy. These and other aspects, advantages and novel features of the invention are set out in the following part of the description and should be understandable to experts in this field after studying the following description and figures, or can be studied by carrying out the invention in practice.

Brief description of drawings

For a more complete understanding of the present invention reference is made to the following description taken in conjunction with the attached drawings, on which:

Figure 1 represents a cross section of a conventional design of the torsion box of an aircraft;

Fig. 2A and 2B are embodiments of the composition of the alloy according to the invention in relation to the major alloying elements Cu and Zn and Mg and Zn in comparison with the alloy compositions family 7085 and 7055 and 7449, respectively;

Fig. 2C-1, 2C-2, 2D-1 and 2D-2 represent different embodiments of the composition of the alloy of the present invention, such chakotay, suitable for receiving sheets of aluminum alloy, having a thickness not more than 2 or 2.5 inches;

Fig. 2E and 2F represent various embodiments of the composition of the alloy of the present invention, such as compositions suitable for receiving sheets of aluminum alloy, having a thickness of at least about 2 or 2.5 inches;

Figa is a graph illustrating the normal dependence of fracture toughness under plane L-T deformation Klcfrom a minimum longitudinal yield strength tensile for (i) examples of alloy A-D in the form of a sheet and to leave T79, and (ii) several other conventional alloys in sheet form;

FIGU is a graph illustrating the normal dependence of fracture toughness under plane L-T deformation, Klcfrom a minimum longitudinal yield strength tensile for (i) examples of alloy A-D in the form of a sheet and leave T79, and (ii) to several other conventional alloys in the form of a sheet;

Figure 4 is a graph illustrating the normal dependence of fracture toughness in a flat L-T stress state, Kappfrom the actual or measured yield strength tensile for (i) examples of alloy A-D in the form of a sheet and leave T79 and (ii) to several other conventional alloys in the form of a sheet;

Figure 5 is a graph, compare the non share of retained strength after exposure to corrosion in the direction of LT for two examples of the composition of the alloy and three times the third stage of aging (0, 6 and 12 hours);

6 is a graph comparing the percentage of retained strength after exposure to corrosion in the direction of LT for example alloy and alloy 7055 prior, when the duration of the second stage aging 12 hours;

7 is a graph illustrating the normal dependence of fracture toughness in the plane T-L strain, Klcfrom the usual LT yield strength tensile sheet (i) of example alloy E (having a thickness of 3.125 inches) and leave T74, and (ii) several other conventional alloys (thickness of about 3 inches);

Fig is a graph illustrating the normal dependence of fracture toughness in the plane T-L strain, Klcfrom the usual LT yield strength tensile sheet of (i) of example alloy F (thickness of 4.0 inches) leave T74 and (ii) for some other conventional alloys (having a thickness of about 4 inches);

Fig.9 is a graph illustrating the normal dependence of S-L fracture toughness under plane deformation, Klcfrom plain ST yield strength tensile sheet of (i) of example alloy E (having a thickness of 3.125 inches) leave T74 and (ii) several other conventional alloys (having a thickness of about 3 inches); and

Figure 10 is a graph illustrating the normal dependence of S-L fracture toughness when packagefile, Klcfrom plain ST yield strength tensile sheet of (i) of example alloy F (having a thickness of 4.0 inches) leave T74 and (ii) several other conventional alloys (having a thickness of about 4 inches).

In all the drawings the same reference positions denoted by the same elements.

Detailed description of the invention

Figure 1 is a schematic illustration of a cross-sectional view of conventional design of caisson 2 wing containing the top 4 of the wing skin and stringers 8, the bottom 6 of the wing skin and stringers 10, separated side members 12 and 20. The stringers 4 and 10 can join separately by attachment or be made as one piece with the casing to eliminate the need for separate stringers and rivets. Depending on the aircraft size and design of the wing usually requires two, three or four wing panels, 4 or 6, to close each of the upper and lower surfaces of the wing. More consoles may be required for one piece design plating and stringers. The combination of consoles, comprising upper and lower casing, usually combined mechanical fastening. These compounds increase the weight of the aircraft.

The spars may have a composite construction consisting of the upper zone of the side member 14 or 22, the lower zone of the spar 16 or 24 and the wall 18 or 26 connected mechanically increase the population, or they may have a solid one piece construction, with each design type has its advantages and disadvantages. Composite spar allows you to use the products of the optimum alloy for each component of the spar and has a superior attitude "purchase/flight" compared with a solid spar. Usually the upper zone of the spar requires high strength in compression and the lower zone of the spar requires less strength, but the best characteristics of resistance to damage, such as fracture toughness and resistance to the growth of fatigue cracks. One-piece spar has a much lower cost to the Assembly, but its performance may be lower than that of composite construction, as its properties are inevitably a compromise between the requirements of the upper casing and the lower casing. Also, the strength and fracture toughness of thick products, used as starting material for the solid spar, usually less than thinner products used for the composite spar.

The torsion box also contains a rib (not shown), which are generally from one side member to the other. These ribs are parallel to the plane of Figure 1, while covering the wing and the side members are perpendicular to the specified plane Figure 1. Like spars, ribs can also have the ü integral or one-piece construction, and each type has advantages and disadvantages similar to the advantages or disadvantages in the case of the spars. However, the optimal properties of the ribs are slightly different because high strength is best for legs rib, which is connected with the upper and lower skins of the wing and stringers, and wall ribs best increased stiffness. More typically, the ribs of the wing have a solid design with a compromise of properties between the requirements of the foot of the rib and the wall of the rib.

New technologies of welding, such as friction welding with stirring and electron-beam welding, make possible a new concept designs, preserving the advantages of modern composite and solid structures, while minimizing their disadvantages. For example, various console 4 wing, used to make the upper panels may be joined by friction welding with stirring, and not a mechanical bond, thereby reducing the weight of the upper cladding. The spars and ribs can be made of several alloys with multiple vacations and/or products that are optimized for each component of the spar or rib connected by friction welding with stirring, thereby retaining the benefits of features and the best relationship purchase/flight" thinner products, as in the composite spar, andat the same time reducing the cost of the Assembly, as in the whole spar or rib. For example, the upper zone of the side member 14 and 22 may be made from forgings of high-strength tempered alloy or alloy, the lower zone of the spar 16 and 24 of forgings from less durable alloy or tempered alloy with less resistance to damage, and the walls of the spar 18 and 26 of the sheet of tempered alloy or alloy with moderate strength, and all component are joined by friction welding with stirring or electron beam welding. Can be applied to designs that are a combination of solid and composite structures to improve the preservation of the health of failure of individual elements and the resistance to damage of the component while reducing the cost of Assembly. For example, the upper zone of the side member 14 and 22 can be joined by friction welding with stirring with the walls of the spar 12 and 20, to reduce the cost of the Assembly, and the lower zone of the spar 16 and 24 can be fastened mechanically to improve resistance to damage. To further improve resistance to damage integral and one-piece welded constructions and structures, which is a combination of both types can be achieved by hardening of the layered materials of metal fibers and other reinforcing materials, as described in U.S. patent 6 595 467.

The alloy described in atente USA 6 972 110, having the trade designation 7085, mainly focused on large calibers, usually from 4 to 8 inches or more, with low sensitivity to hardening. Low sensitivity to hardening is achieved by creating a carefully controlled composition, which allows quenching of thick gauges, still achieving superior combination of high strength and fracture toughness and high corrosion resistance compared with previous thick products from alloys such as 7050, 7010 and 7040. A carefully controlled composition, registered as AA7085 includes low levels of Cu (from about 1.3 to about 1.9 wt.%) and low levels of Mg (from about 1.3 to about 1,68 wt.%), that is within the lowest levels used in alloys for commercial aviation. The levels of Zn (from about 7 to about 9.5 wt.%), at which properties were most optimized, consistent with the levels that are much higher than nominal levels for 7050, 7010 and 7040. This was contrary to previous doctrines that high concentrations of Zn increases the sensitivity to hardening. Conversely, elevated levels of Zn in the alloy 7085 actually turned out to be advantageous in respect of the slow quenching of thick composite parts. U.S. patent 6 972 110 teaches that a significant part of improving the strength and fracture toughness for thick sections of alloy disclosed is Tim patent, required a special combination of alloying elements.

U.S. patent 5 221 377 refers to the alloy 7055, which is typically used for plates and forgings thickness of 2 inches or less, and teaches that the lower levels of Mg leads to improved fracture toughness. In the prior art, it was also widely accepted that increasing the strength by increasing the content of dissolved substances usually leads to reduction of fracture toughness.

The alloy of the present invention is aimed primarily at the more subtle products of alloy, with a thickness of about 4 inches or less, and sometimes a thickness of about 2.0 or 2.5 inches or less for the upper structural members of the wing of a large commercial aircraft, including the covering of the wing, the wing stringers and upper zones of the spar. These applications will benefit from increased strength and in many cases will require a higher strength than that achieved by a composition 7085. Similarly, higher strength can be advantageous in other applications, such as the walls of the spar, ribs and other products of the aerospace industry. To increase the strength, the level of Mg in the alloy of the present invention is increased from about 1.5 or 1.55 to about 2.0 wt%, and the level of Cu from about 1.75 to about 2,30 wt.%. The range of Zn somewhat reduced, to about 6.8-8.5 wt.%. Fig. 2A and 2B illustrate options for the implementation of the management composition of the alloy of the present invention in terms of the main alloying elements Cu and Zn and Mg and Zn in comparison with compositions 7085 (U.S. patent 6 972 110), 7055 (U.S. patent 5 221 377) and 7449. Suitable compositions of the alloy of the present invention is indicated by the rectangle delineated by solid lines. On Figa and 2B also shows the compositions for examples of alloys A-F, described below.

In one embodiment, the alloys of the present invention have the form of a sheet of a thickness of less than 2.5 inches, for example, the thickness of 2.00 inches. In one embodiment, the aluminum alloy sheet contains 6,8-8,5% wt. Zn, 1.5 to 2.0 wt% Mg, about 1.75 to 2.3 wt.% Cu and up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,95 wt.% aluminum (for example, as shown in Fig. 2A and 2B). In other embodiments, implementation and according to Fig. 2C-1, 2C-2, 2D-1 and 2D-2, aluminum alloy contains 7.5 to 8.5 wt.% Zn, 1.9 to 2.3 wt.% Cu, 1.5 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and to about of 89.1 wt.% aluminum (as is given of the embodiment 1 of Fig. 2C-1 and 2C-2). In another embodiment, the aluminum alloy contains about 7.8-8.5 wt.% Zn, 1,95-2.25 wt.% Cu, 1.7 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,55 wt.% aluminum (as is given of the embodiment 2 of Fig. 2C-1 and 2C-2). In one embodiment, the aluminum alloy contains a 7.9-8.2 wt.% Zn, 2,05-2,15% wt. Cu, 1,75-of 1.85 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88.3% wt. aluminum (as is given of the embodiment 3 of Fig. 2C-1 and 2C-2). In one embodiment, the aluminum alloy is gain of 7.4 to 8.0 wt.% Zn, 1,95-2.25 wt.% Cu, 1.7 to 2.0 wt% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,95% wt. aluminum (as is given of the embodiment 4 with Fig.2D-1 and 2D-2). In one embodiment, the aluminum alloy contains 7.5 to 7.9 wt.% Zn, 2,05-2,20% wt. Cu, 1.8 to 1.9 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,65% wt. aluminum (as is given of the embodiment 5 with Fig.2D-1 and 2D-2). In various of these embodiments, the aluminum alloy may contain from 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn, less than about 0.05 wt.% Cr. In any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities.

In another embodiment, the aluminum alloy used in the sheet, having a thickness of from about 2.01 inches or 2.51 inches to about 3.5 inches, 3.75 inch or 4 inch. In one embodiment, the aluminum alloy sheet contains 6,8-8,5% wt. Zn, 1.5 to 2.0 wt% Mg, about 1.75 to 2.3 wt.% Cu, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,95 wt.% aluminum (for example, as shown in Fig. 2A and 2B). In other embodiments, the implementation according to Fig. 2E and 2F, the aluminum alloy contains between 7.4 to 8.0 wt.% Zn, 1.9 to 2.3 wt.% Cu, of 1.55 to 2.0 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 89,15% wt. aluminum (as given option implemented the program with 1 File and 2F). In one embodiment, the aluminum alloy contains 7.5 to 7.9 wt.% Zn, 2,05-2,20% wt. Cu, 1.6 to about 1.75 wt.% Mg, up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and up to about 88,55 wt.% aluminum (as is given of the embodiment 2 with Five and 2F). In various of these embodiments, the aluminum alloy may contain from 0.05 to about 0.3 wt.% Zr, less than about 0.1 wt.% Mn, less than about 0.05 wt.% Cr. In any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities.

Of the indications of U.S. patent 6,972,110 change in the composition of the alloys according to the present invention is to increase the sensitivity of the alloy to a heat treatment in comparison with alloy 7085, and it is quite likely actually. However, the alloys of the present invention, appears to retain some dignity composition 7085, and, in any case, sensitivity to quenching is less of a problem in a more subtle products of alloy, the focus of the alloys of the present invention. It was expected that changes in the composition will have a negative effect on the fracture toughness because of the resulting increase strength, and due to the higher content of Mg. When the level of Mg between content 7085 and existing spaventa the upper part of the wing, 7055 and 7449, it was believed that the strength and fracture toughness of the alloys of the present invention will occupy an intermediate position between these alloys. Indeed it is for strength. However, the combination of strength and fracture toughness of the alloys of the present invention was superior not only in comparison with alloys 7055 and 7449, as expected, but, unexpectedly, also in comparison with alloy 7085. Thus, the alloys of the present invention can be identified as unexpectedly pleasant range of compositions, which offers the best combination of strength and fracture toughness than find the prevailing alloys.

Products from alloys of the present invention can be obtained more or less traditional methods, including casting of the melt and the casting in the mold direct cooling (DC) in the form of an ingot, and show the distinctive features of the internal structure of the origin of the ingot. As is well known in this field can also be used conventional additives, grinding grain, as, for example, supplements containing titanium and boron or titanium and carbon. After the ingot was cast from this composition, it skallerud (if necessary) and homogenized by heating to one or more temperatures from about 800° to about 900°F, or from about 850° to about 900°F. After homogenization of these if the CI process, for example, rolling into sheet or a thin sheet or by stamping or forging in special shapes. For most aerospace applications, products made of alloy, made from a composition of the present invention, have a thickness in the cross section of about 4, 3.75, or 3.5 inch or less, and sometimes a thickness of about 2.5 or 2.0 inches or less. If necessary, the product should then be subjected to heat treatment in solid solution by heating to one or more temperatures from about 850° to about 900°F to enter a substantial portion, sometimes the whole or substantially the whole of the soluble zinc, magnesium and copper in the solution, and it is understood that the physical processes that are not always ideal, it is possible that every last a small number of these basic components of the alloy will not be dissolved during the heat treatment in the solid solution. After heating to elevated temperatures, as described, the product should be cooled down quickly, or to harden to complete the procedure of heat treatment on the solid solution. Such cooling is typically performed by immersing in a suitable size vessel with cold water or a water shower. As an additional or alternate means of cooling can also be used for air cooling. After quenching some products may require mechanical removal of the internal voltage is th, such as stretching and/or compression to about 8%, for example from about 1% to about 3%.

It is believed that the product, after heat treatment for solid solution hardening, with or without cold treatment, is then possible precipitation hardening or ready to artificial aging. Technology can be a two-stage or three-stage and for some applications it may be sufficient even single-stage technologies. However, between each stage or phase may not be clean borders. It is well known that increasing and/or decreasing temperature from the target (or target) value temperature processing itself can effect the precipitation hardening (aging), which can, and often should be taken into account when imposing such conditions of temperature changes and their effects precipitation hardening in a full program of aging. This introduction was described in more detail in U.S. patent 3,645,804, the description of which is included here by reference.

In U.S. patent 6 972 110, the description of which is included here by reference, describes a three-stage mode for aging alloy 7085. Three-stage mode aging temperature ranges that are identical or close to the range in the patent '110 can also be used with the alloy of the present invention, but for some principalin the x applications you can imagine, is also suitable for 2-stage mode. The two-stage mode can be either a stage with a low temperature, followed by high-temperature phase, or Vice versa. For example, 2-stage mode is often used for the upper skin of the wing and stringers. These components are often formed with aging aircraft manufacturers to obtain the profile of the wing. When molding with aging item is held in the mold at an elevated temperature, usually from about 250 to about 400°F in continuation from several to tens of hours, and the desired profile is performed by the processes of creep and stress relieving. Molding with aging often performed together with the operation of artificial aging, especially at the high temperature stage, when creep is the fastest. Molding with aging is usually carried out in the furnace autoclave. Autoclave and molds, which require that molded aging wing for large commercial aircraft that are large and expensive and, as a result, rarely used in the manufacturing process. Therefore, it is desirable that the molding cycle with aging was as short as practicable, but still allowing you to achieve the desired profile and product properties of the alloy, so that the performance was poppy who Kalnay. To achieve this advantageous reduction in the duration of the third stage or complete rejection of it. In two-stage low-temperature mode, the first stage may be carried out by the manufacturer of the alloy, which further reduces time spent in the molding process of aging.

The results of studies of corrosion stress cracking (SCC) on the examples of the alloy indicate that, indeed, the third stage can be reduced and even to abandon it, at the same time meeting the requirements for SCC of the upper skin of the wing and stringers. Three-stage procedure applicable to alloy 7085 in the field of thick products, for alloys of the present invention when used in the upper part of the wing and other high strength applications for several reasons usually not needed. For example, requirements for SCC of the upper components of the wing is less strict than in the area of thick products, such as rib or spar. The upper components of the wing are predominantly compressive stresses, whereas the spar, in particular, the lower part, is subjected to tensile stresses. Only tensile stresses contribute to the SCC. Also, one-piece spar or rib machined thick products can have a significant rated voltage in the direction of ST. For example, belt integrally what about the spar, made of sheets are in the plane of the L-ST of the original sheet. For comparison, the basic design stresses in the upper hull and stringer lie mainly in the plane L-LT, which is less prone to SCC. As a result of this difference the minimum SCC in the direction of ST for lords alloys for the upper wing, 7055 and 7449, be 15 or 16 ksi, which allows to use these alloys for vacation T (high strength), whereas thick products for spars, ribs, and other applications commonly used in conditions of leave T76 and T74 (lower strength), which typically have minimal SCC 25 ksi 35 ksi, respectively.

The alloys of the present invention also provides for applications consisting of several multicomponent alloys spars or ribs connected by mechanical fastening or welding. As already described, these applications are likely to have higher requirements for SCC than for the upper skin of the wing and stringers. However, in multicomponent spar made of thinner products, granular structure can be more favorably oriented in relation to resistance to SCC than for solid spar, machined from thick plate. For example, the belt of the spar may be machined from more than stand the Oh to SCC plane L-LT source worksheet or from forgings instead of plane L-ST. The minimum characteristics of SCC in the directions L and LT are usually greater than 40 ksi, even when less resistant to SCC of high-strength tempered alloys, compared with 25 ksi 35 ksi in the direction of ST for less durable tempered alloys with higher resistance to SCC. Thus, it may be that the third stage of the regime of aging, often used for alloy 7085, you can also reduce or abandon it in the case of the alloys of the present invention, even for spars, ribs, and other applications with higher requirements for SCC. Reduction or waiver of the third stage actually leads to a small decrease in strength, usually from about 1 to about 2 ksi. However, it may be that this reduction in strength can be compensated by using more high-strength holidays, is not feasible in thick products. Even if this is so, for some integral, one-piece or multi-component applications of the present invention may be desirable lower strength after the holidays, such as T74 or T73, or to provide additional corrosion resistance, or to further improve the fracture toughness.

In case consisting of several alloys of the spars or ribs connected by welding, a desirable characteristic is the flexibility of the aging process, exercise alloys on nastasemarian. Methods of welding: welding by melting or welding in the solid phase, for example, friction welding with stirring can be carried out at intermediate vacation and not at the end-leave alloy so as to improve the strength and corrosion properties of the weld is usually desirable aging after welding. For example, the welding alloy of the present invention with another alloy having characteristics of strength and resistance to damage, more suitable for the lower zone of the spar may be carried out after carrying out the first stage, 2 - or 3-stage mode aging of the alloy of the present invention. Another alloy may be another alloy 7XXX series or alloy of a completely different composition, for example, lithium-aluminum alloy, in accordance with U.S. patent 4 961 792, which has its own normal aging, which may consist of one, two or three stages. As aging after welding two of the United products of alloy must inevitably be carried out together, the mode of aging for alloys of the present invention may require two or three stages depending on the requirements of the aging of the alloy with which it is connected. Thus, the flexibility of the alloys of the present invention in relation to the number of stages and the duration of aging, which can be successfully used, is advantageous for welded components, with the present of several alloys. Even in this case, it may require some compromise with normal aging for each alloy, depending on the specific alloys used.

Manufacturing and aging contains several component alloys using the alloys of the present invention, to be joined by welding, can be somewhat simplified by using the 7XXX alloys with compositions similar to the compositions of the alloys of the present invention, but which are more depleted or more enriched with alloying elements added for hardening to achieve the desired balance of strength and fracture toughness in each component. Normal modes of aging before and after welding these alloys would be probably more compatible than for stronger distinguished alloys, as it requires fewer adjustments to their normal operating procedures. Alternatively, the desired difference in strength and fracture toughness in some cases can be achieved, apparently, by the application alone alloys of the present invention, using different releases. For example, spar with multiple vacations, made of alloys of the present invention can use high-strength leave T79 in the top shelf, vacation T76 (moderate strength, high fracture toughness) in the wall of the spar, and the OTP is IC T73 (lower durability the highest fracture toughness) in the lower zone of the spar. Typically, the duration of aging for the holidays T76 and T73 should be higher than for a vacation T79. In welded spar with several types of vacation aging before welding to the upper spar with vacation T79 could, for example, contain only the first stage, for the walls of the spar leave T76 - to contain the first stage and part of the second stage, and for the lower belt spar with vacation T73 - first stage and a more significant part of the second stage. This could be carried out separately on each component, or by dispersing time of their removal from the same furnace. After welding you can use the same mode of aging linked components. With proper selection of the mode of aging before and after welding normal aging can be applied to each component essentially without compromise.

Example 1

Bars A-D having compositions similar to the above-described variants of the implementation of a family of alloys of the present invention, were cast as large ingots industrial scale. In addition, the quality control was cast a single ingot of aluminum alloy 7085. The ingots were scalped and homogenized with a final holding temperature of from about 870 to about 900°F. one ingot of alloys A and B were subjected to the mountains is whose rolling in the sheet, having a thickness of 1.07 inches and a width of 135 inches. Another strand of each of the alloys A and B was subjected to hot rolling into a sheet having a thickness of 1.10 inches and a width of 111 inches. The first will be called hereinafter the Sheet 1, and the last Sheet 2. One of ingots of alloys C and D were subjected to hot rolling to the same thickness and width as the Sheet 2. The size of the Sheet 1 and Sheet 2 is indicative for the top consoles wing ultra-high capacity. Control alloy 7085 was subjected to hot rolling to the same thickness and width as the Sheet 1. The sheets were subjected to heat treatment in solid solution at temperatures from about 880 to about 895°F for about 70-100 minutes, was tempered with the irrigation cooled to ambient temperature and subjected to cold drawing about 1.5-3%. Samples from sheets of alloys A - D and control alloy 7085 aged very early because to high holiday type T79 suitable for the upper components of the wing, using the traditional three-stage mode of aging (e.g., as given by the U.S. patent 6 972 110). Three-stage mode consists of the first stage, lasting about 6 hours at about 250°F, the second stage is about 7 hours at about 308°F and the third stage of about 24 hours at about 250°F. in Addition, samples improved version of the aluminum alloy 7055 (U.S. patent 7 097 719) were cut from several different batches of sheets of the same or close to the width and thickness and subjected to a high-strength leave T7951 and multiple vacations with pedestrianism to reduce strength and increase fracture toughness. The composition of the ingot A-D and formulations of various conventional alloys shown in Table 2. The mode of ageing to leave T7951 improved version of alloy 7055 was the two-stage mode, which consists of the first stage lasts for approximately 10 hours at 302°F and the second phase lasting 6 hours. Perestaranie released alloys were obtained by increasing the duration of the first stage with about 10 hours to about 19-24 hours.

Table 2
Alloywt.% Znwt.% Cuwt.% Mgwt.% Fewt.% Siwt.% Zr
A7,71,811,620,0240,0140,11
B7642,151,650,0280,0210,10
C8,052,081,780,0440026 0,12
D7,832,171,84being 0.0360,0200,11
Sample 70857,61,621,48to 0.0320,0150,11
7085 series AA7,0-
8,0
1,3-2,01,2-
1,8
max 0,08max 0,060,08-0,15
7055 superior7,6-
8,4
2,0-2,61,8-
2,3
Max.
0,09
max 0,060,08-0,25
7055 series AA7,6-
8,4
2,0-2,61,8-
2,3
Max.
0,15
max 0,100,08-0,25
7449 AA7.5 to
8,7
1,4-2,11,8-
2,7
max 0,15max 0,12(1)
(1)Zr + Ti max 0,25.

Measured tensile strength and compression, fracture toughness under plane strain (Klcand apparent fracture toughness in a plane stress state (Kappand resistance to delamination for examples of alloy A-D, alloy 7085 and improved control alloy 7055. Tensile test was carried out according to standard test ASTM E8 and ASTM B557, and tested in compression in accordance with ASTM E9. Tests for fracture toughness under plane strain (Klc) was performed in accordance with ASTM E399. Samples for testing fracture toughness under plane deformation had a full plate thickness and the width W of 3 inches. Tests for fracture toughness in a plane stress state (Kapp) was performed in accordance with ASTM E561 and B646. Professionals need to understand that the numerical value of Kappusually increases with the width of the sample for testing. For Kappalso influenced by the thickness of the sample, the initial crack length and the geometry of the cut sample. Thus, the values of Kappcan reliably be compared only to the test samples equivalent geometry is Irina, thickness and initial crack length. Accordingly, all tests on the examples of the alloy and the control alloy 7085 and 7055 were conducted using specimens with a Central crack M(T)having the same nominal dimensions, width 16 inch, thickness 0.25 inch, and the initial length of the pre-fatigue cracks (2ao) 4 inches. The samples were centered on the mid-thickness (T/2) of the sheet. Also were tested for delamination using the method EXCO in accordance with ASTM G34. Samples for testing were taken from mid-thickness (T/2) and with one-tenth the thickness (T/10).

The measured characteristics of examples of alloy A-D and the nominal composition 7085 are shown in Table 3. Alloy A was found to increase the yield strength at elongation and ultimate tensile strength of about 3 ksi compared with the nominal composition 7085 for the size of the Sheet 1 in the direction as L and LT, i.e. the increase in strength of about 4%; and alloy B was found to increase the yield strength at elongation and ultimate tensile strength of about 5 ksi, i.e. an improvement of about 6%. Alloys C and D showed even higher strength. The increase of yield strength and ultimate tensile strength for both alloys was about 7 ksi, i.e. an improvement of about 8%. Aircraft manufacturers is considered a significant improvement in strength. Improving strength was recip is but while maintaining excellent resistance to delamination, all of the samples of the examples of the alloy was achieved assessment EA.

Table 3
Alloy/PanelE.g.UTS (ksi)TYS (ksi)CYS (ksi)Klc(ksi√inch)Kapp (ksi√inch)EXCO
Sample 7085L
LT
83,7
83,7
79,9
79,6
81,4
no data
50,6
41,1
128,9
102,6
EA (t/2)
EA(t/10)
Example alloy A
Page 1
L
LT
86,7
86,8
83,2
82,6
84,3
no data
50,9
40,8
127,5
94,0
EA (t/2) EA(t/10)
Example alloy A
Sheet 2
L
LT
85,8
85,7
81,7
81,5
83,0
no data
49,1
39,6
to 129.2
91,9
EA(t/2) EA(t/10)
Example alloy B
Page 1
L
LT
89,3
89,2
85,7
85,0
86,7
no data
43,8
34,2
of 113.2
78,6
EA(t/2)
EA(t/10)
Example alloy B
Sheet 2
L
LT
87,8
88,5
84,3
84,1
86,4
no data
43,6
34,5
129,1
86,0
EA (t/2) EA(t/10)
Example alloy CL
LT
90,2
90,2
87,2
84,6
86,5
no data
36,0
30,0
115,6
71,2
EA (t/2)
EA t/10)
Example alloy DL
LT
90,4
90,6
87,1
86,5
86,2
no data
40,1
31,5
107,9
68,8
EA (t/2) EA(t/10)

The combination of strength and is Ascoli destruction for examples of alloy A-D shown in Fig. 3A, 3B and 4, where they are compared with the alloys of the prior art. Figa and 3B compare the fracture toughness under plane deformation Klcorientation L-T, which corresponds to the main direction of the load on the upper part of the wing as a function of the minimum yield strength in tension in the direction L (rolling direction) for examples of alloy A-D, batch control samples of alloy 7085 (table 3), the other four parties of thin sheets of alloy 7085, which was subjected to aging for less strength, more suitable for the lower part of the wing (table 1), and values for an enhanced version of the alloy 7055 after the holidays T7951 and after pereustroennogo vacation. In addition, shows the typical fracture toughness Klcfor other alloys of the prior art in sheet form. For examples of the alloy and perestroennih tempered alloy 7055, for which currently there are no technical requirements for the materials, a minimum yield strength tensile was estimated by subtracting 3 ksi from the measured values. One line minimum specifications for alloys of the present invention indicated by the line A-A, which corresponds to the equation FT = -2,3*(TYS)+229 where TYS has a longitudinal yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, and where FT is eating the L-T fracture toughness under plane deformation of the sheet, in ksi√inch, measured in accordance with ASTM E399.

Figa also includes a shaded region, illustrating the potential properties of thin sheet products of the alloy of the present invention. The shaded region is bounded by the minimum L-T fracture toughness of 36 ksi√inch, a minimum strength of 74 ksi and a line A-A, which corresponds to the equation FT = -2,3*(TYS)+229, as described above. The shaded area on Figa particularly well suited for rolled products of alloy vacation T74, although it is possible to obtain alloys with other leave (e.g., T6, T73, T75, T79), which can have properties that lie in the shaded area.

Figv also includes a shaded region, illustrating the potential properties of thin sheet products of the alloy of the present invention. The shaded region is bounded by the minimum fracture toughness of 30 ksi√inch, a minimum strength of 79 ksi and a line A-A, which corresponds to the equation FT = -2,3*(TYS)+229, as described above. The shaded area on FIGU particularly well suited for rolled products of alloy vacation T76, although it is possible to obtain alloys with other issues (for example, T6, T73, T74, T79), which can have properties that lie in the shaded area.

Figure 4 compares the longitudinal yield strength in tension and the effective fracture toughness in a plane stress state (Kapp for embodiments of the alloys of the present invention in the orientation L-T, again with the five parties alloy 7085, and values for an enhanced version of the alloy 7055. Improved combination of strength and fracture toughness of the alloy 7085 compared with the improved version of alloy 7055 obvious. One line minimum specifications for alloys of the present invention indicated by the line B-B, which corresponds to the equation FT = -4,0*(TYS)+453 where TYS has a longitudinal yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT is an L-T fracture toughness (Kapp) sheet in plane stress, ksi√inch, and FT measured in accordance with ASTM E561 and B646 on the sample of aluminum alloy with a Central crack, taken from the position of T/2 sheets of aluminum alloy, and the sample has a width 16 inch, thickness 0.25 inch, and the initial length of the pre-fatigue crack 4 inches.

Even with significant perestiani, to obtain the same or similar level of strength, as in the alloy of the present invention, the fracture toughness of alloy 7055 much lower. As the levels of Cu and Mg in the alloys of the present invention lie between the levels for alloy 7085 and improved version of alloy 7055, and the levels of Fe and Si similar low, it was expected that the combination of strength and toughness times is osenia, achieved by the alloys of the present invention, will be between the characteristics of the alloy 7085 and improved alloy 7055. It was unexpectedly found that the alloys of the present invention exhibit a superior combination of strength and fracture toughness, not only better than that of alloy 7055, but than alloy 7085. Thus, embodiments of the alloys of the present invention establish a "nice" area of compositions, which offers the best combination of strength and fracture toughness than are alloys of the prior art. Although the values of Kappand a relative improvement corresponds to a test cut samples of the specified type and size, it is expected that a close relative improvement is observed for cut samples of other types and sizes. However, experts in this field should understand that the actual values of Kappcan vary significantly in samples of other types and sizes, as described previously, and that the amplitude difference can also be changed.

Figure 4 also contains a darkened area, illustrating the potential properties of rolled products of alloy according to the present invention. The shaded region is bounded by the minimum fracture toughness (Kapp) 100 ksi√inch, minimum yield strength tensile 80 ksi and a line B-B that soo is the same as equation FT = -4,0*(TYS)+453, as specified above. The shaded region in figure 4 is particularly suitable for rolled products of alloy vacation T79, although it is possible to obtain alloys with other leave (e.g., T6, T73, T74, T76), which can have properties lying within the shaded area. In addition, some sheet products of the present invention may be able to realize the value as fracture toughness in plane stress and yield stress tensile, asked the shaded area of Figure 4, and fracture toughness under plane strain and yield strength tensile defined by the shaded area in Fig. 3A and/or 3B.

Example 2

Four sets of samples, after undergoing a heat treatment for solid solution, conventional quenching and extraction (vacation W51), from a sheet of examples of alloy A and B obtained in Example 1, have been subjected to the first two stages of the three-stage aging mode used in Example 1. Then the first set of samples was subjected to a third stage with a duration of aging 24 hours, the same as was used in Example 1, and the second and third sets were subjected to a shorter duration of aging, 6 and 12 hours. At the fourth set of samples of the third stage of aging was not performed (0 hours). Samples for tensile test with a diameter of 0.125 inch mechanically processed in the length of the nom transverse direction (LT) and short transverse direction (ST) for both tests: the test for resistance to stress corrosion cracking under tension with alternate immersion and testing in Maritime coast (SC) (sometimes referred to here as a positive test for resistance to stress corrosion cracking under stress, in terms of the sea coast). Alternatively, the air-tightness test with immersion was conducted in accordance with ASTM G44, G47 and/or G49. More specifically, the samples were subjected to cycles of immersion in 3.5%aqueous NaCl solution for 10 minutes, which was a 50-minute air drying under constant deformation required to achieve the desired voltage level. Tests in the conditions of the sea coast was held at atmospheric station in Pt. Judith, Alcoa, RI, as described below.

These three duration of the third stage of aging (0, 12 and 24 hours) and two levels of stress (16 and 20 ksi) were selected for ST-direction. The first voltage level is a minimum requirement for modern alloys for the upper wing, 7055 and 7449, in the direction of ST. The second level voltage corresponds to the voltage level is 25% higher. The period of exposure to AI-testing 7XXX alloys in ST-direction is typically 20 or 30 days or until corruption will occur. In these tests for AI expected maximum period of exposure at 150 days to better assess the performance of the different modes of aging. For the conditions of the sea coast the maximum period of exposure was 466 days. The results of the tests on orrosion cracking under tension (SCC) are given in Table 4.

Table 4
AlloyPanelThe time of the 3rd stage (h)Stretching LT (ksi)SCC-testDamage (days)
YSUTSPlaceStress (ksi)The number of trials
A2082,586,2AI16548, 101, 101, 101, 115
AI20 532,59,70, 101,115
SC165297,311
SC205290, 290, 339, 349
A212is 83.887,4AI16578,97, 101
AI 20553,98, 101, 101, 101
SC165325, 339
SC20566, 325, 339, 367
A22483,7of 87.3AI165101, 101, 101, 115, 129
AI 20544,73,98, 101, 143
SC165332
SC205332, 346, 346, 402
A11284,187,6AI16587, 129, 143, 143
AI 20559,98, 101, 101, 101
SC165325, 332, 332, 339
SC205325, 332, 339
B21285,589,2AI165115,135, 135
AI 20529,54, 101,101, 115
SC165234, 332
SC205122,311,325

The results for the sample of alloy A, panel 2, when the duration of the third stage of aging 0 (i.e. without the third stage), 12 and 24 hours indicate that there is no significant difference in the resistance to SCC of the alloys of the present invention with or without the third stage of aging or for a shorter or longer duration of the third stage of aging. In all cases, the number of days prior to the damage exceeded the standard exposure time is 20 or 30 days for 7XXX alloys for AI-tests on SCC as when the voltage level of 16 ksi (minimum requirement for modern alloys for the upper wing), and when the voltage level, 25% higher (20 ksi). The number of days until the damage for three different times of aging was also close Resistance to SCC for three times the third stage of aging was also similar to the resistance in the conditions of the sea. Alloy A, panel 1 and the example of alloy B, panel 2, was estimated only for the midnight of the third stage of aging. Panel 1 is thinner and wider than the panel 2, and therefore it is expected that it will have a different aspect ratio grain and possibly other resistance to SCC. It turned out that the results for alloy A, panel 1, slightly better than the results for the panel 2. Results for alloy B, panel 2, were similar and possibly better than for alloy A, panel 2.

SCC tests were also carried out in the direction of LT. For the LT direction of impact was interrupted after 30, 47 and 90 days, and experiencing the samples were subjected to destructive test load in accordance with ASTM G139. Was determined by the percentage withheld or residual strength of the exposed sample in comparison with the tensile strength is not exposed to the sample. The voltage levels for the LT direction were 42 and 63 ksi, which is approximately 50% and 75% of the yield strength in the LT direction for alloys of the present invention. This test was a means to obtain more quantitative information in a shorter time, and thus, it is useful for the more resistant to SCC LT-direction, where it is expected that the damage of the sample will occur over a longer time, and perhaps with greater variation than the La less resistant to SCC ST-direction. In one experiment tested breaking load was carried out on examples of alloy A and B, for a given duration of the third stage of the aging process 0, 6 and 12 hours, after a period of exposure to 47 days. In the second experiment was conducted destructive testing load on the example of alloy A and the control alloy 7055-T7951 after periods of exposure 30 and 47 days in the AI test and after 90 days of exposure when tested under the conditions of the sea coast, at voltage levels corresponding to 50 and 75% of the yield strength in the LT direction for each alloy. In both experiments also participated relaxed samples. The inclusion of an unstressed and stressed samples allows to divide the strength loss resulting from General corrosion and pitting corrosion and loss from SCC.

The results of the first experiment is shown in Figure 5, where each dot represents the average of 5 samples. Here, the share of retained strength is the ratio of the strength has experienced the impact of the sample to the strength of the sample is not exposed to (i.e., decorationrole), expressed as a percentage. The results indicate that the failure of the third stage of aging was not losses total corrosion resistance (stress) or resistance to SCC (under voltage). Indeed, the samples without the third stage had a higher withheld or residual is prochnosti, than samples with a six - or twelve-hour third stage. For a given length of aging the alloy B was superior to alloy A. the results of the second experiment is shown in Fig.6, each point of which corresponds to an average over 5 samples. 6 is a graph comparing the percentage of retained strength in the direction of LT for alloys of the present invention and alloy prior 7055 for a 12-hour duration of the second stage of aging, followed by maintaining for 30 and 47 days in 3.5%NaCl solution and 90 days in the conditions of the sea coast when the voltage levels of 50 and 75% of the yield strength of each alloy. Example alloy has A higher percentage of retained strength than alloy 7055, for all three effects, as in loose conditions and under stress and at two levels of stress.

In General, the results of corrosion indicate that two-and three-stage mode aging provides acceptable corrosion characteristics of the alloys of the present invention for use in areas of the upper wing. One drawback of the 2-stage mode is that the strength is slightly lower, as shown in Table 4 for an example of alloy A. In comparison with the third stage of ageing duration of 24 hours, the yield strength without the third stage was about 1 ksi above. As described earlier, ibcast mode aging alloys of the present invention is a favorable characteristic. The two-stage mode typical applications, such as the upper wing skin and the stringer, where aging is partly or completely performed in the molding process of aging by aircraft manufacturers or subcontractors, and it is desirable that the molding cycle with aging was as short as practical to maximize performance. In this respect, the alloys of the present invention with the 2-step mode, which has a full time 13 hours, offer improved compared with modern alloys for the upper wing. Depending on the requirements of aging, it could be reduced to about 7 hours, if the first stage is carried out by the manufacturer of the material, and only the second stage is carried out in the molding process of aging.

Three-stage mode can be used when the material is supplied by the manufacturer in full aging, for applications such as upper wing spar or the wall of the spar in the composite structure. Less durable tempered alloy, such as vacation T76 or vacation T74, can also be used for these applications when using any of the 2 - or 3-stage mode depending on the requirements and direction of the estimated voltage relative to the orientation of the grains in the h is x alloy. When the alloys of the present invention should be welded to the products of other alloy and aging after welding as part of a component consisting of several alloys, can be used 2 - or 3-stage mode depending on the mode of ageing of the alloy or alloys, which should connect the alloys of the present invention. The flexibility provided by the alloys of the present invention, may be useful to combine cycles vulcanization adhesives used to attach the reinforcing materials, with aging of the alloys of the present invention.

Example 3

Samples of the sheet of example alloy A, obtained in Example 1 under the conditions of heat treatment on the solid solution hardening and drawing (vacation W51), mechanically processed in panel thickness 0.5 inch, a width of 6 inches and a length of 35 inches. Samples from forgings of alloy 2099 were purchased after the holidays T3511 and machined to the same dimensions. In both cases, the length corresponded to the direction of rolling. Alloy 2099 is commercially available lithium-aluminum alloy, a registered Association of the aluminium industry, with the composition of 2.4 to 3.0 wt.% Cu, 0.1-0.5 wt.% Mg, 0,4-1,0 Zn, 0.1 to 0.5 Mn, 0,05-0,12 Zr and 1,6-2,0 Li, the rest being Al and incidental impurities. The panel of example alloy A and alloy 2099 connected by friction welding with paramasivan the m line melting along the length of the panels. This combination of alloys according to the present invention and alloy 2099, which have very different compositions, can be used, for example, consisting of several alloys of the spar or rib. In the spar alloys of the present invention could be used in the top shelf and the wall where high compressive strength and alloy 2099 in the lower zone of the spar, which benefits high resistance to the growth of fatigue cracks. Similarly, the rib alloys of the present invention could be used in the leg, where high strength and alloy 2099 - in the wall of the spar, where the best high stiffness and low density.

Prior to the friction welding with mixed panel of alloys A and 2099 aged very early because individually. Aging before welding for the alloy consisted of A first phase lasting 6 hours at 250°F, and the mode of aging before welding for alloy 2099 consisted of the first and second stages of different length and/or at other temperatures than those used for alloys of the present invention. The aging process after welding the United panels inevitably was the same and consisted of a first phase lasting 6 hours at 250°F and a second phase lasting 18 hours at 305°F. Aging after welding it is desirable to improve the strength and corrosion properties of the weld zone. To improve the shape characteristics of the weld, in particular, the strength and corrosion resistance after welding should be stronger aging. However, for different alloys the possibility of this may be limited due to the different requirements of aging for each alloy and the final desired vacation for everyone. Modes of aging before welding for each alloy and modes of aging after welding for the panel, consisting of several alloys, was carefully selected in order to achieve the vacation type T76 in the alloys of the present invention and vacation type T in the alloy 2099. Even in this case, was required some compromise between modes of aging in both alloys, and the flexibility of the alloys of the present invention in regard to the number of stages and the duration of aging, which can be used to obtain good properties, was in this respect advantageous.

Mechanical properties, including tensile strength, compressive strength, modules of elasticity in tension and compression and fracture toughness were measured for base metals (i.e. outside of the weld and heat affected zone (HAZ), in the heat-affected zone (HAZ) and the weld seam, followed by aging after welding. The size of each zone and the position of the samples was determined using measurements of hardness Vickers hardness (VHN) across the joint and optical microsync is. The tests were conducted in accordance with appropriate test methods ASTM: ASTM E8 and B557 for tensile test, E9 for compression testing, E111 testing module tension and compression and ASTM E399 for fracture toughness under plane deformation. Mechanical tensile properties were measured in the directions L and LT. The compressive strength and the elastic modulus was measured only in the L-direction. Samples for testing fracture toughness under plane deformation was in orientation T-L, had a width W of 2 inches and a total thickness of the panel. Samples for testing to destruction cut out of the panel so that their machined slots (representing the expected plane of propagation of cracks) were aligned with the interests of the region. Two samples were taken from the weld and zone HAZ, one sample with a machined recess oriented in the same direction, which was held welding tool during friction welding with stirring, and the one with the machined groove oriented in the opposite direction. The results of these tests are given in Table 5.

11,5
Table 5
DescriptionDirectionAlloy A Weld2099
Base metalHAZHAZBase metal
UTS (ksi)L
LT
84,556,861,277,683,1
84,362,4*77,2
TYS (ksi)L
LT
79,8
79,1
43,959,069,976,0
50,3*70,6
CYS (ksi)L82,369,560,976,2
Et (106f/DM2)L10,310,411,5
Ec (106f/DM2)L10,710,711,311,811,9
Klc, Kq (ksi√inch)T-L41,534,21,
36,22
40,51,
38,22
26,41,
27,12
32,1
*samples for testing the tensile strength in the LT direction, crossed the weld and HAZ, breaking in the weak spot.
1The propagation of cracks in the same direction, which is traversed by the welding tool during the welding process.
2The propagation of cracks in the opposite direction, which is traversed by the welding tool during the welding process.

Even when you compromise mode aging carried out at each alloy, the base metal of each of the alloy which was subjected to aging before welding (different for each alloy) and the process of aging after welding (the same for all alloys), has achieved the required level of strength and fracture toughness for the target leaves. Properties in the HAZ area and over REGO seam were lower as is usually observed for joints. The area of the seam was mainly termoobrabotka in solid solution during the process of friction welding with stirring, so that the artificial aging in this area was only by aging after welding. Similarly, the HAZ area is also heated during the welding process, but to temperatures that are lower than used in the heat treatment for solid solution and, therefore, insufficient for the complete dissolution of the alloying elements. This may limit the response to ageing in the zone of the HAZ during aging after welding and impair its strength and fracture toughness. Despite these factors, the achieved efficiency of welding (i.e. the ratio of the strength of the weld to the strength of the base metal) was pretty good. When measured perpendicular to the line of sealing, where the sample for tensile tests included the weld and the HAZ area, welding performance was 71% for yield strength tensile (TYS) and 81% for ultimate tensile strength compared to the strength of the base metal alloy 2099 in the direction of LT.

Fracture toughness obtained in the weld and in the HAZ area, were also satisfactory. In the welding zone fracture toughness was equivalent to the viscosity of the base metal alloy A, whereas the fracture toughness in the HAZ area on the side seam as the JV is ava A, and alloy 2099, was lower than their corresponding base metal, but still sufficient to meet the requirements of most aircraft.

Tests on stress corrosion cracking stress (SCC) and a test for delamination was carried out also in the United panels with subsequent aging after welding. For SCC-flat test samples type of samples for tensile test thickness 0,235 inch mechanically processed at the mid-thickness perpendicular to and across the weld and the HAZ area. Each of the three samples was tested at two levels of voltages, 26 and 35 ksi by alternate immersion in accordance with ASTM G44, G47 and/or G49. No damage was observed after a period of exposure to 250 days. To test for delamination two rectangular sample with the full thickness of the panel containing the weld zone, HAZ and base metal of the alloy was tested using test method EXCO in accordance with ASTM G34. This test method is appropriate accelerated test method for 7XXX alloys, such as alloys of the present invention. The second set of samples from the full thickness of the panel was tested using Dry Bottom MASTMAASIS in accordance with ASTM G85. This test method is appropriate accelerated test method for alloy 2099. Base metal alloy A is alloy 2099 was the evaluation of delamination EA. This evaluation serves as an indicator of good corrosion characteristics and is consistent with the usual characteristics of the target vacations for each alloy. The area of the weld, which contained a mixture of both alloys had the EB evaluation according to test method EXCO, which again indicates a fairly good resistance to exfoliation corrosion. It is expected some deterioration of the corrosion characteristics of the weld, as this area is only aging after welding. The area of the HAZ in the alloy 2099 was the evaluation of P on MASTMAASIS, however, the HAZ in the alloy shows A localized corrosion and has a rating of ED by EXCO. These corrosion characteristics can be unacceptable to the internal structures of the aircraft, such as spars and ribs, but they probably can be improved by optimization of the parameters of friction welding with the mixing or application of cooling processes during welding to reduce the flow of heat in the HAZ. This area can also be protected during the operation, using the methods of protection against corrosion. For example, prior to anodizing and coating anti-corrosive primer, which is already widely used for corrosion protection of internal structures, melting can be applied more anodized aluminum alloy than the alloys of the present invention, by deposition of metal at the high temperature, or in other ways. Electrochemical corrosion that occurs due to the difference of the corrosion potentials of the alloys A and 2099, may contribute to localized corrosion in the area of the HAZ of alloy. In this case, to improve the corrosion resistance in the area of HAZ can be useful to use more depleted and more enriched alloys with similar composition to the alloys of the present invention, which should have a smaller difference between the corrosion potentials than for two very different alloys, or use of the alloys of the present invention with different vacation.

Example 4

Two ingots were cast as large ingots of industrial size. The ingot had a composition corresponding to the ideas of the present invention. First ingot marked with alloy E, and the second strand are indicated by alloy F. in Addition, the quality control was cast four ingots of alloy 7085 according to the Association of the aluminium industry and six ingots of alloy 7050 according to the Association of the aluminium industry. The composition of the alloys E and F, the control of the ingots of the alloys 7050 and 7085 and the range of compositions for alloys 7085 and 7050 registered Association of the aluminium industry, are given in Table 6.

Table 6
Alloythe EC.% Zn wt.% Cuwt.% Mgwt.% Fewt.% Siwt.% Zr
EEUR 7.572,111,630,040,010,11
Fof 7.642,151,650,030,020,1
7050-part 16,07of 2.212,180,080,050,11
7050-part 26,07of 2.212,180,080,050,11
7050-part 36,002,222,150,080,050,11
7050-part 4 6,04to 2.292,170,070,040,11
7050-part 56,04to 2.292,170,070,040,11
7050-party 66,09of 2.262,200,080,040,11
7085-part 17,471,641,500,050,020,11
7085-part 2of 7.481,681,500,050,010,11
7085-part 37,351,651,500,040,020,12
7085-part 4 7,311,65the 1.440,030,020,12
7085 range AA7,0-8,01,3-2,01,2-1,8max 0,08max 0,060,08-0,15
7050 range AA5,7-6,72,0-2,61,9-2,6max 0,15max 0,120,08-0,15

The ingots were skalierbare and homogenized with a final holding temperature of from about 870 to 910°F. the Ingot with a composition E was subjected to hot rolling into a sheet having a thickness of 3.125 inches, and the ingot with a composition F was subjected to hot rolling into a sheet having a thickness of 4.0 inches. These dimensions are typical for a standard in the aerospace industry sheets used for teleoperating details. Party 1-3 bars of the control alloy was subjected to hot rolling into a sheet having a thickness of about 4 inches. Part 4 of the ingot from the control alloy 7085 were subjected to hot rolling into a sheet having a thickness of about 3 inches. Three of the ingot from the control of the 7050 alloy was subjected to hot th rolling into sheet, having a thickness of about 4 inches. The other three of the ingot from the control of the 7050 alloy was subjected to hot rolling into a sheet having a thickness of about 3 inches. All ingots rolled across in a long transverse direction to less than 15%. All sheets were subjected to heat treatment in solid solution at a temperature of about 880-900°F for about 2-4 hours, hardening under irrigation cooled to ambient temperature and the cold extract of up to about 1.5-3%.

Samples have been taken from sheets of alloys E and F. These samples were aged very early because before vacation type T74 (suitable teleoperating components)using a conventional three-stage mode. Three-stage regime consisted of a first phase lasting about 6 hours at 250°F, the second stage about 15-20 hours at a temperature of 310°F and the third stage of about 24 hours at 250°F. Some samples of alloys E and F aged very early because 15 hours during the second stage (sample 1). Other samples of alloy F aged very early because 18 hours in the second stage (model 2). Other samples of alloy E aged very early because 20 hours in the second stage (model 2). 4-inch party control alloy 7085 also aged very early because to leave T74, using this conventional 3-stage mode of aging. Sample 1 of part 4 (3-inch sheet) parties control alloy 7085 aged very early because to leave T76, using a conventional three-stage process, and sample 2 party 4 (3-inch sheet) party control alloy 7085 aged very early because to the of Tuska T74, using a conventional three-stage mode. Party control of alloy 7050 aged very early because to leave T74, using conventional two-stage mode of aging.

Measured mechanical properties of tensile and fracture toughness under plane strain (Klcsamples of alloys E and F and parties control alloy 7085 and 7050. Tensile test was conducted in accordance with ASTM E8 and ASTM B557. Test fracture toughness under plane strain (Klc) was conducted in accordance with ASTM E399. Samples for testing fracture toughness under plane strain for alloy E were thick and 2 inches and had a width W of 4 inches in orientation T-L, and had 1 inch in thickness and width W 2 inch orientation S-L. Samples for testing fracture toughness under plane strain for alloy F were 1 inch thick and had a width W of 2 inches and T-L and S-L orientations. Samples for testing the fracture toughness of alloys E and F were centered on the mid-thickness (T/2) of the sheet. Samples for testing fracture toughness under plane strain 4-duimovich sheets from the control alloy 7085 were 2 inches thick and had a width W of 4 inches in orientation T-L, and had a thickness of 1.5 inches and a width W of 3 inches in orientation S-L. Samples for testing fracture toughness under plane strain and 3-inch sheets of the control alloy 7085 were 1.75 inches and had a width W of 5 inches in orientation T-L, and had a thickness of 1.25 inches and a width W of 2.5 inches in orientation S-L. Samples for testing fracture toughness for 4-inch sheets of the control alloy 7085 were centered at quarter thickness (T/4) of the sheet in the orientation T-L and at mid-thickness (T/2) of the sheet in the orientation of S-L. Samples for testing fracture toughness for 3-inch sheets of the control alloy 7085 were centered on the mid-thickness (T/2) of the sheet and T-L and S-L orientation. Samples for testing fracture toughness under plane deformation in orientation T-L for leaves of control 7050 alloy had a thickness of 2 inches and had a width W of 4 inches. Samples for testing fracture toughness under plane deformation in the orientation of the S-L to 3-inch thick sheets of the control alloy 7050 were 1 inch thick and had a width W of 2 inches. Samples for testing fracture toughness under plane deformation in the orientation of the S-L to 4-inch thick sheets of the control 7050 alloy had a thickness of 1.5 inches and had a width W of 3 inches. Samples for testing the fracture toughness of leaves from control alloy 7050 were centered on the mid-thickness (T/2) of the sheet and T-L and S-L orientation. Test pairs using the method of EXCO was conducted for alloy F in accordance with ASTM G34, and the samples for testing were taken from mid-thickness (T/2), one quarter of the thickness (T/4)and one tenth of the thickness (T/10).

The measured properties of alloys E and F and parties control alloy 7085 and 7050 shown in Table 7. When the thickness of the sheet about 3 inches alloy E shows the increase in yield strength tensile approximately (9-12) ksi and higher ultimate tensile strength of about 6-8 ksi compared to the batch control alloy 7050 in the direction of LT. Similarly, the alloy E shows the increase in yield strength tensile approximately 8-10 ksi and higher ultimate tensile strength of about 6-8 ksi compared to the batch control alloy 7050 in the direction of ST. When the thickness of the sheet 4 inch alloy F shows the increase in yield strength tensile approximately 7-9 ksi and higher ultimate tensile strength of about 3-4 ksi compared to the batch control alloy 7050 in the direction of LT. Similarly, the alloy F shows the increase in yield strength tensile approximately 5-7 ksi and higher ultimate tensile strength of about 4-5 ksi compared to the batch control alloy 7050 in the direction of ST. Alloy F shows the improvement in strength of about 2-5 ksi for yield strength and ultimate tensile strength in the LT and ST directions in comparison with batch control alloy 7085 when vacation T74. This improvement in strength aircraft manufacturers is a significant improvement in strength.

Table 7
AlloyParty/
No. sample
Thickness (inches)DirectionTYS (ksi)UTS (ksi)Elongation (%)OrientationKlc(ksi√inch)
7050-T7451party 13LT66,276,611,4T-L28,2
ST62,073,46,2S-L28,0
party 23LT65,876,211,4T-L29,2
ST61,773,36,7S-L 28,1
game 33LT65,375,311,0T-L30,0
ST61,072,56,6S-L28,8
party 44LT65,275,811,3T-L26,3
ST62,974,65,8S-L22,4
part 54LT65,676,110,8T-L26,4
ST62,473,55,6S-L 26,6
party 64LT66,9of 76.87,9T-L26,5
ST61,673,05,1S-L26,2
7085-T7451party 14LT69,176,410,5T-L29,1
ST64,474,17,5S-L32,3
party 24LT69,976,510,7T-L29,4
ST64,774,37,0 S-L31,0
game 34LT69,576,910,6T-L30,1
ST65,474,86,2S-L32,1
7085-T7X51part 4, sample 23LT69,375,418,2T-L35,4
3ST66,575,013,5S-L39,6
part 4, sample 13LT68,674,519,0T-L37,2
3ST 65,574,113,9S-L37,7
Alloy ESample 13,125LT77,6is 83.89,3T-L25,0
ST71,580,97,8S-L27,6
Sample 23,125LT74,782,0the 9.7T-L26,9
ST69,779,48,6S-L29,4
Alloy FSample 14LT74,580,310,0 T-L26,4
ST69,278,27,8S-L25,1
Sample 24LT73,079,610,0T-L28,3
ST67,377,68,6S-L27,4

Properties of alloy E and various conventional alloys, having a thickness of about 3 inches, illustrated in Fig.7. More specifically, Fig.7 compares the fracture toughness under plane strain (Klcin orientation T-L as a function of the yield stress in tension in the direction of LT (long transverse direction) for alloy E (thickness of 3.125 inches), parties control 7050 alloy (having a thickness of about 3 inches) and a 3-inch party control alloy 7085. Alloy E implements a significantly higher yield strength at elongation at fracture toughness, which is close to the viscosity batch control alloy 7050. Alloy E also implements the relationship between strongly the TEW and fracture toughness, comparable to alloy 7085, but, as described below, alloy 7085 not able to consistently withstand SCC-test in the conditions of the sea. In other words, alloy E achieves equal or better resistance to cracking under stress than having a similar size and get a close way alloy 7085, but at higher LT strength. Thus, the alloy E implements not achievable before the combination of LT strength, T-L fracture toughness and corrosion resistance in the specified range of thickness.

7 also contains a darkened area, illustrating the potential properties of sheet products of the alloy of the present invention. The shaded region is bounded by the minimum fracture toughness of 22 ksi√inch, a minimum strength of 72 ksi and a line C-C, which corresponds to the equation FT_TL = -1,0*(TYS_LT)+98 where TYS_LT LT is the yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, and where FT_TL have T-L fracture toughness under plane deformation of the sheet in ksi√inch, measured in accordance with ASTM E399. The shaded area of figure 7 is particularly suitable for sheet products from alloy, having a thickness of from about 2.0 to 2.5 inches to about 3,0, 3,125 or 3.25 inches, and for the holidays T73, T74, T76 or T79.

Properties of alloy F and various conventional alloys, having a thickness of 4 inches, illustrated in Fig. Over oncrete, Fig compares the fracture toughness under plane strain (Klcin orientation T-L as a function of the yield stress in tension in the direction of LT (long transverse direction) for alloy F (4 " thick), parties control 7050 alloy (having a thickness of about 4 inch and 3-inch party control alloy 7085. Alloy F implements a significantly higher yield strength tensile to fracture toughness, which is close to the viscosity batch control alloy 7050. Alloy F also implements the correlation between strength and fracture toughness, which is close to the ratio for parties control alloy 7085, but, as described below, alloy 7085 not able to consistently withstand SCC-test in the conditions of the sea. In other words, the alloy F achieves equal or better resistance to cracking under stress than having a similar size and get a close way alloy 7085, but at higher LT strength. Thus, the alloy implements F is not reachable before the combination of LT strength, T-L fracture toughness and corrosion resistance in the specified range of thickness.

Fig also contains a darkened area, illustrating the potential properties of sheet products of the alloy of the present invention. The shaded region is bounded by the minimum fracture toughness of 21 ksi√inch, minimum, robust is the capacity of 71 ksi and a line D-D, which corresponds to the equation FT_TL = -1,0*(TYS_LT)+98 where TYS_LT LT is the yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_TL have T-L fracture toughness under plane deformation of the sheet in ksi√inch, measured in accordance with ASTM E399. The shaded area on Fig especially suitable for sheet products from alloy, having a thickness of from about 3.0 to 3,125 inches or 3.25 inches to about 3.5, 3.75, or 4 inches, and when you leave T73, T74, T76 or T79.

Properties of alloy E and various conventional alloys, having a thickness of about 3 inches, also illustrated in Fig.9. More specifically, Figure 9 compares the fracture toughness under plane strain (Klcin orientation S-L as a function of the yield stress in tension in the direction of the ST (short transverse direction) for alloy E (thickness of 3.125 inches), parties control 7050 alloy (having a thickness of about 3 inches) and a 3-inch party control alloy 7085. Alloy E implements a significantly higher yield strength at elongation at fracture toughness, which is close to the viscosity batch control alloy 7050. Alloy E also implements the correlation between strength and fracture toughness, which is close to the ratio for party control alloy 7085, but, as described below, alloy 7085 not able to consistently withstand SCC-test in the conditions of the sea. D. the ugogo words, alloy E achieves equal or better resistance to cracking under stress than having a similar size and get a close way alloy 7085, but at higher ST strength. Thus, the alloy E implements not achievable before the combination of ST strength, S-L fracture toughness and corrosion resistance in the specified range of thickness.

Figure 9 also contains a darkened area, illustrating the potential properties of sheet products of the alloy of the present invention. The shaded region is bounded by the minimum fracture toughness of 22 ksi√inch, a minimum strength of 69 ksi and a line E-E, which corresponds to the equation FT_SL = -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L fracture toughness under plane deformation of the sheet in ksi√inch, measured in accordance with ASTM E399. The shaded area in Figure 9 is particularly suitable for sheet products from alloy, having a thickness of from about 2.0 to 2.5 inches to about 3,0, 3,125 or 3.25 inches, and for the holidays T73, T74, T76 or T79.

Properties of alloy F and various conventional alloys, having a thickness of about 4 inches, also illustrated in Figure 10. More specifically, Figure 10 compares the fracture toughness under plane strain (Klcin orientation S-L as a function of the yield strength at restiani is in the direction of the ST (short transverse direction) for alloy F (thickness of 4.0 inches), parties control 7050 alloy (having a thickness of about 4 inches) and batch control alloy 7085 (having a thickness of about 4 inches). Alloy F implements a significantly greater yield strength at elongation at fracture toughness, which is close to the viscosity batch control alloy 7050. Alloy F also implements the correlation between strength and fracture toughness, which is close to the ratio for parties control alloy 7085, but, as described below, alloy 7085 not able to consistently withstand SCC-test in the conditions of the sea. In other words, the alloy F achieves equal or better resistance to cracking under stress than having a similar size and get a close way alloy 7085, but at higher ST strength. Thus, the alloy implements F is not reachable before the combination of ST strength, S-L fracture toughness and corrosion resistance for a given thickness.

Figure 10 also contains a darkened area, illustrating the potential properties of sheet products of the alloy of the present invention. The shaded region is bounded by the minimum fracture toughness of 20 ksi√inch, a minimum strength of 66 ksi and a line F-F, which corresponds to the equation FT_SL = -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L wascott the fracture under plane deformation of the sheet, in ksi√inch, measured in accordance with ASTM E399. The shaded area in Figure 10 is particularly suitable for sheet products from alloy, having a thickness of from about 2.0 to 2.5 inches to about 3,0, 3,125 or 3.25 inches, and for the holidays T73, T74, T76 or T79.

Alloy F in both the aging was tested for resistance to exfoliation corrosion (EXCO) in accordance with ASTM G34. Alloy F in both the aging was achieved assessment EA, consistent with good resistance to exfoliation corrosion for alloy 7XXX. Close results should be expected if EXCO-test to expose the alloy E. Thus, the alloys of the present invention have the improved strength while maintaining excellent characteristics of resistance to delamination, and all samples of alloy F reach assessment EA by EXCO.

Alloys E, F and alloys 7085 were subjected to two types of tests to stress corrosion cracking under tension. The first test is an accelerated test for stress corrosion cracking stress (SCC) in terms of alternate immersion (AI), was performed for samples 1 and 2 alloys E and F, as well as on the leaves of the control alloy 7085 samples for testing are taken from mid-thickness (T/2) in the direction of ST, in accordance with ASTM G44, G47 and/or G49. The results of SCC tests with AI are shown in Table 8 (for a thickness of 4 inches) and Table 9 (for a thickness of 3 inches).

td align="center"> 35
Table 8
AlloyTYS (LT)Party/No. sampleThickness (inches)Stress (ksi)The number of trialsDays testDamage (days)
Alloy F74,5 ksiSample 1440310053,71,78
50310051,57,58
73,0 ksiSample 2440310074,93, 100
50310063,63,63
7085-T745169,1 ksiParty 14565no
4556544, 50 and 58
to 69.9 ksiParty 243556560, 60 and 62
4556545, 46, 57, 57, 57
for 69.5 ksiGame 3435565no
4556550,57

Table 9
AlloyTYS (LT)Party/
No. sample
Thickness (inches)Stress (ksi)The number of trials Days testDamage (days)
Alloy E77,6 ksiSample 13,12540310054, 54, 99
50310047, 52, 74
74,7 ksiSample 23,12540310068, 74, 76
50310050, 54, 57
7085-T745168,6 ksipart 4, sample 2335510072, 73, 75
45510057,61,61,64, 65
7085-T7651 69,4 ksipart 4, sample 1335510061,65,68, 81
45510048, 65, 65

Alloys E and F gave acceptable characteristics for voltage levels 40 and 50 ksi, which are respectively 5 and 15 ksi above the minimum requirements for the classification of alloy as having vacation T74.

SCC-tests in the conditions of the sea coast was also conducted for samples 1 and 2 alloy E samples for testing are taken from mid-thickness (T/2) in the direction of ST. Were also conducted SCC-tests under the sea coast for alloy 7085. Samples for SCC tests in the conditions of the sea coast was tested in the clamping device at a constant deformation (e.g., similar to used in accelerated laboratory SCC-tests). SCC-test in terms of the sea coast includes a continuous effect on the samples through the racks of the environment of the sea coast, and the sample set of about 1.5 m from the ground and oriented at 45° to the horizontal, and the sample surface is open to the prevailing winds. The samples are located about 100 metres from the shoreline. In od the om embodiment, the coastline is rocky nature, moreover, the prevailing wind is focused on the sample to provide aggressive conditions of salt fog (for example, similar conditions to the station for testing for atmospheric corrosion in the conditions of the sea, Pt. Judith, Rhode Island, USA, Alcoa Inc.). The results of SCC tests in the conditions of the sea coast for alloy E and alloy 7085 shown in Table 10.

td align="center"> 40
Table 10
AlloyTYS (LT)Party/No. sampleThickness (inches)Voltage-tion (ksi)The number of trialsDays testDamage (days)
Alloy E77,6 ksiSample 13,125403262no
50326260, 102
74,7 ksiSample 23,1253262no
503262no
708568,6 ksipart 4, sample 23355525no
708569,4 ksipart 4, sample 1335552576, 132 and 3 without damage

Many samples of alloy E tested (sample fails the test, if it is divided into two pieces or crack becomes visible to the naked eye) when the level of stress of 40 ksi and 50 ksi after 262 days of exposure. Remember that the alloy E was reached LT strength 74,7 ksi and 77.6 ksi for samples 1 and 2, respectively. On the contrary, the alloy 7085 similar thickness, with LT strength just 68,6, and 69.4 ksi, did not pass the test of the null of 5 and 2 out of 5 times, respectively. Note, the trend data on alloy 7085 that only with a slightly higher strength alloy 7085 capability in order to sustain SCC-test conditions the coast falls. It is expected that if the alloy 7085 process to reach level LT strength 72 ksi at a thickness of 3 inches, this alloy 7085 will invariably fail in SCC-test in the conditions of the sea coast (at a voltage of 35 ksi towards ST), whereas the alloy E (and other alloys, as defined by the present invention) will continue to sustain SCC-tests under the sea coast at the same level of strength and SCC-stress.

Thus, the alloys of the present invention is able to achieve not feasible before the combination of strength, fracture toughness and corrosion resistance in the specified ranges of thickness. In one embodiment, given by the product of the aluminum alloy vacation T74. The product of the aluminum alloy can be made from the first sheet, second sheet and/or the third sheet. If you use the first sheet, the first sheet will have a thickness of not more than about 2,00 inches and contain the alloy composition according to any one of embodiments 1, 2, 3, 4 or 5 with Fig. 2C-1, 2C-2, 2D-1 and 2D-2 or on the modalities for the implementation of 1 or 2 with Fig. 2E and 2F, which are described above. If you use the second sheet, the second sheet will have a thickness greater than a 2.00 inch, but not more than about 3.00 inches, and contain the alloy composition according to any one of embodiments 1, 2, 3, 4 or 5 with Fig. 2C-1, 2C-2, 2D-1 and 2D-2 or on the modalities for the implementation of 1 or 2 with Five and 2F, what is described above. If you are using a third sheet, the third sheet will have a thickness greater than about 3.00 inches, but not more of 4.00 inches, and contain the alloy composition according to any one of embodiments 1, 2, 3, 4 or 5 with Figs-1, 2C-2, 2D-1 and 2D-2 or on the modalities for the implementation of 1 or 2 with Five and 2F, which are described above. The product of the aluminum alloy may contain other compounds, for example, other of the above composition ranges. In addition, in any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities.

In this embodiment, any of the first sheet may have a ratio between strength and fracture toughness, which satisfies the expression

FT ≥ -2,3*(TYS)+229 where TYS has a longitudinal yield strength tensile of the first sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT is an L-T fracture toughness under plane deformation of the first sheet in ksi√inch, measured in accordance with ASTM E399, and the first sheet has a TYS of at least 74 ksi, and the first sheet has an FT of at least 36 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 75 ksi, for example, at least about 76 ksi, or IU the greater extent of about 77 ksi, or at least about 78 ksi, or at least about 79 ksi, or even at least about 80 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 40 ksi√inch, for example, at least about 42 ksi√inch, or at least about 44 ksi√inch, or at least about 46 ksi√inch, or at least about 48 ksi√inch, or even at least about 50 ksi√inch.

In this embodiment, any of the second sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_TL ≥ -1,0*(TYS_LT)+98 where TYS_LT have LT yield strength tensile second sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_TL have T-L fracture toughness under plane deformation of the second sheet, in ksi√inch, measured in accordance with ASTM E399, and the second sheet has TYS_LT at least 72 ksi, and the second sheet has FT_TL at least 24,5 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 73 ksi, for example, at least about 74 ksi, or at least about 75 ksi, or at least about 76 ksi, or even at least about 77 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 25 ksi√inch, for example, at m the re about 26 ksi√inch, or at least about 27 ksi√inch, or even at least about 28 ksi√inch.

In this embodiment, any of the second sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_SL ≥ -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength tensile second sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L fracture toughness under plane deformation of the second sheet, in ksi√inch, measured in accordance with ASTM E399, and the second sheet has TYS_ST at least 69 ksi, and the second sheet has FT_SL at least 25 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 69,5 ksi, for example, at least about 70 ksi, or at least approximately 70.5 ksi, or even at least about 71 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 26 ksi√inch, for example, at least about 27 ksi√inch, or at least about 28 ksi√inch, or at least about 29 ksi√inch, or at least about 30 ksi√inch, or even at least about 31 ksi√inch.

In this embodiment, any third sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_TL ≥ -1,0*(TYS_LT)+98 where TYS_LT have LT Pres who ate of the yield strength tensile third sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_TL have T-L fracture toughness under plane deformation, respectively, of the third sheet, in ksi√inch, measured in accordance with ASTM E399, and the third sheet has TYS_LT at least 71 ksi, and the third sheet has FT_TL at least 23 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 71.5 ksi, for example, at least about 72 ksi, or at least about to 72.5 ksi, or at least about 73 ksi, or at least about 73,5 ksi, or even at least about 74 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 24 ksi√inch, for example, at least about 25 ksi√inch, or at least about 26 ksi√inch, or at least about 27 ksi√inch, or at least about 28 ksi√inch, or even at least about 29 ksi√inch.

In this embodiment, any third sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_SL ≥ -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength tensile third sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L fracture toughness under plane deformation of the third sheet, in ksi√inch, measured in accordance with ASTM E399, moreover, the third sheet has TYS_ST at least 66 ksi, and with the third sheet has FT_SL at least 23 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 66.5 ksi, for example, at least about 67 ksi, or at least about to 67.5 ksi, or at least about 68 ksi, or at least about 68,5 ksi, or even at least about 69 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 24 ksi√inch, for example, at least about 25 ksi√inch, or at least about 26 ksi√inch, or at least about 27 ksi√inch, or even at least about 28 ksi√inch.

In this embodiment, any of the first, second or third sheet may invariably be able to withstand one or more tests on stress corrosion cracking under tension. In the private embodiment, and by definition leave T74, leaves invariably stand (described below) test for stress corrosion cracking stress (SCC) in the conditions of the sea coast, at a voltage of at least 35 ksi in the direction of ST, or at least about 40 ksi in the direction of ST, or even at least about 45 ksi in the direction of ST, and for at least 180 days. In some embodiments of the leaves consistently withstand SCC-tests under the sea coast for less is th least 230 days or at least 280 days, or at least 330 days, or even at least 365 days, at the specified level(s) voltage. In the private embodiment, the sheets always stand SCC-test with alternative immersion (in accordance with ASTM G44, G47 and/or G49) for at least 30 days. In some embodiments of the leaves consistently withstand SCC-test with alternative immersion for at least 40 days, or at least 60 days, or at least 80 days, or even at least 100 days. It is known that none of the traditional alloys 7XXX series with vacation T74 not able to achieve all of (i) provide higher strength in the specified range of thickness, (ii) provide higher fracture toughness in the specified range of thickness, (iii) provide higher correlation between strength and fracture toughness in the specified range of thickness, and (iv) the ability to consistently withstand one or both of the above SCC tests in the specified range of thickness.

In another embodiment, given by the product of the aluminum alloy vacation T76. The product of the aluminum alloy can be made from the first sheet, second sheet and/or the third sheet. If you use the first sheet, the first sheet will have a thickness of not more than about 2,00 inches, and contain the alloy composition is about any one of embodiments 1, 2, 3, 4 or 5 on Figs-1, 2C-2, 2D-1 and 2D-2, or on the modalities for the implementation of 1 or 2 with Five and 2F, which are described above. If you use the second sheet, the second sheet will have a thickness greater than a 2.00 inch, but not more than about 3.00 inches, and contain the alloy composition according to any one of embodiments 1, 2, 3, 4 or 5 with Figs-1, 2C-2, 2D-1 and 2D-2 or on the modalities for the implementation of 1 or 2 with Five and 2F, which are described above. If you are using a third sheet, the third sheet will have a thickness greater than about 3.00 inches, but not more of 4.00 inches and contain the alloy composition according to any one of embodiments 1, 2, 3, 4 or 5 with Fig. 2C-1, 2C-2, 2D-1 and 2D-2 or on the modalities for the implementation of 1 or 2 with Fig. 2E and 2F, which are described above. The product of the aluminum alloy may contain other compounds, for example, other of the above ranges of compositions. In addition, in any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities.

In this embodiment, any of the first sheet may have a ratio between strength and fracture toughness, which satisfies the expression

FT ≥ -2,3*(TYS)+229 where TYS has a longitudinal yield strength tensile of the first sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT is an L-T fracture toughness when loscoe deformation of the first sheet, in ksi√inch, measured in accordance with ASTM E399, and the first sheet has TYS at least 79 ksi, and the first sheet has an FT of at least 30 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 80 ksi, for example, at least about 81 ksi, or at least about 82 ksi, or at least about 83 ksi, or at least about 84 ksi, or at least about 85 ksi, or even at least about 86 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 32 ksi√inch, for example, at least about 34 ksi√inch, or at least about 36 ksi√inch, or at least about 38 ksi√inch, or at least about 40 ksi√inch, or even at least about 42 ksi√inch.

In this embodiment, any of the second sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_TL ≥ -1,0*(TYS_LT)+98 where TYS_LT have LT yield strength tensile second sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_TL have T-L fracture toughness under plane deformation of the second sheet ksi√inch, measured in accordance with ASTM E399, and the second sheet has TYS_LT at least 76 ksi, and the second sheet has FT_TL at least 22 ksi√inch. In some of these variant the implementation of the sheet may have a yield strength tensile of at least about 77 ksi, for example, at least about 78 ksi, or at least about 79 ksi, or at least about 80 ksi, or even at least about 81 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 22.5 ksi√inch, for example, at least about 23 ksi√inch, or at least about 23.5 ksi√inch, or at least about 24 ksi√inch, or at least about 24.5 ksi√inch, or even at least about 25 ksi√inch.

In this embodiment, any of the second sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_SL ≥ -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength tensile second sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L fracture toughness under plane deformation of the second sheet, in ksi√inch, measured in accordance with ASTM E399, and the second sheet has TYS_ST at least 71 ksi, and the second sheet has FT_SL at least 22 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 71.5 ksi, for example, at least about 72 ksi, or at least about to 72.5 ksi, or even at least about 73 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 23 ksi√inch, for example, m is Nisha least about 24 ksi√inch, or at least about 25 ksi√inch, or at least about 26 ksi√inch, or at least about 27 ksi√inch, or even at least about 28 ksi√inch.

In this embodiment, any third sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_TL ≥ -1,0*(TYS_LT)+98 where TYS_LT have LT yield strength tensile third sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_TL have T-L fracture toughness under plane deformation, respectively, of the third sheet, in ksi√inch, measured in accordance with ASTM E399, and the third sheet has TYS_LT at least 75 ksi, and a third sheet is FT in TL at least 21 ksi√inch. In some of these embodiments, the sheet may have a yield strength tensile of at least about 75,5 ksi, for example, at least about 76 ksi, or at least about 76,5 ksi, or at least about 77 ksi, or at least about 77,5 ksi, or even at least about 78 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 22 ksi√inch, for example, at least about 23 ksi√inch, or at least about 24 ksi√inch, or at least about 25 ksi√inch, or at least about 26 ksi√inch, or even at least about 27 ksi√inch.

In this embodiment, the status is of any third sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT_SL ≥ -1,1*(TYS_ST)+99 where TYS_ST is ST yield strength tensile third sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT_SL have S-L fracture toughness under plane deformation of the third sheet, in ksi√inch, measured in accordance with ASTM E399, and the third sheet has TYS_ST at least 70 ksi, and the third sheet has FT_SL at least 20 ksi√inch. In some of these embodiments, the sheet may have a yield strength in tension of at least approximately 70.5 ksi, for example, at least about 71 ksi, or at least about 71.5 ksi, or at least about 72 ksi, or at least about to 72.5 ksi, or even at least about 73 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 21 ksi√inch, for example, at least about 22 ksi√inch, or at least about 23 ksi√inch, or at least about 24 ksi√inch, or even at least about 25 ksi√inch.

In this embodiment, any of the first, second or third sheets consistently maintains one or more tests on stress corrosion cracking under tension. In the private embodiment, and by definition leave T76 leaves invariably stand (described below) test for stress corrosion cracking stress (SCC) in terms of the s sea coast in the voltage range of at least about 25 ksi (for example, the range from 25 to 34 ksi ksi) in the direction of ST and for at least 180 days. In some embodiments of the leaves consistently withstand SCC-test in terms of the sea-coast for at least 230 days, or at least 280 days, or at least 330 days, or even at least 365 days, at the specified level(s) voltage. In the private embodiment, the sheets always stand SCC-test by alternate immersion (in accordance with ASTM G44, G47 and/or G49) for at least 30 days. In some embodiments of the leaves consistently withstand SCC-test with alternative immersion for at least 40 days, or at least 60 days, or at least 80 days, or even at least 100 days. It is known that none of the traditional alloys 7XXX series with vacation T76 not able to achieve all of (i) provide higher strength in the specified range of thickness, (ii) provide higher fracture toughness in the specified range of thickness, (iii) provide higher correlation between strength and fracture toughness in the specified range of thickness, and (iv) the ability to consistently withstand one or both of the above SCC tests in the specified range of thickness.

In one embodiment, the aluminum alloy is used as the upper clapboard the wing and Aero-space plane. The upper wing skin may be made from sheet aluminum alloy, having a thickness of not more than about 2,00 inches, with aluminum alloy has any of the compositions according to the options exercise 1, 2, 3, 4 or 5 with Fig. 2C-1, 2C-2, 2D-1 and 2D-2. The product of the aluminum alloy can (rarely) contain other compounds, for example, any of the other above-mentioned ranges of composition. In any of these embodiments, the aluminum alloy may be composed mainly of the above components (in addition to aluminum), and the rest are aluminum and incidental elements and impurities. In these embodiments, the implementation of a sheet of aluminum alloy may have a ratio between strength and fracture toughness, which satisfies the expression FT ≥ -4,0*(TYS)+453 where TYS has a longitudinal yield strength in tension of the sheet, in ksi, measured in accordance with ASTM E8 and ASTM B557, where FT is an L-T fracture toughness (Kapp) sheet in plane stress, ksi√inch, and FT measured in accordance with ASTM E561 and B646 on the sample of aluminum alloy with a Central crack, taken from the position of T/2 sheets of aluminum alloy, and the sample has a width 16 inch, thickness 0.25 inch, and the initial length of the pre-fatigue crack 4 inches. In some of these embodiments, the sheet may have a limit recuces and at least about 80 ksi, for example, at least about 81 ksi, or at least about 82 ksi, or at least about 83 ksi, or at least about 84 ksi, or even at least about 85 ksi. In some of these embodiments, the sheet may have a fracture toughness of at least about 100 ksi√inch, for example, at least about 101 ksi√inch, or at least about 102 ksi√inch, or at least about 103 ksi√inch, or at least about 104 ksi√inch, or even at least about 105 ksi√inch. In addition to improved yield strength tensile and fracture toughness in a plane stress state list for the top covering of the wing can also achieve improved fracture toughness under plane strain (Klc). Thus, in these embodiments, the implementation of the sheet may have a ratio between strength and fracture toughness, which satisfies the expression FT-KlC ≥ -2,3*(TYS)+229 where TYS has a longitudinal yield strength under tension, as described above, and where FT-KlC has an L-T fracture toughness under plane deformation of the sheet in ksi√inch, measured in accordance with ASTM E399, and the sheet is FT-KlC at least 34 ksi√inch. In some of these embodiments, the sheet may have a fracture toughness of FT-KlC at least about 36 ksi√inch, for example, at least about 38 ksi√inch, or at least about 40 ksi√guy is a, or even at least about 42 ksi√inch. It is known that none of the traditional alloys 7XXX series are not able to achieve all of (i) provide higher strength in the specified range of thickness, (ii) provide higher fracture toughness in the specified range of thickness, and (iii) provide higher correlation between strength and fracture toughness in the specified range of thickness. These alloys may also be able to achieve corrosion resistance, installed above in Example 2.

Although much of this description has been provided in respect of the sheets of the alloy, it is expected that similar improvements will be implemented with the alloy of the present invention and in the products of another shape, such as stamping and forging. In addition, although particular embodiments of the present invention have been described in detail, the experts in this field should understand that in the light of the General ideas of the invention can manifest various modifications and alternatives to those details. Accordingly, it is understood that the disclosed private configurations are merely illustrative, and not restrictive of the scope of the present invention, the full width of which shall be applied the formula and all its equivalents.

1. A sheet of an aluminium alloy consisting of between 6.8 and 8.5 wt.% Zn, a 1.75-2.3 wt.% C, 1,5-,84 wt.% Mg and up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and optionally additives, grinding grain, and the rest are aluminum and inevitable impurities, and the sheet has a thickness of not more than 2,00 inches, and the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT_LT≥-4,0*(TYS_L)+453, and the sheet has TYS_L at least 80 ksi and FT_LT at least 100 ksi√inch, where TYS_L - yield strength in tension of the sheet in the direction L, in ksi, measured in accordance with ASTM E8 and ASTM V, FT_LT - fracture toughness (Karr) sheet in plane stress state in the direction L-T, ksi√inch, measured in accordance with ASTM E561 and V on the sample of aluminum alloy with a Central crack in a position of T/2 sheets, the sample has a width 16 inch, thickness 0.25 inch, and the initial length of the pre-fatigue crack 4 inches.

2. The sheet according to claim 1, in which the amount of Mg in the aluminum alloy is of 1.55 to 1.75 wt.%.

3. The sheet according to claim 2, in which the amount of si in the aluminum alloy is from 1.95 to 2.25 wt.%.

4. The sheet according to claim 3, in which the amount of Zn in the aluminum alloy is from 7.0 to 8.5 wt.%.

5. A sheet of an aluminium alloy consisting of between 6.8 and 8.5 wt.% Zn, a 1.75-2.3 wt.% C, 1,5-1,84% wt. Mg and up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and optionally additives, grinding grain, and the rest are al mini and inevitable impurities, when this sheet has a thickness of not more than 2,00 inches, and the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT-K1C≥-2,3*(TYS_L)+229, when this sheet has TYS_L at least 74 ksi and FT-K1C of at least 30 ksi√inch, where TYS_L - yield strength in tension of the sheet in the direction L, in ksi, measured in accordance with ASTM E8 and ASTM B557, FT-K1C is the fracture toughness under plane deformation of the sheet in the direction L-T in ksi√inch, measured in accordance with ASTM E399, and fracture toughness of the samples corresponds to the full thickness of the sheet.

6. The sheet according to claim 5, in which the amount of Mg in the aluminum alloy is of 1.55 to 1.75 wt.%.

7. The sheet according to claim 6, in which the amount of si in the aluminum alloy is from 1.95 to 2.25 wt.%.

8. The sheet according to claim 7, in which the amount of Zn in the aluminum alloy is from 7.0 to 8.5 wt.%.

9. A sheet of an aluminium alloy consisting of between 6.8 and 8.5 wt.% Zn, a 1.75-2.3 wt.% C, 1,5-1,84% wt. Mg and up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and optionally additives, grinding grain, and the rest are aluminum and inevitable impurities, and the sheet has a thickness of from 2.00 to 3.25 inches, and the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT_TL≥-1,0*(TYS_LT)+98, and the sheet has TYS_LT at least 72 ksi and FT_TL at least 22 ksi√inch, the de TYS_LT - the yield strength in tension of the sheet in the direction of LT, in ksi, measured in accordance with ASTM E8 and ASTM B557, and FT_TL - fracture toughness under plane deformation of the sheet in the direction T-L, ksi√inch, measured in accordance with ASTM E399 in the position of T/2 sheets.

10. The sheet according to claim 9, in which the amount of Mg in the aluminum alloy is of 1.55 to 1.75 wt.%.

11. Sheet of claim 10, in which the amount of si in the aluminum alloy is from 1.9 to 2.3 wt.%.

12. The sheet according to claim 11, in which the amount of Zn in the aluminum alloy is from 7.0 to 8.5 wt.%.

13. The sheet according to any one of p-12, in which FT_TL is at least of 24.5 ksi√inch.

14. Sheet according to item 13, in which TYS_LT is at least 76 ksi.

15. The sheet according to any one of p-12, in which the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT_SL≥-1,1*(TYS_ST)+99, and the sheet has TYS_ST at least 69 ksi and FT_SL at least 22 ksi√inch, where TYS_ST - yield strength in tension of the sheet in the direction of ST, in ksi, measured in accordance with ASTM E8 and ASTM B557, and FT_SL - fracture toughness under plane deformation of the sheet in the direction of S-L, in ksi√inch, measured in accordance with ASTM E399 in the position of T/2 sheets.

16. The sheet 15, which FT_SL is at least 25 ksi√inch.

17. Sheet according to clause 16, in which TYS_ST is at least 71 ksi.

18. The sheet on which the yubom of PP-12, in which the sheet withstands the test on resistance to stress corrosion cracking under the strain when AC immersed when the voltage level of 25 ksi, measured in accordance with ASTM G44, G47 and G49 on the sample for testing in the mid-thickness (T/2) in the direction of ST, for at least 30 days.

19. Sheet p, in which the sheet withstands the test on resistance to stress corrosion cracking under stress, in terms of the sea coast when the voltage level measured on the sample for testing in the mid-thickness (T/2) in the direction of ST, for at least 180 days.

20. The sheet according to claim 19, in which the voltage level is 35 ksi for testing the resistance to stress corrosion cracking under the strain when alternating immersion and/or tests on resistance to stress corrosion cracking under tension in the conditions of the sea.

21. A sheet of an aluminium alloy consisting of between 6.8 and 8.5 wt.% Zn, a 1.75-2.3 wt.% C, 1,5-1,84% wt. Mg and up to 0.25 wt.% at least one of Zr, Hf, Sc, Mn, and V, and optionally additives, grinding grain, and the rest are aluminum and inevitable impurities, and the sheet has a thickness of from 2.75 to 4 inches, and the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT_TL≥-1,0*(TYS_LT)+98, when this sheet has TYS_LT for men is her least 71 ksi and FT_TL at least 21 ksi√inch, where TYS_LT - yield strength in tension of the sheet in the direction of LT, in ksi, measured in accordance with ASTM E8 and ASTM B557, a FT_TL - fracture toughness under plane deformation of the sheet in the direction T-L, ksi√inch, measured in accordance with ASTM E399 in the position of T/2 sheets.

22. Sheet according to item 21, in which the amount of Mg in the aluminum alloy is of 1.55 to 1.75 wt.%.

23. Sheet according to article 22, in which the amount of si in the aluminum alloy is from 1.9 to 2.3 wt.%.

24. Sheet according to item 23, in which the amount of Zn in the aluminum alloy is from 7.0 to 8.5 wt.%.

25. The sheet according to any one of p-24, in which FT_TL is at least 23 ksi√inch.

26. Sheet A.25, in which TYS_LT is at least 74 ksi.

27. The sheet according to any one of p-24, in which the sheet has a ratio of yield strength and fracture toughness, which satisfies the expression: FT_SL≥-1,1*(TYS_ST)+99, and the sheet has TYS_ST at least 66 ksi and FT_SL at least 20 ksi√inch, where TYS_ST - yield strength in tension of the sheet in the direction of ST, in ksi, measured in accordance with ASTM E8 and ASTM B557, and FT_SL - fracture toughness under plane deformation of the sheet in the direction of S-L, in ksi√inch, measured in accordance with ASTM E399 in the position of T/2 sheets.

28. Sheet according to item 27, which FT_SL is at least 23 ksi√inch.

29. Sheet p in which TYS_ST is at least 69 ksi.

0. The sheet according to any one of p-24, in which the sheet withstands the test on resistance to stress corrosion cracking under the strain when AC immersed when the voltage level of 25 ksi, measured in accordance with ASTM G44, G47 and G49 on the sample for testing in the mid-thickness (T/2) in the direction of ST, for at least 30 days.

31. The sheet according to any one of p-24, in which the sheet withstands the test on resistance to stress corrosion cracking under stress, in terms of the sea coast when the voltage level measured on the sample for testing in the mid-thickness (T/2) in the direction of ST, for at least 180 days.

32. Sheet p, in which the voltage level is 35 ksi for testing the resistance to stress corrosion cracking under the strain when alternating immersion and/or tests on resistance to stress corrosion cracking under tension in the conditions of the sea.



 

Same patents:

FIELD: weapons and ammunition.

SUBSTANCE: method consists in obtaining the workpiece to be rolled and its heating, hot rolling of plate according to the size requirements, cooling down to room temperature and artificial ageing. Production of workpieces to be rolled involves rolling according to the size requirements of ingots and/or slabs from aluminium alloys for sandwich plate and assembly of a pack using them. Pack is heated at 500-550°C during 5-7 hours. Rolling according to the size requirements is performed at 410-450°C. Additional hardening is performed at 450-480°C. Artificial ageing is performed at temperature of 110-120°C during 24-36 hours.

EFFECT: improving armour properties and durability of sandwich plate.

5 cl, 1 tbl

FIELD: metallurgy.

SUBSTANCE: aluminium-based alloy contains the following, wt %: zinc - 6.35 - 8.0, magnesium - 0.5 - 2.5, copper - 0.8 -1.3, iron - 0.02 - 0.25, silicon - 0.01 - 0.20, zirconium - 0.07 - 0.20, manganese - 0.001 - 0.1, chrome - 0.001 - 0.05, titanium - 0.01 - 0.10, boron - 0.0002 -0.008, beryllium - 0.0001 - 0.05, at least one element from potassium, sodium, calcium group in quantity of 0.0001 - 0.01 each, aluminium is the rest; at total content of zinc, magnesium, copper within 8.5-11.0, and that of zirconium, manganese and chrome - within 0.1-0.35. Method involves loading and melting of charge components, flux treatment of molten metal, molten metal purification, further vacuum treatment of molten metal in mixer and casting of ingots; boron is added to molten metal in the form of Al-Ti-Be alloy which is distributed at least one hour before molten metal pouring to mixer along the whole surface area of mixer bottom; at that, mixer is pre-heated to temperature which is by 15-30°C more than molten metal temperature, and vacuum treatment of molten metal in mixer is performed at temperature of 695-720°C, during 45-90 minutes.

EFFECT: invention allows obtaining high-strength aluminium alloys with absence of primary intermetallic compounds, decreased content in them of non-metallic inclusions and dissolved gases, with stable properties and optimum size of grain on basis of standard furnace and process equipment.

2 cl, 3 tbl

FIELD: metallurgy.

SUBSTANCE: there proposed is aluminium-based alloy intended for manufacture of deformed semi-finished products in the form of sheets, formings, rods, tubes or in any other form to be used in gas centrifuges, low pressure compressors, vacuum molecular pumps and in other heavily loaded items operating at moderately increased temperatures. Alloy contains the following components, wt %: zinc 6.6-7.4, magnesium 3.2-4.0, copper 0.8-1.4, scandium 0.12-0.30, zirconium 0.06-0.20, titanium 0.01-0.07, molybdenum 0.01-0.07, nickel 0.35-0.65, iron 0.35-0.65, silcone 0.10-0.30, and aluminium is the rest.

EFFECT: improving strength properties of alloy at room temperature.

3 tbl, 1 ex

FIELD: metallurgy.

SUBSTANCE: aluminium-based cast alloy has the following chemical composition, in wt %: Cu 3.5-6.0, Mg 0.2-0.9, Ti 0.1-0.4, Zr 0.1-0.5, Mn 0.2-1.2, Zn 0.5-2.5, Sc 0.15-0.5, Al making the rest.

EFFECT: reduced metal consumption, higher reliability in operation.

2 tbl

FIELD: metallurgy.

SUBSTANCE: invention relates to alloy of AA7000 series and to the manufacturing method of products from this aluminium alloy, and namely to aluminium deformed products of relatively large thickness, namely of 30 to 300 mm. Method involves casting of workpiece - ingot of aluminium alloy of AA7000 series, which contains >0.12 to 0.35% Si, pre-heating and/or homogenisation of workpiece, hot deformation treatment of workpiece using one or more methods chosen from the group, which involves rolling, extrusion and forging, optionally cold deformation treatment, solution treatment, workpiece solution treatment cooling, optional tension or compression or other cold deformation treatment for release of stresses, which is performed by straightening or drawing or by cold rolling, ageing of workpiece in order to achieve the required state. At least one heat treatment is performed at temperature in the range of more than 500°C, but lower than solidus temperature of the considered aluminium alloy. The above heat treatment is performed either: (i) after heat treatment by homogenisation prior to hot deformation treatment, or (ii) after solution treatment, or (iii) both after heat treatment by homogenisation prior to hot deformation treatment, and after solution treatment.

EFFECT: obtaining the product from deformed aluminium alloy, which has improved balance of properties, and namely destruction viscosity, tensile yield point, tensile ultimate strength and relative elongation.

30 cl, 8 tbl, 3 ex

FIELD: metallurgy.

SUBSTANCE: group of inventions can be used at manufacture of semi-finished products in the form of forgings, formings, pressed rods and channels, rolled plates and sheets from high-strength alloys of Al-Zn-Mg-Cu system, which are intended to be used in power structures of aerospace equipment and transport means, on which stringent strength, crack resistance, fatigue life, corrosion resistance requirements are imposed. In order to solve the set task, high-strength alloy on the basis of aluminium, which contains the following, wt %, is proposed: Zn 6.2-8.0, Mg 1.5-2.5, Cu 0.8-1.2, Zr 0.05-0.15, Fe 0.03-0.15, Ti 0.01-0.06, at least one element of the group of metals: Ag 0.01-0.5, Sc 0.01-0.35, Ca 0.0001-0.01, Al and inevitable impurities are the rest. In particular version of alloy the inevitable impurities include not more than 0.05 of Si, Mn, Cr, Ni and not more than 0.01 of Na, H2, O2, B, P. Method for obtaining an item from this alloy involves ingot casting, its homogenisation, hot deformation and strengthening heat treatment of the item, which includes hardening and staged ageing; at that, during the ingot casting there performed is melt purification by blowing with argon or mixture of argon with chlorine and out-of-furnace purification using rotor and/or filtering devices, and homogenisation is performed as per one-stage mode at temperature which is by 55-130°C lower than unbalanced solidus (tu.s.) temperature of this alloy with exposure during 8-36 h or as per two-stage mode at temperature at the first stage at the temperature which is by 175-280°C lower than tu.s. temperature, and at the second stage at the temperature which is by 75-125°C lower than tu.s., with exposure at each stage to 24-36 h; hot deformation is performed at temperature of 300-420°C, hardening is performed at temperature which is by 50-120°C lower than tu.s. during the time determined with the item thickness, with further cooling to temperature of not more than 80°C.

EFFECT: improving the set of mechanical and corrosive characteristics, and characteristics of crack resistance, life time and manufacturability.

10 cl, 1 tbl

FIELD: metallurgy.

SUBSTANCE: invention refers to alloys on base of aluminium, particularly to Al-Zn-Cu-Mg alloys on base of aluminium, and also to procedure of fabrication of rolled or forged deformed product of it and to rolled or forged deformed product proper. The procedure consists in following stages: a) casting an ingot, containing wt % Zn 6.6-7.0, Mg 1.68-1.8, Cu 1.7-2.0, Fe 0-0.13, Si 0-0.10, Ti 0-0.06, Zr 0.06-0.13, Cr 0-0.04, Mn 0-0.04, additives and other side elements ≤0.05 each, b) homogenising of the said ingot at 860-930°F or, preferably, at 875-905°F, c) hot deformation treatment of the said ingot with temperature at input 640-825°F, but preferably - 650-805°F by rolling or forging to a plate with finish thickness from 2 to 10 inches, d) thermal treatment for solid solution and quenching the said plate, e) drawing the said plate with residual deformation from 1 to 4 %, f) ageing the said plate by heating at 230-250°F during from 5 to 12 hours and 300-360°F during from 5 to 30 hours during equivalent time t(eq) between 31 and 56 hours. Equivalent time t(eq) is determined from formula:

where T corresponds to instant temperature in K during annealing, while Tcontr corresponds to control temperature equal to 302°F (423K), and t(eq) is expressed in hours.

EFFECT: production of deformed product possessing improved combination of mechanical strength for corresponding level of crack resistance and resistance to corrosion cracking under load.

8 cl, 2 dwg, 10 tbl, 4 ex

FIELD: metallurgy.

SUBSTANCE: alloy on base of aluminium contains following components wt %: zinc 5-8, magnesium 2-3.1, nickel 1-4.2, iron 0.02-1, zirconium 0.02-0.25 %, copper 0.05-0.3 %. Also, temperature of equilibrium solidus of material is as high as 550°C and hardness is as high as 180 HV. Alloy has a structure corresponding to matrix formed with solid solution of aluminium with uniformly distributed disperse particles of secondary discharges in it and particles of aluminides containing nickel and iron of eutectic origin uniformly distributed in matrix. Also, alloy contains matrix and aluminides at the following ratio, vol % aluminides containing nickel and iron 5.0-6.3, matrix - the rest.

EFFECT: production of new high-strength alloy thermally hardenable and designed both for fabrication of shaped casting and of deformed semi-products.

4 cl, 5 tbl, 4 ex

FIELD: metallurgy.

SUBSTANCE: product consists of following components, wt %: Zn 9.0-14.0, Mg 1.0-5.0, Cu 0.03-0.25, Fe <0.30, Si <0.25, Zr from 0.04 to less, than 0.3 and one or more elements chosen from group consisting of: Ti <0.30, Hf <0.30, Mn <0.80, Cr <0.40, V <0.40 and Sc <0.70, random elements and impurities, each <0.05, totally <0.15, and aluminium - the rest. The procedure for fabrication of product out of aluminium alloy consists in casting an ingot, in homogenisation and/or in preliminary heating the ingot upon casting, in hot treatment of the ingot into preliminary finished product with one or more methods, chosen from the group including rolling, extrusion and forging. Not necessarily, the preliminary treated product can be heated or hot treated and/or cold treated to a required shape of a blank; further formed blank is subjected to heat treatment to solid solution, to hardening blank heat treated to solid solution; not necessarily, hardened blank can be stretched or compressed, or cold treated by other way to stress relief, for example, by levelling sheet products or artificial ageing, till obtaining a required condition.

EFFECT: product with reduced tendency to forming hot cracks and with improved characteristics of strength, fracture toughness and hardness over 180 HB at artificially aged state.

32 cl, 6 tbl, 6 ex

FIELD: metallurgy.

SUBSTANCE: invention refers to deformed alloys of aluminium-zinc-magnesium-scandium system and to procedure for their production. Aluminium alloy contains from 0.5 to 10 wt % Zn, from 0.1 to 10 wt % Mg, from 0.01 to 2 wt % Sc, at least 0.01 wt % at least one alloying additive chosen from Ag at amount of up to 1 wt % and Sn at amount of up to 0.5 wt %, aluminium and unavoidable additives - the rest. The procedure consists in production of the said aluminium alloy, in homogenisation, in extrusion, in treatment for solid solution, in quenching, in straightening with drawing and in ageing.

EFFECT: alloys possess good qualities such as relatively high strength and excellent corrosion resistance.

33 cl, 3 dwg, 4 tbl

FIELD: metallurgy.

SUBSTANCE: aluminum based protective alloy comprises, in mass %, 4-5 of zinc, 0.01-0.06 of indium, 0.01-0.1 solder, 0.01-0.1 of zirconium, and aluminum the remainder.

EFFECT: enhanced corrosion protection.

2 tbl

Aluminum-base alloy // 2280092

FIELD: metallurgy.

SUBSTANCE: invention relates to aluminum-base alloys used for making deformed semifinished products used in industry and building. Proposed alloy comprises the following components, wt.-%: zinc, 4.5-5.6; magnesium, 1.6-2.1; manganese, 0.2-0.8; scandium, 0.03-0.09; zirconium, 0.05-0.12; copper, 0.1-0.3; titanium, 0.01-0.07; molybdenum, 0.01-0.07; cerium, 0.001-0.01, and aluminum, the balance, wherein the ratio content of zinc to magnesium = 2.6-2.9. Invention provides the development of alloy providing enhancing corrosion resistance of articles.

EFFECT: improved and valuable properties of alloy.

2 tbl, 1 ex

FIELD: metallurgy of aluminum alloys; manufacture of wrought semi-finished products for transport engineering and other industries.

SUBSTANCE: proposed alloy includes the following components, mass-%: zinc, 3.6-4.1; magnesium, 0.6-1.1; manganese, 0.2-0.5; zirconium, 0.05-0.12; chromium, 0.05-0.15; copper, 0.1-0.2; titanium, 0.01-0.06; molybdenum, 0.01-0.06; the remainder being aluminum.

EFFECT: enhanced corrosion resistance and technological ductility of semi-finished items at plastic metal working.

2 tbl, 1 ex

FIELD: metallurgy.

SUBSTANCE: invention relates to aluminum-base material. Proposed material comprises the following components, wt.-%: zinc, 6-8; magnesium, 2.5-3.5; nickel, 0.6-1.4; iron, 0.4-1.0; silicon, 0.02-0.2; zirconium, 0.1-0.3; scandium, 0.05-0.2, and aluminum, the balance wherein the temperature of equilibrium solidus of material is 540°C, not less, the hardness value of material is 200 HV, not less. Invention provides the development of the novel high-strength material designated for both producing fashioned ingots and deformed semifinished product possessing high mechanical properties. Invention can be used in making articles working under effect of high loading, such as car articles and sport inventory articles.

EFFECT: improved and valuable properties of material.

4 cl, 2 dwg, 4 tbl, 3 ex

FIELD: nonferrous metallurgy.

SUBSTANCE: invention relates to ultrastrong economically alloyed aluminum-based alloys belonging to system Al-Zn-Mg-Cu. Alloy and article made therefrom are, in particular, composed of, %: zinc 3.5-4.85, copper 0.3-1.0, magnesium 1.2-2.2, manganese 0.15-0.6, chromium 0.01-0.3, iron 0.01-0.15, silicon 0.01-0.12, scandium 0.05-0.4, at least one element from group: zirconium 0.05-0.15, cerium 0.005-0.25, and aluminum - the rest.

EFFECT: increased characteristics of corrosion resistance, bondability with all welding techniques, and lowered fatigue crack growth rate.

2 cl, 2 tbl

FIELD: metallurgy.

SUBSTANCE: invention proposes alloy containing the following components, wt.-%: zinc, 5.4-6.2; magnesium, 2.51-3.0; manganese, 0.1-0.3; chrome, 0.12-0.25; titanium, 0.03-0.10; zirconium, 0.07-0.12; beryllium, 0.0002-0.005; sodium, 0.0001-0.0008; copper, 0.2, not above; iron, 0.3, not above; silicon, 0.2, not above, and aluminum, the balance. Alloy provides enhancing uniformity of armor structure and its welded seams, stable armor resistance of extended armor welded seams independently on disposition of units to bed welded, elimination of splits from armor rear site in case its resistance to a missile impact, elimination possibility for reducing tenacity of armor during its exploitation including using under conditions of combination with dynamic protection of armored-body and armor-carrying mechanized objects. Invention can be used in producing armor for individual protection and for protection of mechanized armor-carrying objects against effecting agents.

EFFECT: improved and valuable properties of alloy.

1 tbl

FIELD: metallurgy.

SUBSTANCE: invention proposes alloy comprising the following components, wt.-%: zinc, 4.7-5.3; magnesium, 2.1-2.6; chrome, 0.12-0.25; titanium, 0.03-0.10; zirconium, 0.07-0.12; beryllium, 0.0002-0.005; iron, 0.05-0.35; silicon, 0.05-0.25; boron, 0.0003-0.003; sodium, 0.0001-0.0008; copper, 0.2, not above, and aluminum, the balance. Proposed alloy provides enhancing the armor structure uniformity and its welded joints, to provide stable armor resistance of extended welded joints of armor and independently of location of units to be welded, to exclude splitting off from rear side of armor in case armor not piercing by a missile, to exclude possibility for decreasing tenacity of armor in exploitation including using under conditions of combination with external dynamic protection of armored-carcass and armored-carrying mechanized objects. Invention can be used in producing armor for armored-carrying equipment for protection against effect of affection agents.

EFFECT: improved and valuable technical properties of alloy.

FIELD: metallurgy.

SUBSTANCE: invention proposes alloy comprising the following components, wt.-%: zinc, 4.7-5.3; magnesium, 2.1-2.6; manganese, 0.05-0.15; chrome, 0.12-0.25; titanium, 0.03-0.10; zirconium, 0.07-0.12; beryllium, 0.0002-0.005; iron, 0.05-0.35; silicon, 0.05-0.25; sodium, 0.0001-0.0008; copper, 0.2, not above, and aluminum, the balance. Proposed alloy provides enhancing armor structure uniformity and its welded joins, to provides stable armor resistance of armor welded joints being independently on location of units to be welded, to exclude splitting off from rear side of armor in case armor not piercing by missile, to provide high tenacity of armor including its using under conditions of combination with external dynamic protection of armored-carcass and armor-carrying mechanized objects. Invention can be used in producing armor for armor-carrying equipment for its protection against protection of affecting agents.

EFFECT: improved and valuable properties of alloy.

FIELD: metallurgy.

SUBSTANCE: the present innovation deals with obtaining aluminum-based alloys necessary for manufacturing stampings, particularly those of automobile wheels disks. The alloy in question has got the following composition, weight%: copper 0.8-2.2; magnesium 1.2-2.6; manganese 0.2-0.6; iron ≤0.25; silicon ≤0.20; zinc 5.0-6.8; titanium ≤0.1; chromium 0.08-0.17; zirconium 0.01÷0.12; boron 0.0008-0.005; antimony 2.5-3.5; indium 2.5-3.5; boron 0.4-0.5; hydrogen (0.3-4.1)10-5, aluminum - the rest. The alloy in question is of optimal combination of strength and plasticity that guarantee the required level of performance characteristics of automobile wheels disks, the decrease of their weight in combination with high technological effectiveness at volumetric stamping, especially complex-shaped articles.

EFFECT: higher strength and plasticity of the alloy.

2 cl, 1 ex, 3 tbl

Aluminum-base alloy // 2319762

FIELD: metallurgy, alloys.

SUBSTANCE: invention relates to compositions of deformable aluminum-base alloys. Proposed alloy comprises the following components, wt.-%: zinc, 5.0-7.0; magnesium, 0.4-0.8; copper, 0.8-1.2; manganese, 0.8-1.2; zirconium, 0.2-0.3; titanium, 0.2-0.3; niobium, 0.2-0.3; nickel, 3.0-5.0; boron, 0.02-0.03, and aluminum, the balance. Proposed alloy possesses the enhanced strength. Proposed alloys can be used in aircraft construction and automobile construction.

EFFECT: improved and valuable property of alloy.

1 tbl

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