Products from aluminium alloy and method of artificial age-hardening

FIELD: metallurgy.

SUBSTANCE: invention relates to constructional elements from aluminium alloy, in particular, for space industry. The slab is made with the thickness at least 4 inches from aluminium alloy, which contains (wt %): Zn - from 6.4 up to 8.5, Mg - from 1.4 up to 1.9, Cu - from 1.4 up to 1.85, Zr - from 0.05 up to 0.15, Ti - from 0.01 up to 0.06, Fe - up to 0.15, Si - up to 0.12, the rest is aluminium, accompanying elements and impurities.

EFFECT: improved combination of durability and resistance to fracturing is provided, and also the resistance to a fracturing as a result of stress corrosion is provided, especially in conditions of marine atmosphere.

10 cl, 14 dwg, 14 tbl, 3 ex

 

The technical field to which the invention relates.

The present invention relates to aluminum alloys, in particular aluminum (Al) alloys of the 7000 series (or 7XXX) in accordance with the designation of the Aluminum Association. More specifically, the present invention relates to products made of aluminum alloy with a relatively large thickness, that is, about 2-12 inches (5.1 to 30.5 cm). Although the present invention is mainly directed to products of the type rolled plate, it can also be used for the products obtained by the process of extrusion or forged products. When using the present invention in practice, the parts made from such raw materials/products with thick cross-section, have exceptional combinations of strength properties-viscosity, which makes them suitable for use as structural components in different applications in the aerospace industry, in the form of details with a large thickness profile or parts with thinner cross-sections, which were manufactured by machining from a material with a profile with a large cross-section. Thanks to the present invention are also achieved a significant increase in corrosion resistance, in particular resistance to cracking due to corrosion under on ruscoe (or "SCC"). Examples of items of structural elements made of this alloy include integral elements of the spars and similar items manufactured by the method of mechanical processing of the processed pressure sections with a large cross-section, including rolled plate. Such elements of the spar can be used in the design of the compartment of the wing of high-capacity aircraft. The present invention, particularly suitable for manufacturing high-strength aircraft parts, obtained by extrusion and forging, such as, for example, the main beams of the chassis plane. Such aircraft include commercial passenger jet Airliners, cargo transport aircraft (such as used by the postal services) and some military aircraft. To a lesser extent, the alloys in accordance with the present invention, suitable for use in other aircraft, including but not limited to, turbo-prop aircraft. In addition, in accordance with the present invention can be manufactured parts, not intended for the aerospace industry, such as various mouldings in the form of a plate with a profile with a large cross-section.

With increasing size of new jets or with increasing demands to existing models of jet liners to ensure the possibility of increasing the load is stretch that far and/or increase the range, to improve performance and increase savings need to reduce the weight of structural elements, such as the fuselage, wings and parts of the side members. Industry aircraft satisfies this need by establishing technical conditions determining the higher strength metal parts, which provides the possibility of reducing the thickness of their cross-section to reduce weight. In addition to the strength, are also critical parameters such as durability, resistance to damage of materials to ensure safe aircraft structure. Such conditions, the defining characteristics of a variety of materials used in aircraft construction, allow for the possibility of creating modern designs, resistant to damage, which combined the principles of safe design technology periodic inspection.

The traditional design of the wing contains a compartment of the wing which is generally indicated by the numeral 2 on the attached figure 1. It is located outside of the fuselage and is the main load-bearing element of the wing, passing essentially perpendicular to the plane of figure 1. Such compartment 2 wing comprises upper and lower casing 4 and 6 of the wing, between which there is a vertical structural members or rails 12 and 20 were Asia between the upper and lower skins of the wing or connecting them. Compartment of the wing also contains edges that can be, in General, from one side member to the other. These ribs are parallel to the plane of figure 1, while the wing and the side members are perpendicular to the plane of figure 1. During the flight, the upper structural elements of the wing commercial aircraft experience a compression load, which creates tension compression high level at an acceptable level of resistance to cracking. Covering the top of the wing is the largest of modern aircraft usually made of aluminum alloys of the 7XXX series, such as aluminum alloy 7150 (re-edition of the American patent # US 34008) or aluminum alloy 7055 (U.S. patent # US 5221377). Since the structural details of the lower part of the same wing aircraft experience during flight action voltage, they require a large damage than for parts of the upper part of the wing. Although for maximum weight reduction you may want to construct the lower part of the wing using a more durable alloy, the characteristics of resistance to damage such alloys are often insufficient for use in these structural elements. In this regard, the most modern manufacturers of jet aircraft is used to lower the wing more damage-resistant aluminum alloy series 2XXX, such as aluminum alloy 2024 or 2324 (U.S. patent # US, 294625), both of these alloy 2XXX have a lower strength than the alloy 7XXX series, which made parts of the upper part of the wing. Used elements and the alloy compositions correspond to well-known standards products Aluminum Association.

The stiffness of the casings 4 and 6 are respectively top and bottom of the wing, as shown on the attached figure 1, usually given by a longitudinally passing items 8 and 10 of the stringers. Such elements of the stringers can be made with different shape, including "J", "I", "L", "T" and/or "Z" configuration of the cross section. The elements of the stringer is usually fixed on the inner surface of the covering of the wing, as shown in figure 1, and as the attachment usually use rivets. Item 8 stringer top wing and shelves 14 and 22 of the upper side members currently produce from alloys series hhha, the stringer 10 of the lower part of the wing and the shelves 16 and 24 of the lower side members are made of alloy 2XXX series for the same structural reasons described above, to provide the relative strength and resistance to damage. The elements 18 and 26 vertical wall of the side member is also made of alloys of 7XXX and attached to both the upper and lower shelves of the spar, passing in the longitudinal healthy lifestyles the research Institute of the wing, constituent elements of the side members 12 and 20. This traditional design of the spar, also known as "built-in" spar, includes top shelf 14 or 22 of the side member, the wall 18 or 20 and the bottom shelf 16 or 24 spar with fasteners (not shown). It is obvious that the fasteners and holes for fasteners in connections with these spars are structurally weak connections. To ensure the structural integrity of the built spar, such as 18 or 20, many details of items such as wall and/or shelf spar must be made thicker, which increases the weight of the whole structure.

One possible design approaches aimed at reducing the above-mentioned unwanted weight of the spar is in the manufacture of the upper part of the spar, the walls and bottom of the spar by means of mechanical processing of simple sections thicker profile, such as a plate, obtained as the product of the aluminum alloy, usually with removing enough metal for more complex parts, albeit with a smaller cross-section or shape, such as a spar. Sometimes such a machining operation is called the "bow" part of the product in the form of a plate. With this design it is possible to eliminate the necessity of using compounds of the wall with the upper part of LON is erona and walls with the lower part of the spar. Spar of this type, made in the form of a single detail, sometimes referred to as "solid spar", and it can be manufactured by machining from a thick plate by extrusion or forging. Solid spars must not only weigh less than the multiple items, they must also be less costly in the manufacture and Assembly by eliminating the need to use fasteners. The ideal alloy for the manufacture of solid spars must have the strength characteristics of the alloy used in the upper part of the wing, combined with requirements for resistance to cracking/resistance to damage to alloy the lower part of the wing. Existing commercial alloys used in aircraft, do not satisfy this combination of requirements preferred properties. Lower strength alloy used for the lower skin of the wing, such as alloy 2024-T, it is not possible to provide a safe level of resistance to the loads transmitted from at high loads the upper part of the wing, if only the area of its cross section will not be significantly increased. This in turn leads to undesirable increase in the weight of the whole structure of the wing. Conversely, the use of structural elements of the upper part of the wing strength properties of alloy 2XXX also p is igodit to the total increase in weight.

In the design of large jet aircraft requires the use of very large wings. For the manufacture of one-piece side members of such wings, you must use the products of a thickness of 6-8 inches (15,2-20.3 cm) or more. For the manufacture of parts with large cross-section is often used alloy 7050-T. The industry standard for plates of alloy 7050-T thickness of 6 inches (15.2 cm), as indicated in the Specification AMS 4050F aerospace materials, contains a requirement to provide minimum values of yield strength in the longitudinal (L) direction at the level of 60 thousand pound/inch2(414 MPa) and cracking under conditions of plane strain, or Klc(L-T), 24 thousand (lb/in2)inch1/2. For the alloy with the same hardness and thickness values are defined in the transverse direction (LT and T-L) - 60 thousand (lb/in2) (414 MPa) and 22 thousand (pounds/inch2)inch1/2respectively. For comparison, an aluminum alloy 7055-T for the upper parts of the wing thickness approximately from the 0.375 to 1.5 inches allows to provide the minimum value of the yield strength of 86 thousand (lb/in2) (593 MPa) in accordance with MIL-HDBK-5H. If one-piece spar alloy 7050-T with a minimum value of the yield strength of 60 thousand (lb/in2) (414 MPa) to use with the above-mentioned alloy 7055, the overall strength of the top features of the cu plating the La will not be fully used with the most effective weight reduction. Therefore, for the manufacture of solid configurations of the spar with a large cross-section that are required currently for the design of new jet liners required high-strength aluminum alloys with sufficient resistance to cracking. The above is just one concrete example of the advantages of aluminum material with high strength and stability in large cross sections, but in modern aircraft construction there are many other examples of parts with similar requirements, such as wing ribs, partitions or stringers, panels or sheathing elements of the wing frame of the fuselage floor beams or frames, and even beams landing gear or various combinations of these structural elements of the aircraft.

Known for the ability to change the stiffness of the metal as a result of various processing by artificial aging, which provides different levels of strength and other performance characteristics, including resistance to corrosion and resistance to cracking. 7XXX series alloys are often manufactured and supplied commercially in such a state of artificial aging, as alloys peak strength (Type"T6"or "perestaranie" alloys (Type T7"). In U.S. patent No. 4863528, 4832758, 4477292 and 5108520 described compositions of the alloys of the 7XXX series with diazonamide strength and performance. The entire contents of these patents is fully herein by reference.

Specialists in the art it is well known that for a given forged alloy 7XXX series modes vacation, providing the peak value of the strength or type T6 provide the highest strength value, but combined with a relatively low value resistance to cracking and corrosion resistance. For these alloys also know that most perestroennih formulations, such as typical compositions treated in vacation mode type T73, have the highest value of resistance to cracking and corrosion resistance, but have substantially lower value of relative strength. In the manufacture of parts intended for use in the aerospace industry, the designers of the parts must choose the appropriate composition having properties somewhere between these two extremes, in order to ensure the possibility of its use in a particular case. A more complete description of the modes of vacation, including the designation "T-XX"can be found in the publication of the Aluminum Association's Aluminum Standards and Data 2000, which is well known in the art.

Most treatment options alloys intended for use in the aerospace industry require COI is lesofat thermal processing of solid solution (or "SHT"), followed by annealing and subsequent artificial aging to obtain the strength and other characteristics. However, finding ways to improve properties in large cross sections is faced with two natural phenomena. First, increasing the thickness profile of the product, the rate of hardening in the internal cross section of the product, of course, reduced. This reduction in turn leads to loss of strength and resistance to cracking to form the product with a large cross-section, in particular in the inner regions of the cross section. Specialists in this field of technology called this phenomenon "sensitivity to hardening". Secondly, there is also a well-known inverse relationship between strength and resistance to cracking so that if construction details are developed for large values of strength load, their relative characteristics resistance decreases and Vice versa.

For a better understanding of the present invention is useful to consider some well-known trends in the use of commercial 7XXX series alloys used for aerospace industry. In aluminium alloy 7050, for example, Cr is replaced by Zr, which is used as a colloidal agent for increasing the degree of control structure is Roy grains and increases the content of Cu, and Zn compared to the older alloy 7075. Alloy 7050 has significantly better (i.e. lower) sensitivity to hardening compared with the previous alloy 7075, making aluminum alloy 7050 is the main alloy used for parts in the aerospace industry with a large cross-section, made in the form of plates, extruded parts and/or forged parts. For use in the upper part of the wing where you want to use materials with high strength-resistance, the minimum values of Mg and Zn in the composition of the aluminium alloy 7050 slightly increase, so it's registered with the Aluminum Association alloy 7150, which is a variation of alloy 7050. Compared to its predecessor 7050 minimum content for Zn alloy 7150 increased from 5.7 to 5.9 wt.%, and the minimum level of Mg was increased from 1.9 to 2.0 wt.%.

In the end, we have developed the new composition of the alloy for the upper skin of the wing. This alloy 7055 has 10% of the best value of the yield strength in compression, in particular, through the use of a wider range of Zn, from about 7.6 to 8.4 wt.%, at similar levels of Cu and slightly more narrow range of Mg (1.8 to 2.3% wt.) compared with the alloy 7050 7150 or.

Made in the past, attempts to further increase the strength by increasing l giraudy elements and optimization of composition) were associated with the need to increase the purity of the metal and control of the microstructure using thermomechanical processing ("TSR") to provide, among other properties, improve the strength and time to fatigue fracture. In U.S. patent No. 5865911 described a significant improvement in the resistance with equivalent strength plate alloy 7XXX series. However, the sensitivity of this alloy tempering with more thickest cross-sectional profile, probably leads to another significant undesirable properties.

Alloy 7040, registered with the Aluminum Association, provides the following ranges of the main alloying elements: 5,7-6,7% wt. Zn, 1.7 to 2.4 percent wt. Mg and 1.5-2.3% of wt. Cu. In the literature, namely in the publication authors Shahani and others, "High Strength 7XXX Alloys For Ultra-Thick Aerospace Plate: Optimization of Alloy Composition", PROC. ICAA 6, v. 2, pp/ 105-1110 (1998) and in U.S. patent number US 6027582, stated that the developers alloy 7040 sought to balance optimization between alloying elements to improve the strength and other characteristics, avoiding excessive additives to minimize sensitivity to hardening. Although alloy 7040 profiles with a large cross-section announced some improvements to properties in comparison with alloy 7050, these improvements are still insufficient for the current needs of designers commercial aircraft.

The present invention differs in several key aspects from the alloys supplied at present is its time on a commercial basis for use in the aerospace industry. The main alloying elements for several currently used in aerospace alloys 7XXX specified by the Aluminum Association in the following form:

It should be noted that aluminum alloys 7075, 7050, 7010 and 7040 put in the aerospace industry both in the form of a profile with a large cross-section, and in the form of a thin (up to 2 inches (5.1 cm) profile, while other alloys (7150 and 7055), mostly delivered in the form of a thin profile. In contrast to these commercial alloys the preferred alloy in accordance with the present invention, contains approximately 6.9 to 8.5 wt.%. Zn, from 1.2 to 1.7 wt.%. Mg, from 1.3 to 2% wt. Cu, 0.05 to 0.15 wt.%. Zr, and the rest of the content, essentially, comprise aluminum, incidental elements and impurities.

The present invention solves the above problems of the prior art using the new aluminum alloy 7XXX series, which in profile with a large cross-section exhibits a significantly reduced sensitivity to hardening, providing a significantly higher level of durability and resistance to cracking than was possible until now. The alloy, in accordance with the present invention, has a relatively high content of zinc (Zn) at a lower content of copper (Cu) and magnesium (Mg) is compared with the above commercial alloys 7XXX, used in the aerospace industry. For the purposes of the present invention the combined content of Cu+Mg is usually less than about 3.5% and preferably less than approximately 3.3%. When the above composition is subjected to a preferred practice, a 3-stage aging, which is described in more detail below, the resulting thick form forged product (plate, the parts obtained by the method of extrusion or forging) are more preferred combination of performance, such as durability, resistance to cracking and fatigue, along with exceptional resistance to cracking due to corrosion under stress (SCC), especially in atmospheric conditions or in the test conditions in the marine atmosphere.

Known three-stage or three-stage examples of aging 7XXX aluminum alloys of the prior art. Their examples are described in U.S. patent number US 3856584, 4477292, 4832758, 4863528 and 5108520. The first stage/stage many of the above methods of the prior art usually performed at a temperature of about 250°F (121°C). The preferred first stage of aging of the composition of the alloy, in accordance with the present invention, is carried out in the temperature range of about 150-275°F (66-135°C), preferably about 200-275°F (93-135°C) and more prefer the Ino from about 225 or 230°F (107 or 110°C) up to about 250 or 260°F (121 or 127°C). This is the first stage or step may include processing at two temperatures, for example at 225°F (107°C) for approximately 4 hours and 250°F (121°C) for approximately 6 hours, both these period are only the first stage, i.e. the stage preceding the second (for example, processing at a temperature of approximately 300°F (149°C) stage described below). Most preferably, the first stage of aging, in accordance with the present invention, performed at a temperature of approximately 250°F (121°C) for at least about 2 hours, preferably for about 6 to 12 and sometimes up to 18 hours or more. However, it should be noted that smaller periods of exposure may be sufficient depending on the size of the part (i.e. its thickness) and the complexity of the shape given the rate of temperature increase in the use of this equipment (i.e. taking into account the relatively slow increase of heating temperature), the time to heat the workpiece must be taken into account when calculating the shorter the exposure time at the processing temperatures for these alloys.

Preferred treatment during the second stage in some methods of the prior art, including practice 3-step artificial aging, is usually held at a temperature of about 350 or 360°F (177 or 182°C) or you is e, followed by the third stage of aging, similar to the first stage performed at a temperature of approximately 250°F (121°C). In contrast, the preferred second stage of aging, in accordance with the present invention, characterized in that the treatment is carried out at substantially lower temperatures, which are approximately 40-50°F (4.4 to 10°C) below. For the preferred embodiments of this 3-step method of aging the alloy 7XXX described here is the second of three phases or stages must occur at a temperature of from about 290 or 300°F (149°C) to about 330 or 335°F (166 or 168°C). More precisely, this second phase or stage of ageing should be carried out at a temperature of from about 305 to 325°F (152-163°C), the preferred temperature range of the second stage of aging is from about 310 to 320 or 325°F (154-160, 163°C). The preferred time for such processing of the second stage has an inverse dependence on the used temperature (temperature). For example, if you want to work, essentially, at a temperature of 310°F (154°C) or very close to it, would be enough to use the total exposure time of approximately 6-18 hours. More preferably, the aging in the second step must be within approximately 8 or 10 to 15 hours in total at this operating temperature is ur. At a temperature of approximately 310°F (160°C). the total processing time of the second stage should be in the range of about 6-10 hours, with the range from 7 or 8 to 10 or 11 hours is preferred. There is also a preferred aspect of the specified property, which must be taken into account when choosing the time of aging of the second stage and the temperature selection. In particular, it should be noted that the shorter the processing time at a given temperature allows to obtain relatively higher values of strength, while longer exposure allows you to provide the best properties corrosion resistance.

After the previous second stage of aging followed by a third stage of aging is performed at a lower temperature. When using blanks with a large cross-section, preferably, not slowly move from the second stage to the implementation of the third stage without taking extraordinary care in handling to ensure accurate matching of the temperature of the second stage and the total exposure time to prevent unnecessary processing at higher temperatures (temperatures used in the second stage). Between the second and third stages of aging metal products, in accordance with the present invention, can be intentionally removed from heating the furnace and subjected to a rapid cooling using a fan or similar device to a temperature of approximately 250°F (121°C) or below, sometimes even completely to room temperature. In any case, the preferred periods of the exposure time/temperature in the third stage of aging, in accordance with the present invention, similar to the values described for the first stage of aging, above, at temperatures from about 150 to 275°F (135°C), preferably from about 200 to 275°F (93-135°C) and more preferably from about 225 or 230°F (107 or 110°C) up to about 250 or 260°F (121 or 127°C). Although the above method improves specific properties, in particular the resistance to SCC for this new family of alloys 7XXX, it should be understood that the same combination of improved properties can be implemented using the same method 3-step aging for other 7XXX series alloys, including, but not limited to, alloys H (for example, aluminum alloys 7050 7150 or), aluminum alloys 7010 and 7040.

For newer and larger aircraft manufacturers need products of aluminum alloy profile with a large cross-section, with a yield strength in compression, which is approximately 10-15% higher than usually obtained by using the currently used aluminum alloys 7050, 7010 and/or 7040. In accordance with this requirement alloy type 7XXX in accordance with the present invention, satisfactory the et of the above requirements to ensure the values of the yield strength, at the same time having good properties of resistance to development of cracks. In addition, this alloy also exhibits exceptional resistance to cracking under corrosion under load under the condition of aging using the preferred three stages according to the method of artificial ageing, described in this application. Specimens plate thickness of six inches (15.2 cm), made of this alloy have been tested in laboratory scale to cracking due to corrosion under stress (SCC) under alternate immersion (AI) in a 3.5% salt solution. In accordance with these test metal samples with a large cross-section withstand without cracking for at least 30 days minimum load of 25 thousand (lb/in2) (173 MPa)applied in the direction of the shorter cross-section (or in the direction of "ST") to meet the conditions of leave kzt76, in accordance with the requirements of one of the main manufacturers of jet aircraft. Such samples of metal with a large cross-section must also meet the other requirements of this manufacturer of jet aircraft to ensure static and dynamic parameters.

When performing the initial wave of laboratory tests with alternating immersion (AI) SCC at higher load levels 35-45 implement the (lb/in 2) (242-311 MPa) samples of alloys with a large cross-section, in accordance with the present invention, past the artificial aging using well-known two-way vacation, showed some unexpected failures due to corrosion, even when the load level 25 thousand (lb/in2) (173 MPa) when first exposed to the test conditions SCC in the marine atmosphere. This was unexpected, because accelerated laboratory tests AI SCC is usually well correlated with atmospheric tests in the marine atmosphere and in an industrial environment. During this testing in an industrial environment the samples of the alloys in accordance with the present invention, during aging in 3 stages, as stated in the description of the present invention, does not result in failure after an 11-month aging in such atmosphere, as under load at 25 (lb/in2) (173 MPa), and at the level of 35 thousand (lb/in2) (242 MPa). Even though indicators of SCC in different atmospheres were not explicitly expressed in the specifications of the manufacturer of the aircraft for aircraft of the next generation, they, nevertheless, are important for critical applications in the aerospace industry, such as spars and ribs of the compartments of the wing of a jet plane. Thus, although the products are processed is by aging in two stages, can meet their requirements in the application in practice of the present invention, it is preferable described processing with artificial aging in three stages.

Known way to "improve stability" SCC of some 7XXX alloys hitherto consisted in the aging of the material usually by reducing its strength. This decrease in strength is undesirable for the whole of the wing spar, because the last machining a part with a large cross-section must still meet the extremely high standards of the yield strength in compression. Thus, there is a clear need to develop a method of artificial ageing, which will not require excessively sacrificing strength properties while improving corrosion resistance of aluminum alloys 7XXX. In particular, you want to develop a way of aging which will allow you to boost the performance of SCC in the conditions of the marine environment for these alloys to higher levels without compromising strength and/or other combinations of properties. The above method of aging in three stages, in accordance with the present invention, satisfies this requirement.

An important aspect of the present invention is based on the developed new aluminum alloy that PR is is significantly reduced sensitivity to hardening when the profile with a large cross-section, that is, when the thickness is greater than approximately 2 inches (5.1 cm), and more preferably at a thickness in the range from approximately 4 to 8 inches (a 10.2-20.3 cm) or more. With a wide range of content, the composition of the alloy essentially consists of: from about 6% Zn to about 9, 9.5 or 10 wt.%. Zn; from about 1.2 or 1.3% Mg to about 1,68, 1.7 or even 1,9% wt. Mg; from about 1.2, 1.3 or 1.4% of wt. Cu to an estimated 1.9 or even the 2.2 wt.%. Cu at % Mg ≤ (% Cu + 0.3 maximum); there is one or more elements selected from the group consisting of: up to approximately 0.3 or 0.4% wt. Zr, up to about 0.4% wt. Sc and up to approximately 0.3 wt.%. Hf, the rest of the content, essentially, comprise aluminum and incidental elements and impurities. Except when stated otherwise, such when you specify "present", the expression "up to" when referring to a number of item means that the item is used, if necessary, and includes zero amount of this specific element of the composition. Unless indicated otherwise, all percentages in the composition represent the weight percent (% wt.).

Used herein, the term "essentially does not contain" means that no special additives of this alloying element, but due to the presence of impurities and/or due to leaching when K is ntake with production equipment small number of such elements, however, can get into the final alloy product. However, it should be understood that the scope of the present invention should not/cannot be changed by simple addition of any such item or items in quantities which, otherwise, could affect the combination of properties required and achieved in the present invention.

When referring to any numerical range of values, such ranges should be understood as including any number and/or a fractional value between the specified minimum and maximum range. For example, range from approximately 6 to 10% wt. zinc should definitely include all intermediate values that are estimated at 6.1, or 6.2, 6.3 and 6.5%, continuing, therefore, and including values of 9.5, 9.7 and 9.9% zinc. The same applies to every other quantitative property in the description of the method of heat treatment (i.e. temperature) and/or to the range of content of the elements referred to in the present description. The maximum or "max" refers to the total value up to the specified values for content elements, time value and/or to other property values, for example a maximum of 0.04% wt. Cr, and the minimum ("min") applies to all values above the minimum value.

The term "random elements" may include relatively small amounts of Ti, b and D. the natives elements. For example, titanium with boron or carbon is used as an auxiliary substance spill alloy to control the grain size. The alloy described in this invention may contain as random elements to approximately 0.06% of wt. Ti or approximately from 0.01 to 0.06 wt.%. Ti and, if necessary, to approximately 0,001 or 0.03% wt. Sa, about 0.03% wt. Sr and/or approximately a 0.002% wt. Be. Random elements can also be present in significant quantities and can improve preferred or other characteristics alone without going beyond the scope of the present invention, if only the alloys remain the required characteristics specified in the present description, including reduced sensitivity to hardening and improved combination of properties.

This alloy may additionally contain a smaller number of other elements that are less preferred. Chromium, preferably, do not use, that is, maintain at or below about 0.1 wt.%. Cr. However, it happens that a very small amount of Cr may improve some properties in one or more specific embodiments, the use of the alloy in accordance with the present invention. In preferred at the present time options run-level Cr supports rivets below about 0.05 wt.%. Manganese also deliberately maintained at a low level, below about 0.2 or 0.3 of the total content % wt. The content of Mn is preferably not exceed about 0.05 or 0.1 wt.%. And at the same time can be one or more specific variants of the composition of the alloy, in accordance with the present invention, in which the intentional addition of manganese can have a positive result.

In the composition of the alloy can be used small amounts of calcium, primarily as a good deoxidiser item on the stages of the molten metal. Calcium supplements in the amount of about 0.03% wt. or more preferably approximately from 0.001 to 0.008 wt.%. (or 10-80 ppm) Ca also help prevent unpredictable cracking when filling larger molds alloy of the above composition. In cases where cracking is less critical, such as a round billet, forged parts and/or parts produced by extrusion, Ca do not want to add to the alloy or it can be added in smaller quantities. Strontium (Sr) can be used as a replacement or in combination with the above amounts of Ca for the same purpose. Usually supplements beryllium is used as deoxidiser element/means to eliminate cracking ex the Cai. Although for reasons of environmental protection, health and safety preferred options alloy, in accordance with the present invention essentially do not contain beryllium.

The content of iron and silicon must be maintained at a substantially low level, for example, not exceeding roughly 0.04 or 0.05 wt.%. Fe and about 0.02, or 0.03% of wt. Si or less. In any case still provides slightly higher levels of both impurities, about 0.08% wt. Fe and approximately 0.06% of wt. Si, which may be acceptable, although it is less preferred in accordance with the present invention. Even less acceptable iron at the level of about 0.15% wt. and the silicon level to 0.12 wt.%. may be present in the alloy in accordance with the present invention. For embodiments of the present invention in forming plates are acceptable even higher levels approximately to 0.25 wt.%. Fe and approximately 0.25 wt.%. Si or less.

We know from experience in the application of aerospace alloys of the 7XXX series, the iron may be contacted with the copper during curing. Therefore, the present description will be periodically made references to "effective concentrations of Cu, i.e. the amount of copper, not related to the presence of iron, or, in other words, the number is in Cu, in fact, available for solid solution alloy. In some cases it may therefore be preferable to consider an effective amount of Cu and/or Mg, in accordance with the present invention than, respectively, to regulate (or increase) the range of valid contents of Cu and/or Mg, measured to calculate the levels of Fe and/or Si, present and possibly interacting with Cu, Mg, or both elements. For example, the increase in preferred acceptable content of Fe from roughly 0.04 or 0.05 wt.%. up to about 0.1% wt. max can make the preferred increases the real, measurable minimum and maximum values of Cu, indicated as about 0.13 wt.%. Manganese is similar to copper in the presence of iron. Similarly, in relation to magnesium is known that silicon is associated with magnesium during solidification of alloys of 7XXX series. Therefore, in this description, it may be preferable to make reference to the amount of magnesium present as "effective Mg content", which realize the amount of Mg that is not associated with Si, and thus available for the solution at a temperature or temperatures used for dissolution of 7XXX alloys. Same as above valid adjusted to a range of copper content enhancement is the preferred maximum silicon concentration from about 0.02 to about 0.08 to or even up to 0.1 or 0.12 wt.%. Si can lead to the possibility of a similar regulation acceptable/measurable amount (both maximum and minimum) of magnesium present in the alloy in accordance with the present invention, in the direction of increase, perhaps to the level of the order of approximately 0.1 to 0.15 wt.%. Mg.

When a narrow range of composition in accordance with the present invention may contain from about 6.4 or 6.9 to 8.5 or 9% wt. Zn, about 1.2 or 1.3 to 1.65 or 1,68% wt. Mg, about 1.2 or 1.3 to 1.8 or 1.85% wt. Cu and approximately from 0.05 to 0.15 wt.%. Zr. If necessary, the latter composition may contain up to 0.03, and 0.04, or 0.06% wt. Ti, up to about 0.4% wt. Sc and up to approximately 0,008% wt. Sa.

If a more narrow definition is preferred in the present ranges of compositions in accordance with the present invention, contain approximately 6.9 or 7 to about 8.5 wt.%. Zn, from about 1.3 or 1.4 to about 1.6 or 1.7% by weight. Mg, from about 1.4 to about 1.9% of wt. Cu, and from about 0.08 to 0.15 or 0.16% of wt. Zr. % Mg does not exceed (% Cu+0,3), preferably does not exceed (% Cu+0,2) or, even better, (% Cu+0,1). For the above preferred embodiments, the content of iron and silicon is maintained at a relatively low level, at or below roughly 0.04 or 0.05 wt.%. for each element. The preferred composition contains: approximately 7-8% wt. Zn, approximately from 1.3 to 1.68 wt.%. Mg and approximately 1.4 to 1.8 wt.%. Cu, more preferably magnesium content equal to the content of copper or, even better, % wt. Mg < % wt. Cu. Also preferably, the ranges of the content of magnesium and copper, in accordance with the present invention, when combined, does not exceed about 3.5% wt. in sum, when, in the most preferred embodiment, % wt. Mg+% wt. Cu is approximately 3.3V.

Alloys in accordance with the present invention, can be prepared using more or less conventional ways, including smelting and direct casting (DC) in filling the form. Can also be used in conventional additives, grinding grain, such as containing titanium and boron or titanium and carbon, as is well known in the art. After the usual removal of the surface layer of the ingot (if necessary) and homogenization of such ingots undergo further processing such as hot rolling, in the form of a plate or extrusion, or forging with the receiving section of a special form. Typically, the thickness of the cross section is about 2 inches (5.1 cm) or more and, more typically, of order 4, 6, 8 or up to 12 inches (10,2, 15,2 20,3, 30.5 cm) or more. In the case of a plate thickness of approximately 4-8 inches to 10.2-20.3 cm) above the plate is subjected to heat treatment in solution (SHT) and let go, then remove stress by using, for example, stretching and/or compression by about 8%, for example, approximately 1-3%. Then give the desired structural shape by mechanical processing plate of these sections, cooked, often after artificial aging for forming parts of desired shape, for example, one-piece wing spars. Similarly, after the operations SHT, vacations, often relieving mechanical stresses and artificial aging is followed by the production of sections with a large cross-section by way of extrusion and/or processing forging. A good combination of properties is required for all values of thickness, but especially useful when the size of the profile, when used with increasing thickness also increases the sensitivity of the product to be hardening. Therefore, the alloy in accordance with the present invention, particularly suitable for the manufacture of products with a profile with a large cross-section comprising, for example, from 2-3 inches (5.1 to 7.6 cm) in thickness and up to 12 inches (30.5 cm) or more.

Brief description of drawings

Figure 1 shows a view in cross section of conventional design section of an airplane wing, comprising front and rear spars built using conventional composite construction of the three parts.

On Phi is .2 shows a graph, representing the two calculated curves cooling, intended to approximate the cooling rate in the plane passing through the middle of the profile, for plate thickness of 6-8 inches (15,2-20.3 cm), manufactured in the factory when quenched by spraying, are imposed on two experimental cooling curves, simulating the speed of the cooling plate thickness of 6 inches and 8 inches (15,2 and 20.3 cm).

Figure 3 shows a graph representing the value of the yield stress in tension in the longitudinal direction TYS (L) depending on the resistance to longitudinal cracking Kq(L-T) relationships for selected alloys in accordance with the present invention, and other alloys, including alloys of type 7150 and 7055, which are used for comparison or as "reference values", all data were obtained on the basis of a simulated speed of hardening in the plane of the mid profile (or "T/2") plate thickness of 6 inches (15.2 cm), detail, obtained by extrusion, or forging.

Figure 4 shows a graph similar to figure 3, representing the value of the yield stress in tension in the longitudinal direction TYS (L) depending on the resistance to cracking of the Kq(L-T) for selected alloys in accordance with the present invention, and other alloys, including reference values for alloys 7150 and 7055, all data received is based simulation speed of hardening in the plane of the mid profile for slab thickness of 8 inches (20.3 cm) details obtained by extrusion, or forging.

Figure 5 shows a graph depicting the effect of zinc content on the sensitivity to hardening, which are represented by arrows indicating the direction of change of the values of the yield strength tensile TYS simulated tempering plate thickness of 6 inches (15.2 cm).

Figure 6 shows a graph depicting the effect of zinc content on the sensitivity to hardening, which are represented by arrows indicating the direction of change of the values of the yield strength tensile TYS simulated tempering plate thickness of 8 inches (20.3 cm).

7 shows a graph depicting the diagram of the interdependence of the values of the yield strength tensile TYS (L) and resistance to cracking under plane deformation Klc(L-T) in a quarter plane (T/4) profile plate for full-scale production of a thickness of 6 inches (15.2 cm) with the use of the alloy in accordance with the present invention, where the plotted line (M-M) current extrapolated minimum values for comparison with values found in the literature for aluminum alloys 7050 and 7040.

On Fig shows a graph depicting the influence of the thickness of the cross section values of the yield strength tensile TYS, in the form factor of the properties of sensitivity to hardening to study the properties with the lava of full-scale production molded by forging forgings in accordance with the present invention, as compared with aluminum alloy 7050.

Figure 9 shows a graph which shows a comparison of the longitudinal values of the yield strength tensile TYS (in thousands of pounds per square inch) depending on the conductivity of the EU (in % of values IACS (international annealed copper) for specimens plate thickness of 6 inches (15.2 cm) of the alloy, in accordance with the present invention, after aging using the well-known 2-step method compared to the preferred method 3-step aging, which is described below. In this drawing, most notably an unexpected and significant increase in strength observed at the same level of the EU, or a substantial increase in the level of the EU, observed at the same value of strength for the samples that underwent the 3-step aging, compared with samples that undergo a 2-stage aging. In each case, the first stage of aging was carried out at a temperature of 225°F (107°C)250°F (121°C) or at both temperatures, followed by the second stage of aging at a temperature of approximately 310°F (154°C).

Figure 10 shows a graph representing the characteristics of corrosion under load SCC in terms of the marine atmosphere alloys underwent 2-stage aging in comparison with the 3-step aging on the one I preferred composition of the alloy under different levels of stress in the short transverse direction (ST), and that is the visual representation of the data collected in table 9 below.

Figure 11 shows a graph representing the characteristics of corrosion under load SCC in terms of the marine atmosphere alloys underwent 2-stage aging, compared with the 3-step aging for one preferred composition of the alloy under different levels of stress in the short transverse direction (ST), which is the visual representation of the data collected in table 10, below.

On Fig shows a graph of the values of the time to fatigue failure with the holes in the direction of orientation of the L-T for sample plates with different sizes, in accordance with the present invention, on which were drawn the strip 95% confidence interval for the values of S/N (dashed line) and extrapolated current characteristics preferred minimum (solid line a) and the results of their comparison with the values specified one of the manufacturers of jet aircraft, for products in the form of plates of alloys 7040/7050-T and 7010/7050-T, albeit in the other direction orientation (T-L).

On Fig shows a graph of the values of the time to fatigue failure with the holes in the direction of orientation of the L-T for forgings of different sizes, in accordance with the present invention, on which were drawn the line is renego values (dashed) and extrapolated current characteristics preferred minimum (solid line).

On Fig shows a graph depicting curves of the rate of growth of fatigue cracking (FCG) in the direction L-T and T-L for plates and forgings of different sizes, in accordance with the present invention, on which was drawn the curve extrapolated current characteristics of the preferred maximum FCG (solid line With), and it held its comparison with FCG curves presented on Fig specified one of the manufacturers of jet aircraft for the same range of sizes of commercially available plates of alloys 7040/7050-T, in the same directional orientation (L-T and T-L).

Detailed description of the invention

Of interest, the following mechanical properties of the plate with a large cross-section, obtained by extrusion or forging, for the structural elements of the aircraft, as in the case of use as structural elements in other industries, except aircraft, which include strength as compression for covering the upper side of the wing, and when the tension is for covering the lower side of the wing. Also important parameters are: resistance to cracking as in plane strain and in the plane stress state and characteristics of corrosion resistance, such as resistance to delamination and resistance to cracking under corrosion under loading is coy, and fatigue parameters representing the values of the time to fatigue fracture smooth material and material with an open hole (S/N), as well as resistance to growth of fatigue cracking (FCG).

As described above, one-piece wing spars, ribs, walls and panels covering the wing with integral stringers can be made by machining of the plates with a large cross-section or other products of extruded or forged shapes, which were subjected to heat treatment in the solid solution hardening, removing mechanical stress (if necessary) and artificial aging. This is not always possible to conduct heat treatment in the solid solution and rapid hardening themselves finished structural elements, since rapid cooling during hardening can create residual stress and lead to distortions of shape and size. Residual stresses caused by quenching, can also lead to cracking due to corrosion under stress. Similar distortions of the size due to the rapid quenching can lead to the need for re-treatment for straightening parts, form distortion which complicate the standard Assembly. Using the present invention can be manufactured more options of components/products for aerocon the practical industry, including, but not limited to: large frame and fuselage ribs for commercial jet aircraft, pressure treated boards for the top and bottom cladding smaller aircraft designed for flight at the local lines, beams landing gear or gender different jets, even frames, the elements of the fuselage and the lining elements of the wing fighter aircraft. In addition, the alloy in accordance with the present invention may be molded in the form of various small forged parts and other pressure treated structural elements of the aircraft, which currently is made from aluminium alloy 7050 or 7010.

Although the best mechanical properties easier provided with thin cross sections (because of the more rapid cooling of these parts prevents bicrystalline alloying elements), rapid hardening can cause excessive distortion. Within the practical feasibility of such details may be subject to mechanical alignment and/or flattening, using the methods of removing residual stress, after which these parts are subjected to artificial aging.

As described above, the heat treatment in the solid solution and the hardening of the parts with a thick profile very important indicator of the varieties is by the sensitivity of the aluminum alloy to be hardened. After heat treatment in the solid solution, it is preferable to provide rapid cooling of the material for the preservation of various alloying elements in solid solution, not letting them yet to crystallize from solution in the form of large forms, which are formed during slow cooling. When such large forms are formed large crystals, which leads to deterioration of mechanical properties. In products with a large cross-section greater than 2 inches (5.1 cm) thick at the point of greatest cross-section, and more specifically, the thickness of 4-8 inches (a 10.2-20.3 cm) or more, the environment hardening acting on the external surface of this piece (plate, forgings or items obtained by extrusion), can not effectively remove heat from the inner areas, including the Central region (or region of a plane mid profile (T/2) or the area of a quarter plane profile (T/4) of such material). This is due to the physical distance to the surface and due to the fact that heat is released through the metal, taking into account the heat conduction, the value of which depends on the distance. When the small cross section of the product of the rate of hardening in the plane of the middle of the profile, of course, will be higher than the rate of hardening of the product with a thicker cross section. Therefore, the total sensitivity is titelliste alloy tempering is often not so important at a small thickness profile, as in parts with a thicker profile, at least from the point of view of providing strength and durability.

The present invention is primarily aimed at improving the properties of strength-durability aluminum alloy 7XXX series with a thick profile that is larger than a thickness of approximately 1.5 inches (3.8 cm). Low sensitivity to hardening of the alloy, in accordance with the present invention, is of extraordinary importance. For profiles with a large cross-section, the lower the sensitivity to hardening, the easier it is for you to provide material properties to maintain the alloying elements in solid solution (preventing thereby the formation of playing a negative role coarse and another highlight formed during slow cooling from the temperature level SHT), in particular, in the more slowly cooled areas of the planes of the mid and quarter profile specified thick billet. The present invention achieves the desired objective of reducing the sensitivity to hardening through the use of carefully controlled composition of the alloy, which allows hardening thicker profiles, providing an exceptional combination of strength-durability and properties of corrosion resistance.

For illustration of the present invention were formed twenty-eight the ears with a diameter of 11 inches (27.9 cm), obtained by the method of direct casting (DC), which were homogenized and processed by extrusion to obtain a rectangular bars 1.25×4 inch (3,2-10.2 cm). Before hardening all these bars were heat treated in the solid solution at different speeds to simulate the conditions of cooling of thin sections, as well as to create conditions, approximately corresponding to the cross section of the billet thickness of 6 and 8 inches (of 15.2-20.3 cm). These rectangular test bars were then subjected to cold stretching approximately 1.5% to relieve residual stresses. The compositions of the alloys subjected to studies in which the content of zinc was chosen in the range from about 6.0 percent by weight. to a level slightly in excess of 11.0 wt.%, shown in table 2 below. For these test samples, the total content of copper and magnesium were changed in each sample in the range of about from 1.5 to 2.3 wt.%.

For all other alloys, except the control, preset values: Si=0,03, Fe=0,05, Zr=0,12, Ti=0,025.

For control alloy 7150 (sample No. 27) defined values: Si=0,05, Fe=0,10, Zr=0,12, Ti=0,025.

For control alloy 7055 (sample No. 28) defined values: Si=0,07, Fe=0,11, Zr=0,12, Ti=0,025.

Were investigated various approaches to hardening to obtain in-plane sired the us profile extruded bar with a thickness of 1.25 inches (3.2 cm) cooling rate, simulating the velocity in the plane of the middle of the profile plate thickness of 6 inches (15.2 cm), hardened by the water spray at a temperature of 75°F (24°C)that occurs in the case of full-scale production. The second dataset included simulation under identical circumstances, the cooling rate of the bar corresponding to the plate thickness of 8 inches (20.3 cm).

The above simulated tempering included the modification of the characteristics of the heat transfer medium hardening and workpiece surface by immersing the hardened obtained by the method of extrusion bars while using three well-known methods of hardening: (i) tempering at a certain temperature warm water, (ii) saturation of water with carbon dioxide CO2and (iii) chemical treatment of the bars to give a bright etched surface to reduce surface heat transfer.

To simulate the conditions of cooling of the plate thickness of 6 inches (15.2 cm) was performed with the following procedure: the temperature of the water when quenched by immersion was maintained at a level of approximately 180°F (82°C), and the degree of dissolved CO2maintained at a level of approximately 0,20 LAN (a measure of the concentration of dissolved CO2LAN - standard amount of CO2/volume of water). The sample surface was chemically treated to obtain the standard bright baiting is Noah's surface.

To simulate the cooling of a plate thickness of 8 inches (20.3 cm) water temperature raised to about 190°F (88°C) with a corresponding level of dissolved CO2that changed from 0.17 to 0.20 LAN. As for samples with thickness of 6 inches (15.2 cm)above a thicker plate was chemically treated to obtain the standard bright etched surface.

The cooling rate was measured using thermocouples installed in the plane of the middle of the profile of each sample of the bar. To enable comparisons were made two calculated curves cooling, approximating the cooling rate in the plane of the mid profile when quenched by spraying the prefabricated plate thickness 6 and 8 inches (of 15.2-20.3 cm)shown on the attached figure 2. The drawing shows a superimposed two groups of graphs: the lower group (temperature scale) represents the curves simulated cooling rate in the plane of the middle of the profile plate thickness of 6 inches (15.2 cm) and the top - imitation in the plane of the mid profile for slab thickness of 8 inches (of 15.2-20.3 cm). These simulated cooling rate were very close to those for plates of plant production in the important temperature range above about 500°F, although the simulated cooling curves for the experimental materials ex is cialis indicators plates prefabricated at a temperature below 500°F (260°C), what was not considered critical.

After heat treatment in the solid solution and hardening studied the behavior of artificial aging using different periods of aging to obtain acceptable values of specific electrical conductivity (EC) and the values of the corrosion resistance detachment ("EXCB"). The first method is a two-stage aging of the alloy, in accordance with the present invention, consisted of: slow heating (for about 5-6 hours) to a temperature of approximately 250°F (121°C), the period of exposure for 4-6 hours at a temperature of approximately 250°F (121°C), followed by the second stage of aging at a temperature of approximately 310°F (160°C) with a variable period of time ranging from about 4 to 36 hours.

Then collected data to test the resistance to cracking plane strain tensile and pressure of compression on the samples treated with different minimum value of time of aging, are required to obtain visual assessments EXCB at the level of S or better (EA or pitting) for acceptable performance level resistance to corrosion, delamination and minimum values of specific conductance of the EU, exceeding approximately 36% IACS (international annealed copper); the latter value was used to indicate the degree required is electrelane and ensure specific performance enhancing corrosion resistance, as is known in the art. All tensile testing was performed in accordance with specification ASTM E8, and all tests for resistance to cracking under conditions of planar load - in accordance with the specification E ASTM, and these specifications are well known in the art.

Figure 3 shows a graph of the results of strength-resistance for samples of the alloy in table 2, which was slowly tempered from temperature SHT to simulate the hardening of the product thickness of 6 inches (15.2 cm). One family of compounds was significantly different from the others shown in the graph, namely the samples with numbers 1, 6, 11 and 18 (in the upper part of figure 3). All samples with those numbers showed very high resistance to cracking, together with high strength properties. Suddenly, all the compositions of these alloy samples had low levels of copper and low levels of magnesium in the lower range of compositions in accordance with our choice, namely at the level of about 1.5% wt. Mg, 1.5% wt. Cu, while the levels of zinc in this regard ranged from approximately 6.0 to 9.5% of wt. Specific measured values of the level of zinc for these improved alloys amounted to 6% wt. Zn for sample No. 1, 7.6% by weight. Zn for sample No. 6, 8,7% wt. Zn for sample No. 11 and 9.4% wt. Zn for sample No. 18.

Significant street is Ksenia strength and durability, you could also see when the above characteristics of the alloy was compared with two control samples of aluminium alloy 7150 (sample No. 27 above) and aluminium alloy 7055 (sample No. 28); both samples were treated identically (including holidays). Figure 3 by the dashed line connected to the data points for the last two test grades for presentation "trends in properties-strength-resilience", where you can see that the higher strength is accompanied by a low resistance characteristics. It should be noted that the line presented in figure 3 for the control alloy 7150 and 7055, is significantly below the data points for the alloy, in accordance with the present invention, for samples No. 1, 6, 11 and 18, described above.

Graphics, also shown in figure 3 represent the results of alloys containing approximately 1.9% of wt. Mg and 2.0% wt. Cu at different levels of zinc: 6,8% wt. (for sample No. 5), 8.2% of wt. (for sample No. 10), 9.0% of wt. (for sample No. 17) and 10.2% wt. (for sample No. 26). These results again graphically illustrate the decrease in resistance observed for these alloys compared to alloys containing 1.5% by weight. Mg and 1.5% wt. Cu, with appropriate levels of total zinc content. And, although thick profile properties-strength-durability for products from alloys with a higher content of magnesium and copper were anal is Hecny or slightly higher than properties for the control alloy 7150 and 7055 (dotted line trends), these results clearly demonstrate a significant reduction properties of both strength and durability, which occurs at moderate increase in the content of copper and magnesium: (1) higher levels of copper and magnesium alloys in accordance with the present invention and (2) the approximation of the levels of Cu/Mg compositions of many commercially available at present alloys.

A similar set of results are graphically presented on the attached figure 4 for an even slower conditions of hardening than presented and described with reference to the above figure 3. Figure 4 conditions approximately corresponds to the conditions plate thickness of 8 inches (20.3 cm) for cooling conditions in the plane of the mid profile. For the data presented in figure 4, we can make similar conclusions as for figure 3, for the slower simulation of hardening held to represent a thicker product in the form of a plate.

Thus, in contrast to the description of the prior art were obtained best qualities of strength-resistance to some of the lowest levels of copper and magnesium, which thus differed significantly from commercially available at present alloys used in the aerospace industry. In addition, the levels of zinc, in which these properties were the most optimized, adhere to much higher levels than described for the products in the form of a plate of aluminum alloy 7050, 7010 7040 or.

It should be assumed that substantially improved properties of strength and durability, observed for products with a thick cross section of the alloy, in accordance with the present invention, due to the specific combination of ingredients of the alloy. For example, figure 5 presents the values of the yield strength tensile TYS gradually increases with increasing zinc content of sample No. 1 to sample No. 6 and sample No. 11 and exceed the properties of control samples of known prior art. Thus, in contrast to the description of the prior art higher values of dissolved zinc does not necessarily increase the sensitivity to hardening when the alloy was accordingly formed, in accordance with this description. In contrast, higher values of zinc, in accordance with the present invention, in fact, proved the advantage in terms of slow hardening of workpieces with a thick cross-section. However, when further increasing the level of zinc to 9.4% wt. the strength may decrease. In accordance with this, the yield strength tensile TYS sample No. 18 (containing 9,42% wt. Zn) falls lower than in other cases, the CPF is Bob, in accordance with the present invention, with a lower content of zinc, as shown in figure 5.

On the attached 6 also presents the conditions of slow hardening for the simulated thickness of 8 inches (20.3 cm). From these data we can see that the sensitivity to hardening may increase even when the levels of 8.7% wt. Zn, as represented by the values of the yield strength tensile TYS for sample No. 11, which offset lower than that for sample No. 6 with a total zinc content of 7.6% by weight. The influence of the high content of dissolved substances on the sensitivity to hardening is also represented in the accompanying drawings the relative positions of the axes of the yield strength tensile TYS for control alloy 7150 (sample No. 27) and 7055 (sample No. 28). Here, in the slow hardening, the alloy 7055 proved to be more durable than alloy 7150 (figure 5), but the relative scale was reversed at even slower conditions of hardening (in accordance with figure 6).

It should also be noted characteristics for sample No. 7, above, which is in accordance with table 2 contained 1,59% wt. Cu, 2,30% wt. Mg and 7,70% wt. Zn (so that the content of magnesium was higher than the copper content). As shown in figure 3, this sample showed a high value of the yield strength tensile TYS greater than 73 thousand (pounds/inch2) (504 MPa), but considers the super low resistance to cracking KQ(L-T), which constitutes approximately 23 thousand (pounds/inch2)inch1/2. For comparison, the sample No. 6, which contained 7,56% Zn, 1,57% Cu and 1,51% Mg (Mg<Cu)showed, as shown in figure 3, the value of the yield strength tensile TYS in excess of 75 thousand (pounds/inch2) (504 MPa), and a higher value of the resistance to cracking of approximately 34 thousand (pounds/inch2)inch1/2(in fact, increasing the resistance 48%). These comparative data show the importance of: (1) maintain the magnesium content at or below about 1,68 or 1.7% wt., and (2) maintain the specified magnesium content below or equal to the contents of Cu+0,3% wt. and more preferably lower copper content, or at least no higher than the copper content in the alloy in accordance with the present invention.

It is preferable to provide optimal and/or balanced properties resistance to cracking (KQand properties strength (TYS) in alloys in accordance with the present invention. As best seen and understood when comparing the compositions according to table 2 with the corresponding to them the values of the resistance to cracking and durability presented on figure 3, the samples of the alloys within the composition, in accordance with the present invention, provides such a ball the NS properties. In particular, for samples No. 1, 6, 11 and 18, they either have a value (KQ) (L-T) resistance to cracking in excess of approximately 34 thousand (pounds/inch2) inch, when the value TYS exceeding approximately 69 thousand (lb/in1/2) (476 MPa), or they have a value of resistance to cracking in excess of approximately 29 thousand (pounds/inch2)inch1/2in combination with a higher value TYS, constituting approximately 75 thousand (pounds/inch2)inch1/2or higher (518 MPa).

The upper limit of the content of zinc, it is important to achieve an appropriate balance between the properties of durability and strength. Samples, the content of zinc in which more than approximately 11,0% wt., such as samples No. 24 (11,08% wt. Zn) and No. 22 (11,38% wt. Zn), has not ensured that the minimum of the combined levels of strength and resistance to cracking, above, for alloys in accordance with the present invention.

Preferred alloy compositions in accordance with the present invention, therefore, provide a high resistance to damage in thick structural elements designed for the aerospace industry, due to their combined properties, resistance to cracking and yield strength. In respect of certain property values, presents the present description, it should be noted that values of KQrepresents the result of the test of resistance to cracking under plane deformation, which do not meet currently accepted criteria standard E ASTM. In current tests, which were obtained values of KQfor which this description was not provided an exact match criteria of validity are: (1) Pmax/PQ<1,1 basically, and (2) In (thickness) >2,5 (KQYS)2from time to time, where KQ, σYSPmaxand PQdetermined in accordance with ASTM E399-90. These differences are a consequence of the high values of resistance to cracking observed with an alloy in accordance with the present invention. To obtain reliable results for plane strain Klcfor testing, you would need to use a thicker and wider sample than can be achieved with the use of the bar is obtained by extrusion (1.25 inch thick x 4 inches wide (3,2×10.2 cm). Accurate value of Klcgenerally regarded as the property of a material that is relatively independent of the size and geometry of the sample. On the other hand, the value of KQmay not be a true material property in the strict academic sense, because it can change is if I change the size and geometry of the sample. Typical values of KQfor samples smaller than you want, however, are conservative with respect to Klc. In other words, presented in the report, the values of resistance (KQ) to cracking, in General, were lower than the standard values of Klcobtained when the size of the sample met the criteria of reliability standard E-90 ASTM. Values of KQwere obtained in the present description, using sample testing in compression in accordance with the standard E with ASTM thickness equal to 1.25 inches (3.2 cm), and width, which varied from 2.5 to 3.0 inch (6.4 to 7.6 cm) for different samples. Samples that cracked in the fatigue stress, had cracks in length And 1.2 to 1.5 inches (3,1-3,8 cm) (A/W=from 0.45 to 0.5). Test material obtained factory method described below, which really satisfy the criterion of validity of the standard E ASTM for Klc, was performed using samples in compression with thick=2.0 inches (5.1 cm) and a width W=4.0 inches (10.2 cm). These samples were subjected to pre-fatigue cracking to the length of the cracks of 2.0 inches (5.1 cm) (A/W=0,5). All the cases of comparative data among different alloy compositions are presented using the results obtained on samples of the same size and analogues of the conditions of the test.

Information confirming the possibility of carrying out the invention

Example 1: Test plates manufactured in factory conditions

Tests in the factory were conducted using a standard full-sized castings, cast with the following composition of the alloy, in accordance with the present invention: 7,35% wt. Zn, 1,46% wt. Mg, 1,64% wt. Cu, 0.04% of wt. Fe, 0.02% of wt. Si and 0.11 wt.%. Zr. This casting was cleaned from the surface layer, homogenized at a temperature of from 885 to 890°F (474°C-477°C) for 24 hours and subjected to hot rolling to obtain a plate thickness of 6 inches (15.2 cm). Laminated plate was then subjected to heat treatment in the solid solution at a temperature of from 885 to 890°F (474°C-477°C) for 140 minutes, was tempered by spraying to ambient temperature and was out in a cold condition to a value from about 1.5 to 3% for relieving residual stresses. Sections of this plate was subjected to processing by a two-step aging, which consisted of 6 hours of the first stage of aging at 250°F (121°C), followed by the second stage of aging at a temperature of 310°F (160°C) for 6, 8 and 11 hours, respectively, which are indicated as the time periods T1, T2 and T3 in the table below. The test results for tensile strength, resistance to cracking, alternating immersion SCC, EXCB shall conductivity are presented in table 3, below. 7 shows a cross-graph of resistance (Klc) to cracking in the direction of the L-T plane strain conditions depending on the yield strength tensile TYS in the longitudinal direction (L), and two samples were taken in a quarter plane (T/4) profile thickness of the plate. The linear trend of the correlation strength-the resistance (line T3-T2-T1) was drawn so that it was determined among the presented data, the second aging. The preferred line (M-M) minimum characteristic values were also plotted in the drawing. Figure 7 also presents typical properties for plate thickness of 6 inches (15.2 cm) of alloy 7050-T received in accordance with the industry specification BMS 7-323C, and typical values for alloy 7040-T for plate thickness of 6 inches (15.2 cm) in accordance with the project specifications AMS D99AA (see Preliminary Materials Properties Handbook; both descriptions are known in the art. From these preliminary data related to the plate, subjected to aging in two phases, the alloy compositions in accordance with the present invention, clearly show much the best combination of strength-resistance in comparison with plates of alloys 7050 7040 or. Compared with plate 7050-T option, for instance, in accordance with the present invention, the resulting two-stage aging, the chalk superior value TYS approximately 11% (72 thousand (pounds/inch 2) (497 MPa) against 64 thousand (pounds/inch2) (442 MPa), while the equivalent value of Klcequal to 35 thousand (pounds/inch2)inch1/2. In other words, there was obtained a significant increase in the values of Klcin accordance with the present invention, at equivalent levels TYS. For example, in the version of the aging in two stages of this product in the form of a plate resistance values of Klc(L-T) reached 28% increase (32,3 thousand (pounds/inch2)inch1/2compared with 41 thousand (pounds/inch2)inch1/2) when compared with equivalent alloy 7040-T with the same level TYS (L) - 66,6 thousand (pounds/inch2) (560 MPa).

Example 2: Testing at the factory forging

Tests forging and stamping sample of the alloy, in accordance with the present invention was conducted during testing at the factory using a full-sized castings production in the form of sheet/plate, marked SOUR and SOUR, with the following composition:

The OMRS 1: 7,35% wt. Zn, 1,46% wt. Mg, 1,64% wt. Cu, 0,11% wt. Zr, 0,038% wt. Fe, 0,022% wt. Si, 0.02% of wt.

The OMRS 2: 7,39% wt. Zn, 1,48% wt. Mg, 1.91 per cent wt. Cu, 0,11% wt. Zr, being 0.036 wt.%. Fe, 0,024% wt. Si, 0.02% of wt.

Standard casting alloy 7050 was also used as a control sample. All of the above castings were homogenized is at a temperature of 885°F (474°C) for 24 hours and cut into dies for forging. Part in the form of stamped forgings was obtained to evaluate properties at three different values of thickness: 2 inches, 3 inches and 7 inches (5,1, 7,6, 17.8 cm). The processing steps carried out in respect of these metals included: two operations preliminary molding using a hand-forged, followed by the operation of blister forming and operation of the final stamping using a press 35000 tons. Used forging temperature, therefore, was within about 725-750°F (385°C-399°C). All forged parts were then subjected to heat treatment in the solid solution at temperatures of from 880 to 890°F (471°C-477°C) for 6 hours, was tempered and subjected to cold machined with stretching 1-5% to relieve residual stresses. The details are then processed by aging annealing type T to enhance the performance of SCC. Processing by aging consisted of processing at a temperature of 225°F (107°C) for 8 hours, followed by treatment at a temperature of 250°F (121°C) for 8 hours, then at a temperature of 350°F (177°C) for 8 hours. Test results of tensile strength, made in longitudinal, longitudinal-transverse direction and in the short transverse direction, presented at the applied Fig. In all three directions of orientation values of the yield strength tensile (TYS) for the lava, in accordance with the present invention, remained almost unchanged for values of thickness in the range from 2 to 7 inches (5.1 to 17.8 cm). In contrast, the specification for 7050 allowed to fall TYS values with increasing thickness from 2 to 3 and up to 7 inches (5.1 to 7.6 for-17.8 cm) in accordance with known characteristics of the alloy 7050. Thus, the results presented in Fig, clearly demonstrate the advantage of the present invention in relation to a low sensitivity to hardening or, in other words, the ability of forgings made from this alloy to be insensitive to changes in strength in a wide range of thicknesses in contrast to the observed fall of the properties of the comparative strength of the samples when more thick cross-section forgings of alloy 7050 prior art.

The present invention obviously does not fit into the conventional design concept using alloys of the 7XXX series, in which it is believed that the high content of magnesium is preferred to provide a high level of strength. Although this approach may be valid for 7XXX aluminum alloy with a thin cross-section, it does not match the forms of the product with a thicker profile, because the higher magnesium content actually increases the sensitivity to hardening and snige the strength in thick section.

Although the present invention is mainly directed on the product with a thick cross-section, the hardening of which is as fast as it is practically acceptable, for specialists in the art will understand that in other types of applications can use the advantages of the present invention in relation to the low level of sensitivity to hardening, and can be used deliberately slow hardening of parts with thin cross-section to reduce the residual stresses induced by quenching, and the amount/degree of distortion caused by a rapid hardening, without undue loss of strength or durability.

Other possible applications arising from such low sensitivity to quenching observed when using alloy, in accordance with the present invention, is a products having thick and thin sections, such as forging-stamping and some of the details obtained by extrusion. In such products must be provided with a lesser level differences of the values of yield strength between areas of thick and thin cross-section. This in turn should reduce the likelihood of bending or breaking the dimensions after stretching.

Usually for any C is given alloy 7XXX series as a serial advanced processing during artificial aging after reaching the peak strength of the product, released in the mode type T6 (i.e. treated with purestream"), the strength of this product, as it is known, consistently and systematically decreases as consistent and systematic growth, resistance to cracking and corrosion resistance. Thus, the modern designers items have learned to choose certain conditions leave, to achieve desired for specific applications, a compromise combination of strength, resistance to cracking and corrosion resistance. Undoubtedly, the same applies to the alloy, in accordance with the present invention, which illustrates a cross-schedule Kiccracking under plane deformation in the direction of the L-T and the yield strength in tension in the direction L 7, both values were measured in a quarter plane (T/4) thickness profile in the longitudinal direction of the product in the form of a plate thickness of 6 inches (15.2 cm). 7 shows that the alloy in accordance with the present invention, provides the following combination of properties: approximately 75 thousand (pounds/inch2) (518 MPa) yield strength tensile at approximately 33 thousand (pounds/inch2)inch1/2tensile cracking at time T1 of aging on table 3 or approximately 72 thousand (pounds/inch2) (497 MPa) yield strength with age is position at approximately 35 thousand (pounds/inch 2)inch1/2resistance to cracking at time T2 ageing in table 3, or approximately 67 thousand (pounds/inch2) (462 MPa) yield strength tensile at approximately 40 thousand (pounds/inch2)inch1/2resistance to cracking at the time of aging T3 in table 3.

For experts in the field of technology, in addition, it will be clear that within the constraints of the specific alloy series 7XXX line trends durability-resistance to cracking can be interpolated and, to some extent, extrapolated for the given combination of strength and resistance to cracking beyond the characteristics of the three examples of the alloy, in accordance with the present invention, the above and presented in Fig.7. The desired combination of many properties can be thus obtained by selecting the appropriate processing artificial aging.

Although the present invention has been mainly described in relation to options for use as structural components in the aerospace industry, it should be understood that variants of the end use of this product are not necessarily limited to this industry. On the contrary, the alloy in accordance with the present invention and the manner of its preferred treetops is ageing, likely to have many other options for end use not related to the aerospace industry, in the form of a relatively thick cast products, rolled products in the form of plates, extruded or forged products, especially in application forms, which may require the provision of a relatively high strength under conditions of slow quenching from temperatures SHT. An example of one of such applications is the plate form different configurations for the manufacture of which you want to apply significant mechanical processing used to create the form and/or profiling in various other industrial processes. In such embodiments, the application required that the material, while providing the features of high strength and low distortion during machining. When using 7XXX alloys for the manufacture of plate forms would need to use the slow quenching after heat treatment in the solid solution to obtain a low value of the residual voltage, which, otherwise, would lead to distortion during machining. Slow hardening also leads to a decrease in strength and other properties of existing 7XXX series alloys because of their higher sensitivity the spine of tempering. Unique, very low sensitivity to hardening of the alloys in accordance with the present invention, allows a slow hardening from SHT, while maintaining a relatively high strength properties, which allows the use of this alloy in such non structural options for use that is not associated with the aerospace industry, as a thick plate shape. For this particular application it is not necessary to perform the preferred method 3-step aging described above. Should be sufficient even processing in a single step or using a standard operation 2-stage aging. Plate form can even be made of a product in the form of a cast slab.

In the present invention essentially eliminates the problems associated with the prior art, when using a family of products aluminum alloy 7000 series which exhibit substantially reduced sensitivity to hardening, providing, thus, significantly higher levels of strength and resistance to cracking than was possible hitherto in structural aerospace parts with a thick profile or parts obtained by the method of computer processing of thick products. It also describes how aging, to improve the properties stay the spine to corrosion these new alloys. Measurement of yield strength tensile (TYS) and the conductivity of the EU (as % of IACS) was performed on representative samples of several new compositions of alloy 7XXX and by comparative methods of aging, in accordance with the present invention. The above dimension of the EU, probably correlate with the actual properties of resistance to corrosion so that the higher the measured value of the EU, the greater the resistance to corrosion shall be alloy. As an illustration, commercial alloy 7050 get with three modes vacation, improving the resistance to corrosion: kzt76 (with a typical minimum value of SCC or "guarantee" of approximately 25 thousand (pounds/inch2) (173 MPa) and a typical EU, equal to 39.5% of IACS); T74 (with typical guaranteed SCC value of approximately 35 thousand (pounds/inch2) (242 MPa) and the value of the EU 40,5% of IACS) and T73 (with typical guaranteed value SCC approximately 45 thousand (pounds/inch2) (311 MPa) and the value of the EU 41,5% of IACS).

In the aerospace industry, marine industry or other structural variants use quite often, engineer, structural elements and materials chooses the materials for a specific element from the calculation of the failure mode weakest link. For example, since the alloy for the upper part of the wing is mainly exposed to load the m in compression, for him to have relatively low requirements of SCC resistance, including the load in tension. While the alloys covering the upper part of the wing and their modes vacation is usually chosen so that they provide a higher strength, albeit at relatively low resistance of SCC in the short transverse direction. Within one compartment of the wing aerospace apparatus elements of the spar are efforts stretching. Although the design engineer would like to use for this option, use materials with higher strength in order to reduce the weight of the item, the weakest link requires high SCC resistance for these parts of the elements. Modern details of the spar, thus, traditionally made from a more corrosion-resistant, but with a lower strength alloy, such as alloy, processed in vacation mode T74. Based on the observed increase of the EU with the same strength and test results AI SCC described above, preferred new ways 3-step aging, in accordance with the present invention, allow design engineers/engineers materials science and constructors parts aerospace devices the ability to achieve levels of strength products 7050/7010/7040-kzt76 with levels of resistance to corrosi is, close to get in vacation mode T. Alternatively, the present invention can provide resistance to corrosion of the material released by the mode kzt76, in combination with significantly higher levels of strength.

Examples

Three representative product family new 7XXX alloys were cast in the form of a given size castings commercial sample with the following structure:

These materials castings, of course, after machining, i.e. rolling, to obtain a plate with a profile thickness of 6 inches (15.2 cm), heat treatment of solid solution and so on, were subjected to comparative methods of aging and change, as presented in table 5 below. In fact, if such a 3-stage evaluation compared two different first stage, one was a single exposure at a temperature of 250°F (121°C), the other was divided into two padata: 4 hours at 225°F (107°C), followed by the second step of 6 hours at 250°F (121°C). These two procedures substages are referred to here as the first processing of the first stage, that is carried out before processing of the second stage at a temperature of approximately 310°F (154°C). In any case, when processing in the course of these two "types" of the first stages - single treatment at 250°F (121°C) compared with the divided clicks the processing at steps 225 and 250°F (107 and 121°C) - it was not observed significant differences of the properties. Therefore, the reference to any such step in this description covers all such options.

Samples of each plate thickness of six inches (15.2 cm) were challenged with obtaining average values of the properties for two-stage and three-stage aging, which were measured the following values:

Figure 9 shows a graph for comparing the values of the yield strength in tension and values of the EU, which was used to obtain interpolated data presented in table 6 above. In particular, as noted above, a significant increase in EC was observed for the above alloys a, b or C, the last 3-step aging at the same level of yield strength. From these data it was also noted that there has been a surprising and significant increase in strength at the same level of the EU for the above conditions 3-step aging compared with 2 stages, the second stage which is performed at a temperature of approximately 310°F (154°C). For example, the yield strength of the alloy specimen And held 2-stage aging, with 39.5% of IACS made up 72.1 thousand (pounds/inch2) (497 MPa). But it is TYS increased to 75,4 thousand (pounds/inch2) (520 MPa) for a given 3-stage aging is accordance with the present invention.

Research AI SCC was performed in accordance with standard D-1141 ASTM by alternating immersions in the specified solution of the synthetic ocean water (or SOW), which is more aggressive than the typical 3.5% NaCl solution required in accordance with standard ASTM G44. Table 7 shows the results for different samples of alloys a, b and C (all in the direction of ST) when using only 2-stage aging, and the second stage has held various periods of time (6, 8, and 11 hours) at a temperature of approximately 310°F (160°C).

It is interesting to note that there were no failures of samples under identical testing conditions after the first 93 days of exposure. Thus, a new method of 3-stage aging, in accordance with the present invention, probably gives the unique advantages of strength/SCC exceeding the parameters that can be achieved using conventional 2-stage aging, promising the best properties in new products and providing additional improved combination of properties in other presently used lines of products for the aerospace industry.

When comparing table 7 with table 8 we can see that though aging using 2 phases/stages can be used in order to find for the alloy, in accordance with the present invention, the preferred operation 3-stage aging, described in the present application is, in fact, gives a measurable improvement in properties when tested SCC. Tables 6 and 7 also contain data "indicator" operating characteristics of SCC values of the EU (such as % of IACS) together with the respective measured values TYS (T/4). These data cannot be directly compared to determine the relative values of products that have aging in two stages, with products that undergo aging in three stages, however, because the test of the EU have been performed in different areas of the product, the table 7 values were used for surface measurements, and table 8 measure-level T/10 (it is known that the indicator values of the EU are usually reduced when measuring from the surface inward of a given test sample). Values TYS cannot be used for accurate comparison due to significant variations in size, and also due to the dependence on the place of the tests (laboratory tests and tests at the factory). Instead, they should use relative data in figure 9 (below) to compare the extent to which a 3-stage aging provides improved combinations of strength and corrosion resistance using values TYS (thousand pounds/inch2) the longitudinal direction compared with the conductivity of the EU (% of IACS) for direct comparison of test specimens plate thickness of 6 inches (15.2 cm) of the alloys in accordance with the present invention.

Test data SCC in terms of the marine atmosphere confirm a significant increase in corrosion resistance, implemented using a new three-step aging in respect of the above new family of alloys 7XXX. For the composition of the alloy, designated as alloy a in the above table 4, the SCC tests were conducted during the period of 568 days when using 2-stage aging compared with a probation period of 328 days for a 3-stage aging, when compared with data for 2-stage ageing and 3-stage aging characteristics of the SCC noted in the following table 9 (last (3-stage), the tests were started after the beginning of the previous (2-pass) tests; therefore, for samples that passed two-stage aging was observed over long periods of time trials).

Note: the 2-step method of ageing include: 6 hours at 250°F (121°C);

6 or 8 hours at 310°F (160°C);

3-stage method of ageing include: 6 hours at 250°F (121°C);

7 or 9 hours at 310°F (160°C)for 24 hours at 250°F (121°C).

These data are represented graphically in the attached figure 10 with time values shown in this drawing in the upper left of the pointer, which always refers to the time of the second stage of aging at 310°F (160°C), even for samples, p is Osadchy 3-step aging, which is usually done here is the link.

The second composition, the alloy according to table 4 (containing 7.4% of wt. Zn, 1.5% wt. Mg, 1.9 percent by weight. Cu and 0.11 wt.%. Zr) were subjected to comparative tests of 2-stage ageing and 3-stage aging, as for the alloy As described above. Long-term results obtained when testing the SCC in terms of the marine atmosphere, is presented in table 10, below.

Graphically the data of table 10 presents to the applied 11 with time values in the upper left pointer to this Fig., which always refers to the time of the second stage of aging at a temperature of 310°F (160°C), even for samples that have passed 3 stages of aging, which is usually done here link. From the data for both alloys a and C can be seen that when using the preferred 3-step aging, in accordance with the present invention in relation to preferred alloy compositions are obtained significant improvement in characteristics of the alloy when tested SCC in terms of the marine atmosphere, in particular, when the number of days to failure for materials that have passed a 3-stage aging directly compared with samples that undergo a 2-stage aging. However, long before these tests SCC in the Maritime environment, the materials, the past 2-stage aging is s, showed some improvement in the operating characteristics of SCC in simulated tests, and can be used in some embodiments of the application, in accordance with the present invention, even though the improved operation of aging 3 phase/stage is preferred.

As for the 3-stage aging, which is preferred for the above alloy compositions, it can be noted that the first stage of aging should preferably take place in the temperature range from about 200 to 275°F (93-135°C), more preferably from about 225 or 230 to 260°F (107 or 110-127V°C) and most preferably at about 250°F (121°C). Although aging for about 6 hours at the above temperature, or the temperature is sufficient, it should be noted that in any broad sense of the duration of time required for the first stage of aging, should be sufficient to ensure a sufficient degree of dispersion hardening. Thus, a relatively short exposure, for example, approximately at level 2 or 3 hours at a temperature of approximately 250°F (121°C) may be sufficient (1) depending on size, detail and complexity of the form and (2)in particular, when the above "brief" period of treatment/exposure use in conjunction with relatively slow the heating rate within a few hours, for example 4-6 or 7 hours in total.

The preferred second stage of aging in relation to one of the preferred compositions of the alloy, in accordance with the present invention, can be intentionally carried out directly after the first of the above stage of heat treatment, or may be purposefully used an explicit break time/temperature between the first and second stages of processing. In a broad sense, the second stage should take place within a temperature range from about 290 or 300 to 330 or 335°F (143 or 149-166 or 168°C). Preferably, the second stage of aging is performed within a temperature range from about 305 to 325°F (152-163°C). Preferably, the second stage of aging is carried out from about 310 to 320 or 325°F (154-160 or 163°C). The preferred time for this critical process the second stage, to some extent, inversely depends on the current temperature (temperature)used for processing. For example, if the processing is carried out essentially at a temperature very close to the value of 310°F (154°C), the total time will be approximately 6 to 18 hours, preferably, from about 7 to 13 hours or may even be enough for 15 hours. More preferably, the second stage of aging can be carried out during the total period of treatment around the about 10 or 11 and even 13 hours at the operating temperature. When the temperature of the second stage of aging approximately 310°F (160°C) the total time of the second stage can be from about 6 to 10 hours, preferably, from about 7 or 8 to 10 or 11 hours. When choosing the time and temperature of the second stage of aging must also be considered a preferred aspect of the specified property. More specifically, the shorter the processing time at a given temperature results in higher strength values, while a longer time helps to obtain the best properties of corrosion resistance.

Finally, in relation to the preferred third stage of aging, it is better not to carry out a gradual decrease in temperature after the second stage to perform the necessary third stage on such heavy workpieces, unless paid exceptional attention to ensure its coordination with the temperature of the second stage, taking into account the total processing time that is required to prevent excessive heating at the temperature of the second stage of aging. Between the second and third stages of aging metal products, in accordance with the present invention, can be intentionally removed from the heating oven and subjected to a rapid cooling using a fan or similar device to a temperature priblizitelen is 250°F (121°C) or below, maybe even with a full cooling to room temperature. In any case, the preferred periods of the exposure time/temperature for the third stage of aging, in accordance with the present invention, so expect them to be similar to the parameters of the first stage of aging described above.

The alloy, in accordance with the present invention, preferably made in the form of a product, in particular of the product obtained in the form of castings, suitable for hot rolling. For example, large castings can be cast using semi-continuous method using the above composition, and can then be removing the surface layer or mechanical treatment to remove surface defects that it is necessary or required to ensure a good surface rolling. The casting can then be pre-heated for homogenization and dissolution of the internal structure, while the corresponding processing preheating is heated to relatively high for this type of formulations temperature below 900°F (482°C). While it is preferable to carry out heating to a first, lower temperature, for example, to temperatures above 800°F (427°C), for example, approximately 820°F (434°C) or higher, or 850°F (454°C) or higher, preferably 860°F (460°C) or above, for example, CA is approximately up to 870°F (466°C) or higher, and to withstand the casting approximately at this temperature or temperatures for a substantial period of time, say 3 or 4 hours. Then the casting is heated to a temperature of about 890°F or 900°F (477°C or 482°C) or possibly higher during another time of ageing, which is a few hours. Such a stepped or gradual heating promotes homogenization, as is known in the art for many years. Preferably, the homogenization was carried out during the total exposure time of about 4-20 hours or more, and the temperature of homogenization may exceed approximately 880-890°F (471-477°C). That is, the total time of exposure at temperatures in excess of about 890°F. (477°C)should be at least 4 hours and preferably should be greater than, for example, 8 to 20 or 24 hours or more. As you know, the large size castings and other parameters may be required to use a longer time of homogenization. Preferably, the total volume percentage of insoluble and soluble components are maintained at a low level, for example, not exceeding 1,5% vol., preferably not higher than 1% vol. The use described here is relatively significant pre-heating or homogenization and high temperature heat treatment of solid Rast is ora contributes to this condition, although high temperatures require careful control to prevent partial melt. Such precautions can relate to the need for careful control in the process of heating, including slow or speed up or both.

The casting is then subjected to hot rolling, and, preferably, it is required to provide precrystallization the structure of the grains in the rolled product in the form of a plate. Therefore, casting to hot rolling may come out of the oven at a temperature of essentially greater than about 820°F. (438°C), for example, from about 840 to 850°F (449°C-454°C) or, perhaps, above, and the operation of the rolling is carried out at initial temperatures above 775°F (413°C) or better yet above 800°F (427°C), for example, approximately 810 or even 825°F (432°C or 441°C). This reduces the likelihood of recrystallization, and is preferred in some situations to conduct rolling without surgery re-heat by using power rolling mill and heat preservation during rolling to maintain the temperature of the rolling above the required minimum, for example, 750°F (399°C) or such order. Usually when performing the present invention, it is preferable to provide maximum recrystallization at about 50% or less, preferably in the Aries approximately 35% or less and most preferably not more than approximately 25% of recrystallization it should be understood that the smaller will be the resulting level of recrystallization, the better will be provided with the properties of resistance to cracking.

Hot rolling usually continue in reversing mills for hot rolling as long until you get the desired thickness. In accordance with the present invention the product in the form of a plate designed for machining, fabrication of structural elements of the aircraft, such as one-piece side members, may have a thickness of approximately 2-3 inches (5.1 to 7.6 cm) up to about 9 or 10 inches (22,3-25.4 cm) or more. Typically, this plate has a thickness of approximately 4 inches (10.2 cm) for a relatively small aircraft to thicker plates constituting from about 6 or 8 inches (of 15.2-20.3 cm) to about 10 or 12 inches (for 25.4-30.5 cm) or more. In addition to the preferred embodiments, it is assumed that the present invention can be used for the manufacture of the lower skin of the wing of small commercial jets. Other applications can include details obtained by forging and extrusion, in particular varieties of such parts with a thick profile. When performing extrusion alloy, in accordance with the present invention, is extruded in the temperature range from about 600 to 750°F (316°C-399°C), such as the er, approximately 700°F (371°C) and preferably produce a reduced cross-sectional area (extrusion) at the level of approximately 10:1 or more. Forging can also be used in the present invention.

Hot rolled plate or other wrought product is subjected to heat treatment in the solid solution (SHT) by heating in the range from approximately 840 or 850°F (445°C-454°C) up to 880 or 900°F (471-482°C) for transfer to the solution of significant parts, preferably all or essentially all parts of zinc, magnesium and copper, soluble at a temperature of SHT, it should be understood that the physical processes are not always perfect, and probably the remains of these major alloying ingredients may not be completely dissolved in during SHT (dissolution). After heating to elevated temperature, as described above, the product must be hardened to complete the procedure of heat treatment in the solid solution. Such cooling is typically performed either by immersion in a tank of suitable size with cold water or by spraying water, although the cooling air can also be used as an additional or replacement cooling means for conducting a cooling. After hardening certain products must be processed in a cold state, for example, the R, by stretching or compression in order to relieve stress or to straighten the product, and even, perhaps, in some cases, to further enhance the strength of the product in the form of a plate. For example, the plate can be stretched or compressed to 1 or 1.5 or maybe 2% or 3% or more, or may be subjected to cold working of another species, usually an equivalent degree. After heat treatment in the solid solution (and hardening) of the product with cold or without processing, it is considered that the product is in a state where it can be subjected to dispersion hardening or ready for artificial aging in accordance with the preferred methods of artificial aging, described herein, or using other methods of artificial aging. Used herein, the term "heat treatment in the solid solution", unless specified otherwise, includes hardening.

After hardening and cold working, if necessary, the product (which may be a product in the form of a plate) subjected to artificial aging by heating to the appropriate temperature to improve the strength and other properties. In one of the preferred variants of heat treatment aging product in the form of a plate of an alloy, which can be subjected to dispersionnom hardening, processed using the three main stages of aging, the process described above, although there might be no distinct boundaries between each stage or phase. It is well known that raising and/or lowering the temperature to the desired or target temperature of the processing itself can create the effect of hardening (aging), which can and often should be considered a joint consideration of such conditions increase or decrease the temperature and their influence on the dispersion hardening when calculating the total duration of treatment with aging.

Aggregate processing should also be taken into account when calculating method of aging in accordance with the present invention. For example, the temperature in a programmable air furnace after completion of the first stage heat treatment at a temperature of 250°F (121°C) within 24 hours may gradually consistently rise to the level of approximately 310°F (154°C) or around this value for the duration of time, and, even if it is not used the proper time, after this period the temperature rises, the metal can then be immediately transferred to another oven, pre-stable at a temperature of 250°F (121°C), followed by exposure there for 6-24 hours. This largely continuous mode start the I does not contain the return to room temperature between treatments at the transitions from the first to the second and from the second to the third stages of aging. This integration of ageing stages are described in more detail in U.S. patent US 3645804, the full content of which is given here as a reference. With increasing and decreasing temperature with appropriate comprehensive account of the stages of this method in one programmable furnace can run two or, less preferably, may three phases artificial aging product in the form of a plate. However, for convenience and ease of understanding, the preferred embodiments of the present invention have been described in more detail, as if each phase, stage or phase was carried out separately from the other two stages of artificial aging is carried out for details. Generally speaking, the first of these three stages or phases, as is intended for precipitation hardening of the corresponding product of the alloy; the second (held at a higher temperature) phase of the alloy, in accordance with the present invention, is exposed to higher temperatures to improve its corrosion resistance, particularly resistance to cracking due to corrosion under stress (SCC) as under normal conditions and in industrial conditions, the simulation of the marine atmosphere. The third and final stage then provide additional dispersion hardening alloy, in accordance with this is their invention, obtaining a high level of strength, giving it additional properties superior corrosion resistance.

Low sensitivity to hardening of the alloy, in accordance with the present invention, allows specialists in the art to find potential application in class processes, in General, described as "hardening under pressure". You can illustrate "the way" hardening under pressure when considering the standard production flow extrudable alloy subjected to hardening by aging, such as an alloy belonging to the series alloys 2XXX, 6XXX, 7XXX or 8XXX. Typical mass production contains: direct casting (DC) dies, molds, homogenization, cooling to ambient temperature, reheating to a temperature of extrusion using furnaces or induction heaters, heated extrusion dies for giving the final shape, cooling the extruded part to ambient temperature, thermal processing parts in the solid solution hardening, stretching, and natural aging at room temperature or artificial ageing at elevated temperatures up to the end of the vacation mode. The process of "hardening under pressure" contains the control of the temperature of extrusion and other conditions of the extrusion, so after exiting the extrusion head part has a heating temperature of the solid solution or close to the desired temperature, and soluble components effectively turn into solid solution. As the detail emerges from the extrusion press it then immediately and continuously directly quenched with water, compressed air or other environment. Item held hardening under pressure, then treated in the usual stretching, followed by natural or artificial aging. As a result, compared with conventional continuous production, expensive separate heat treatment process of solid solution is removed from this version of hardening under pressure, which substantially reduced production costs and energy consumption.

For most alloys, in particular alloys related to the relatively sensitive to the hardening of the 7XXX series alloys, hardening by quenching under pressure usually is not as effective compared to hardening in the course of thermal treatment of solid solution, so as a result of application of such hardening under pressure may experience a significant deterioration of certain material attributes, such as durability, resistance to development the cracks, corrosion resistance and other properties. Because the alloy, in accordance with the present invention, has a very low sensitivity to hardening, it is expected that the deterioration of the properties of the alloy in the course of hardening under pressure will be either eliminated or substantially reduced to acceptable levels for many applications.

For variants of the molding plate, in accordance with the present invention, the resistance to SCC is not so critical, so that these compounds can also be used in known ways to handle a two-step artificial aging instead of the preferred three-step aging, described in the present description.

The reference to the minimum value (for example, the value of the strength or resistance) may relate to the level specified in specifications for the purchase or designation of the materials, or to the level that can be guaranteed for the material, or to the level that the manufacturer of the frame of the aircraft (safety factor) can be taken into account in their design decisions. In some cases, you can use a statistical approach, in which 99% of the product correspond to or are expected to meet at 95% level of confidence, using standard statistical methods. Due to the insufficient amount of data Nelly who cannot make statistically accurate reference to certain minimum or maximum values, in accordance with the present invention, as really "guaranteed" values. In these cases, calculations were made on the basis of currently available data to extrapolate the values (e.g., maxima and minima). See, for example, extrapolated currently, the minimum value of S/N shown in the graph for the slab (solid line a-a in Fig) and forging (solid line b-b In Fig), and extrapolated to the present time, the maximum value FCG (solid line C-C in Fig).

Resistance to cracking is an important property of the developer frame structure of the aircraft, especially when good durability can be combined with good strength. For comparison, the yield strength tensile or ability to withstand the load without cracking structural element can be defined as the load divided by the area of the smallest cross-section of the element in the plane perpendicular to the tensile load (the value of the net load per unit cross-sectional area). For simple shapes with straight sides, the strength of the section can be easily linked to the tensile strength or yield strength tensile sample for testing with a gradual increase in load. Thus, conduct the test limit fluid is ti tensile. However, in the part containing the crack or fissure defect, the strength of the structural element depends on the crack length, the geometry of the structural element and material properties, known as the resistance to cracking. Resistance to cracking can be represented as the resistance of a material detrimental or even catastrophic propagation of cracks under load.

Resistance to cracking can be measured in several ways. One of them consists in the application of tensile load to the test specimen with a crack. The load required for cracking of the test sample divided by the net cross-sectional area (the cross-sectional area minus the area of the cross section containing the crack), known as the residual resistance, expressed in units of thousands of pounds of force per unit area (thousand pounds per square inch). When the strength of the material and the geometry of the sample are constant, the residual resistance is a measure of the resistance to cracking of the material. Because this option to the extent depends on the strength and geometry of the sample, the residual resistance is usually used as a measure of the resistance to cracking when other methods are not as practical as it is desirable, because the poison restrictions associated with the size or shape of the available material.

When the geometry of the structural element is such that it cannot plastically deform in thickness during the application of tensile loads (plane strain), the resistance to cracking is often measured as resistance to cracking under conditions of plane strain, Klc. It usually refers to a relatively thick products or sections with a thickness of, for example, preferably 0.6 or 0.8 or 1 inch (2-2 .5 cm) or more. The ASTM has established a standard test conditions using a sample voltage compression with pre-fatigue cracking for measuring values of Klcwhich is expressed in units thousand (pounds/inch2)inch1/2. These tests are usually used to measure the resistance to cracking of thick material, because it is believed that this property does not depend on the sample geometry, if only satisfied the standards for width, crack length and thickness. The symbol To be used in the designation of index Kicis called the intensity of the load.

Structural elements that are deformed under plane deformation, are relatively thick, as described above. More fine structural elements (less than 0.8 to 1 inch (2-2 .5 cm) in thickness is) usually deformed under plane stress state, or more frequently when under mixed mode loading. The measurement of the resistance to cracking under such conditions may cause variations in the results, since the values obtained from these tests, to a certain extent depend on the geometry of the test specimen. One way of testing is the application of an increasing load on a rectangular test specimen containing a crack. Thus can be obtained a graph of intensity depending on the increase of cracks, known as the R-curve (curve of resistance to fracture). The load value with a certain degree of increase of the crack based on the secant shifted by 25% on the curve of the load and the magnitude of cracks, and the effective crack length when the load is used to calculate the measure of resistance to cracking, known as the KR25. The 20% offset clipping this index is denoted as KR20. He also is expressed in units thousand (pounds/inch2)inch1/2. The well-known ASTM E561 refers to the definition of the R-curve and, as such, is well recognized in the art.

When the geometry of the product from alloy or structural element is such that it is possible to perform plastic deformation on its thickness during the application of tensile load, resistance to cracking h is a hundred is measured as resistance to cracking under plane stress state, which can be determined using tests on the Central stretch of the crack. As a measure of the resistance to cracking use the maximum load obtained in a relatively thin sample with a wide pre-formed crack. When the crack length at maximum load is used to calculate the index of intensity of load, this load value the index of the intensity of load is called cracking resistance in a plane stress state Kc. However, when the intensity of the load is calculated using the length of the cracks before application of the load, the result of the calculation is known as the apparent resistance to cracking Kappmaterial. Because the calculation of Kcusually use a long crack, the values of Kcare usually higher than the values of Kappfor this material. Both of these criterion of resistance to cracking are expressed in thousand units (pounds/inch2)inch1/2. For materials with high resistance digital value obtained as a result of these tests, usually increase with increasing width of the sample or the reduction of its thickness, as is known in the art. Unless specified otherwise in this specification, the values in the flat voltage the nom condition (K c)given here refer to the test panel width 16 " (40,6 cm). Specialists in the art will recognize that the results of these tests can vary depending on the width of the test panel and it is intended to cover all such tests when referring to the parameters of durability. Therefore, resistance, essentially equivalent or essentially corresponding to the minimum value of Kcor Kappwhen the characteristics of products in accordance with the present invention, in a broad sense when referring to tests using panel width 16 " (40,6 cm), are intended to encompass variations of values of Kcor Kappbenefits of using panels with different width that is understandable for specialists in this field of technology.

The temperature at which carry out the measurement of the resistance can be significant. When flying at high altitude the temperature is low enough, for example minus 65°F (-54°C), and for new projects commercial jets resistance at minus 65°F (-54°C) is a significant factor, this requires that the material in the lower part of the wing showed a level of resistance Klcapproximately 45 thousand (pounds/inch2)inch1/2at a temperature of minus 65°F (-54°C)for KR20- at the level of the e 95 thousand (pounds/inch 2)inch1/2preferably 100 thousand (pounds/inch2)inch1/2or higher. Due to such high values of resistance, it becomes possible to use the details of the lower part of the wing is made of such alloys, instead of the currently used structural parts made from alloy series 2000 or series 2XXX) to the detriment of the corresponding property (i.e. strength/durability). Due to the use in practice of the present invention, it becomes possible to manufacture the upper wing skin of the same alloy, using only his or in combination with integrally formed elements such as stiffeners, ribs and stringers.

Resistance superior products, in accordance with the present invention is very high and in some cases may allow aircraft designers to focus on the durability of the material and its resistance to damage in order to emphasize the fatigue resistance and resistance to cracking. A high value of resistance to cracking under the action of fatigue is a very desirable property. Fatigue cracking occurs as a result of repeated cycles of loading and unloading or cycles between high and low load when, for example, the wing moves up and down. T is Kai cyclic loading may occur during flight due to wind gusts or other sudden changes in air pressure or on the ground when towing aircraft. Fatigue failures are a significant percentage of failures of structural elements of the aircraft. Such fatigue failures are insidious because they may occur under normal operating conditions, without excessive overloads without notice. The development of cracks is accelerated due to inhomogeneities of the material, which act as the starting point or place, contributing to the connection of smaller cracks. Therefore, the method of processing or alteration of the composition, improve the quality of the metal by changing the stiffness or harmful levels, improve fatigue resistance.

Tests for fatigue cycle time to fatigue failure under load (S-N or S/N) characterize the resistance of a material to initiate fatigue failure and the growth of small cracks, which form a major part of the total operating time to fatigue fracture. Therefore, the improvement of the fatigue properties of S-N can provide the work item at higher loads during the design life or work at the same load for extended service life. The latter figure can be converted to significant weight savings by reducing the size or reducing the cost of manufacture of the item, or because of structural simplification, while the latter can translate into fewer floor is surface maintenance at low cost maintenance. Load during fatigue testing below the static tensile strength or yield strength in tension of the material, as measured by tests on the yield stress tensile, and usually are below the yield strength of the material. Test initiation fatigue fracture is an important indicator for buried or hidden structural element, such as a wing spar that is not directly accessible for visual or other checks when searching for cracks or cracks.

If the details are a crack or a similar defect, repeated cyclic or fatigue loading can lead to the growth of this crack. This is called fatigue crack propagation. The propagation of cracks during fatigue can lead to the formation of a sufficiently large cracks, which can drastically be distributed when the combination of the size of the crack and the load will be sufficient, so that they will exceed the material resistance to cracking. Thus, performance in resistance of a material to the propagation of cracks during fatigue are significant advantages to ensure the long service life of the structure of the aircraft. The smaller the spread of trash the us, so much the better. The rapid spread of cracks in the structural element of the aircraft can lead to catastrophic failure without adequate time for discovery, while the slow propagation of cracks leaves time to detect and take measures to remedy or repair. Therefore, a low rate of growth of fatigue cracks is the preferred property.

The velocity of propagation of cracks in the material during the load cycle is affected by the length of the crack. Another important factor is the difference between the maximum and minimum loads between cycles. One dimension assesses the influence of the crack length and the difference between the maximum and minimum loads, which is called the band index of the intensity of cyclic load or ΔK, which is expressed in thousands (pounds/inch2)inch1/2similarly, the index of the intensity of the workload used to measure the resistance to cracking. Range (ΔK) index of intensity of load is the difference between the indices of the intensity of the load at maximum and minimum loads. Another feature influencing the fatigue crack propagation, is the ratio between the minimum and maximum loads during the cycle, and this number is R is called the ratio of the load and is denoted as R, when this value of the ratio of 0.1 means that the maximum load is 10 times higher than the minimum load. Load or the ratio of the load can be positive or equal to zero. Test the rate of growth of fatigue cracks usually performed in accordance with ASTM E647-88 (and in accordance with other techniques, well known in the art). Used here the notation Kt denotes theoretical concentration ratio of the load in accordance with the description ASTM E1823.

The speed of propagation of fatigue cracks of the material can be measured using a test sample containing a crack. One such test sample has a length of approximately 12 inches and a width of 4 inches (30,5×10.2 cm) incision in the center, passing in the transverse direction (width perpendicular to the length). The incision is about to 0.032 inches in width and about 0.2 inches (0.08 to 0.51 cm), includes 60° bevel on each side of the slot. The test sample is subjected to cyclic loading, the crack grows on the end (ends) cut. Once the crack reaches a predetermined length, the length of the cracks periodically measured. The crack growth rate can be calculated for a given crack growth by dividing the change of crack length (denoted Δa) by the number C is clov load (ΔN), in the result of which was obtained the value of the crack growth. The speed of propagation of cracks is presented in the form Δa/ΔN or 'da/dN' and is expressed in units of inches/cycle. The velocity of propagation of fatigue cracks of the material can be determined on the basis of the stretchable panel with a Central crack. For comparison, when using values of R=0,1, which was obtained during tests with relative humidity greater than 90%, when ΔK in the range of from about 4 to 20 or 30 material, in accordance with the present invention, showed relatively good resistance to growth of fatigue cracks. However, the exceptional performance of fatigue S-N make the material, in accordance with the present invention, much more suitable for use as a buried or hidden item, such as a wing spar.

Products, in accordance with the present invention show a very good resistance to corrosion in addition to very good strength and durability, as well as the properties of resistance to damage. Corrosion resistance detachment for the products in accordance with the present invention, may be the level of S or better (meaning "EA" is only pitting) (pitting) when testing EXCB intended for test images, the wire is held either by the average thickness (T/2) profile or one tenth of the thickness from the surface (T/10) (T" represents the thickness profile), or for both values. Tests EXCB known in the art and described in the well-known ASTM No. G34. Figure EXCO at the mark "S" is seen as a good resistance to corrosion, because it is believed that it is acceptable for some commercial aircraft; the indicator "EA" is even better.

Resistance to corrosion cracking when under load in the direction of the short section is often seen as an important property, in particular, for relatively thick elements. Resistance to cracking under conditions of corrosion under load products, in accordance with the present invention, in the direction of the shorter cross-section may be the equivalent having to go for round bar with a diameter of 1/8 inch (0,32 cm) tests with alternating immersion for 20 or, alternatively, 30 days at a load of 25 or 30 thousand pound/inch1/2(173 or 207 MPa) or higher, when using the test procedures in accordance with ASTM G47 (including ASTM G44 and G38 for samples in the form of a C-ring and G49 for bars with a diameter of 1/8 inch (0,32 cm), these standards ASTM G47, G44, G49 and G38 are well known in the art.

As a General indicator of corrosion detachment the corrosion resistance under load plate can typically have conductivity, at least about 36 or preferably 38-40%, or more in accordance with the International standard annealed copper (% of IACS). Thus, good value corrosion resistance detachment, in accordance with the present invention, had the indicator EXCB at the level of "EB" or better, but in some cases can be defined by other criteria corrosion resistance or may require developers aircraft design, such as resistance to corrosion cracking when under load or the electrical conductivity. To meet the requirements of one or more of these specifications is considered as a good corrosion resistance.

The present invention has been described with a certain emphasis on wrought plate, which is preferred, but this assumes that other forms of the product can also be used advantages of the present invention, including components obtained by extrusion and forging. This goal was highlighted in the description of the stringers plating wing or fuselage of the type of ribs, which may be J-shaped, Z - or C-shaped, or can even be made in the form of spooring channel. The purpose of these ribs is to strengthen the lining of the wing or fuselage of an aircraft or any other form, which can be fixed on n the th without substantial increase in weight. Although in some cases it is preferable to save production to use separately attached to the stringers, which can be manufactured by machining from a much thicker plate by removing the metal between the geometric elements of the ribs, leaving only the shape of the ribs, which is integral with the main profile of the sail wing thickness, eliminating thus the need to use all of the rivets. In addition, the present invention has been described in relation to a thick plate intended for making it through mechanical processing element of the wing spar, as described above, the element of the spar, in General, corresponds to the length of the cladding material of the wing. In addition, significant improvement in the properties of materials, in accordance with the present invention provide highly practical their use as a thick plate moulds.

By reducing the sensitivity of hardening is considered that, when carried out welding product of the alloy, in accordance with the present invention, with the second product in the weld zone, exposed to heat, will provide increased preservation of strength, resistance to fatigue, resistance to cracking and/or corrosion resistance splaat true regardless do welding products of this alloy with the use of welding in the solid state, including welding by friction during the rotation, or known technologies, or technologies that will be developed in the future, including, but not limited to, electron beam welding and laser welding. In the application in practice of the present invention, both the welded parts can be made of an alloy of the same composition.

For some parts/products, manufactured in accordance with the present invention, it is possible to form such parts/products by forging. Forming forging reduces the cost of production, enabling them to produce more complex shapes of parts of the wing, usually with elements of thinner cross-section. In the course of morphogenesis by aging these details mechanical limit in the matrix at an elevated temperature, usually about 250°F or higher during the period up to tens of hours, and required paths lead after stress relaxation. In particular, during processing of high-temperature artificial aging, such as processing at temperatures above approximately 310°F (160°C), may be forming or deformation of the metal to obtain the desired shape. Usually this kind of deformation is relatively p is ostim and includes giving a very slight curvature to the width of the element in the form of a plate, while giving a slight curvature along the length of the specified element in the form of a plate. You may need to provide such conditions of negligible curvature during artificial aging, in particular, during high-temperature second stage of artificial aging. Typically, the plate material is heated to a temperature above approximately 300°F (149°C), for example, to temperatures in the range from approximately 320 or 330°F (160 or 166°C), and the plate can usually be installed on a convex shape and loaded by clamping or application of the load to the opposite edges of the plate. Stove, to a greater or lesser extent, takes the path of this form within a relatively short period of time, but after cooling deleting a force or load is some elastic straightening. The expected degree of elastic straightening compensate by calculating the curvature or contour shape, somewhat enlarged in relation to the desired shape of the plate to compensate for the elastic straightening. Most preferably, after forming by aging followed by a third step artificial aging at low temperature, such as approximately 250°F (121°C). Before or after this treatment, shaping aging element in the form of a plate can be machined, for example, to make the plate of the clinoid shape in which the part is intended for installation closer to the fuselage, the imp is both thicker, and the part located closer to the end of the wing, perform more subtle. Additional machining or other operations of forming, if necessary, can also be performed both before and after shaping by aging. For airplanes with a high capacity can require relatively thicker plates and a higher level of formation than previously used, mainly, for plates with thinner cross-section.

Different product forms of the alloy, in accordance with the present invention, that is, as a thick plate (Fig), and forgings (Fig), were manufactured, processed by aging, and samples of the appropriate size were fabricated for testing time to fatigue fracture (S/N), which coincided with the known stages of testing time to fatigue fracture with through holes. We used the following exact formulations for such forms of products:

For these estimates, time to fatigue fracture with a through hole orientation L-T specific test parameters for plates and wrought forms of products included: the value of Ktequal to 2.3, frequency 30 Hz, the value of R=0.1 and the value of the relative humidity (RH) greater than 90%. The test results for plates represent the go in the form of schedule annexed Fig, and the test results for forgings presented on Fig. As the stove, and wrought the form was tested for several values of the thickness of the product (4, 6 and 8 inches (10,2, of 15.2 and 20.3 cm)).

As shown in Fig line (solid) average S/N conducted through both sets of data for plate thickness of 6 inches (15.2 cm) (alloys D and E, described above). Then was drawn strip 95% confidence level (corresponding to the upper and lower dotted lines) around the above line average value for the thickness of 6 inches (15.2 cm). These data was obtained set of points representing the extrapolated currently, the minimum value of operating time (S/N) to fatigue fracture with a through hole. The exact values for these presents points were:

Then pig was plotted solid line And connecting the above values are extrapolated to the present time, the minimum S/N in table 12. These preferred values for the minimum S/N were imposed specified one of the manufacturers of jet aircraft line values of S/N for plates of alloys 7040/7050-T (thickness from 3 to 8.7 inches (7,6-22,1 cm)) and a plate of alloy 7010/7050-T (thickness of 2-8 inches (5.1 to 20.3 cm)). Line a-a represents the relative improvement of the alloy, in accordance with the present invention, the characteristic of NARBO the key S/N to fatigue fracture in comparison with the known commercial alloys 7XXX, used in the aerospace industry, even when the comparative data for the last known alloys have been specified for different orientation (T-L).

According to the operating time (S/N) to fatigue fracture with a through hole for forgings with different sizes (i.e. 4 inch, 6 inch and 8 inch (to 10.2, and 15.2 and 20.3 cm)) was plotted as a dotted line, intended for the mathematical representation of the average values for forging thickness of 6 inches for the comparative sample e, and a thickness of 8 inches (20.3 cm) for the comparative sample D. it Should be noted that several of the samples subjected to the tests were not destroyed during these trials, they were grouped together in the circle on the right side on Fig. After that display multiple points representing values extrapolated at present, the minimum operating time (S/N) to fatigue fracture with a through hole. The exact values of these mapped points were:

Then pig was plotted solid line (-) connecting the above values extrapolated at present, the minimum S/N for forgings in accordance with the above table 13.

On Fig shows curves of the rate of fatigue crack growth (FCG) to plate (thickness in 4 and 6 inch (10.2 and 15.2 cm)) in the direction L-T, that is in direction T-L) and the forged product (only in the direction L-T, 6 inches (15.2 cm), which were manufactured in accordance with the present invention. In fact, the tested compounds are shown in the table above 11. During these tests, conducted using FCG procedures described above are used, in particular: frequency = 25 Hz, a value of R=0.1 and relative humidity (RH) above 95%. On these curves for different shapes and thickness of the product was obtained by a set of data points representing the extrapolated at the present time, the maximum value of FCG in accordance with the present invention. The exact values of these points were:

For extrapolated currently, the maximum values FCG was drawn solid curve line (C-C) for thick plates and forgings in accordance with the present invention, which were imposed FCG values specified one of the manufacturers of jet plate 7040/7050-T (thickness from 3 to 8.7 inches (7,6-22,1 cm)), and these values were taken as the direction of orientation of the L-T and in the direction of orientation of T-L.

The shape of the product in the form of a plate, in accordance with the present invention were also tested for cracking with a hole, including the specified drilling holes (diameter less than 1 inch (2.5 cm)in the sample to be tested can be installed in drilled is twistie split sleeve with subsequent pulling of the mandrel with a variable size is exceeded through the said split sleeve and a pre-drilled hole. In these tests the product in the form of a plate thickness of 6 and 8 inches (of 15.2-20.3 cm), in accordance with the present invention, did not show trends in the formation of any cracks from the drilled holes, showing, therefore, a very good performance.

After the description of the preferred in the present embodiments should be understood that the present invention may have other ways to perform within the scope of the attached claims.

1. Plate of aluminum alloy containing components in the following ratio, wt.%:

Znfrom 6.4 to 8.5
Mgfrom 1.4 to 1.9
Cfrom 1.4 to 1.85
Zrfrom 0.05 to 0.15
Tifrom 0.01 to 0.06
Feto 0.15
Sito 0.12
aluminium co
elements and impuritiesthe rest,

the plate has a thickness of at measures is 4 inches, and aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+92,5, where FT is a K1c(L-T) resistance to cracking in the direction L-T, is equal to at least 29 thousand (pounds/inch2)inch1/2and TYS is the yield strength tensile equal to at least 66 thousand (pounds/inch2).

2. Stove according to claim 1, characterized in that the aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+93.

3. Stove according to claim 1, characterized in that the aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+93,5.

4. Stove according to claim 1, characterized in that the aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+94.

5. Stove according to claim 1, characterized in that the aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+94,5.

6. Stove according to claim 1, characterized in that the aluminum alloy has a yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+95.

7. Stove according to claim 1, characterized in that the aluminum alloy is within the yield strength in tension and resistance to cracking, satisfies the relation FT≥-0,8333∙TYS+95,5.

8. Stove according to any one of claims 1 to 7, characterized in that the aluminum alloy has a resistance to cracking under corrosion load equal to at least 25 thousand (pounds/inch2), and the value of conductivity of at least 39.5 percent from the values of the International standard annealed copper (IACS).

9. Stove according to any one of claims 1 to 7, characterized in that the aluminum alloy has a resistance to cracking under corrosion load equal to at least 35 thousand (pounds/inch2), and the value of conductivity equal to at least 40.5 percent from IACS values.

10. Stove according to any one of claims 1 to 7, characterized in that the aluminum alloy has a resistance to cracking under corrosion load equal to at least 45 thousand (pounds/inch2), and the value of conductivity of at least 41.5 percent from the values of IACS.



 

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FIELD: metallurgy.

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4 cl, 2 tbl, 10 dwg, 2 ex

FIELD: metallurgy.

SUBSTANCE: 50-305 mm-thick article is made from alloy of the following chemical composition, in wt %: Zn - 3-11, Mg - 1-3, Cu - 0.9-3, Ge - 0.03-0.4, Si - not over 0.5, Fe - not over 0.5, Ti - not over 0.3, aluminium and common or unavoidable elements and impurities making the rest. Proposed method comprises slab casting, heating and/or annealing treatment of cast slab, slab hot treatment, optional cold treatment, heat processing to solid solution of slab (PSS), cooling of PSS slab, optional expansion or compression of cooled PSS slab or any other cold treatment of PSS slab to remove strain, ageing of cooled PSS slab to reach required state.

EFFECT: high hardness, ductility and lower sensitive to quenching.

20 cl, 2 tbl, 1 ex

FIELD: metallurgy.

SUBSTANCE: aluminium-based alloy is used for production of strained semis as forgings and tubes for gas centrifuges, low-pressure compressors, vacuum molecular pumps and heavy-duty units operated at mildly increased temperatures. Proposed composition contains the following substances, in 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, beryllium - 0.0001-0.005, cobalt - 0.05-0.15, nickel - 0.35-0.65, iron - 0.25-0.65, aluminium making - the rest.

EFFECT: higher strength at sufficient ductility and reduce density of alloy.

1 ex, 3 tbl

FIELD: metallurgy.

SUBSTANCE: proposed method comprises casting the ingots of alloy containing the following components in wt %: zinc - 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, aluminium making the rest, at the temperature of 690-710°C and casting rate of 25-50 mm/min. Ingots are homogenised at 450-470°C for 8-12 hours and subjected to hot forming at 10-530°C and outflow rate of 0.1-4.0 m/min. Besides it includes air or air-water mix quenching and two-step ageing: at 90-110°C for 6-12 hours and at 160-190°C for 4-10 hours.

EFFECT: production of log parts with high operating properties.

4 cl, 1 ex, 6 tbl, 4 dwg

FIELD: metallurgy.

SUBSTANCE: alloy contains, wt %: 3.5-4.5 zinc, 3.5-4.5 magnesium, 0.6-1.0 copper, 2.0-3.0 nickel, 0.25-0.3 zirconium, aluminium - balance, at the same time after strengthening thermal treatment the alloy has yield point of 570 MPa, strength limit of 600 MPa, hardness of 160 HY, and after deformation at 440-480°C with speed of 0.001-0.01 1/s the alloy has elongation of more than 500%.

EFFECT: production of alloy with equiaxial homogeneous fine-grain structure.

4 ex

FIELD: metallurgy.

SUBSTANCE: alloy contains the following, wt %: silicon 6.6-7.4, magnesium 0.31-0.45, copper 0.18-0.32, manganese 0.15-0.45, iron 0.15-0.4, aluminium is the rest, at that, the alloy has liquidus temperature within 608 to 620°C; temperature of balanced solidus of not less than 552°C and structure after heat treatment as per mode T66, which contains the amount of inclusions of silicon phase within 6.4 to 7.5 vol. %; iron in the alloy structure is completely bound to skeletal inclusions of phase Al15(Fe,Mn)3Si2, and magnesium is completely bound to secondary extractions of phase Al15Cu2Mg8Si6.

EFFECT: creation of alloy for obtaining high-duty shaped castings and having high technological and operating characteristics.

2 cl, 2 tbl, 2 ex, 2 dwg

FIELD: metallurgy.

SUBSTANCE: proposed composition contains the following substances, in wt %: zinc - 5.0-5.8, magnesium - 1.1-1.2, chromium - 0.2-0.3, copper - 0.1-0.4, titanium - 0.05-0.15, cerium - 0.005-0.05, samarium - 0.005-0.05, silicon - not over 0.3, iron - not over 0.3, zirconium - not over 0.005, aluminium making the rest. Method of making semis from said aluminium alloy comprises first thermal treatment at up to 480°C, cooling to room temperature, and second thermal treatment at up to 200°C.

EFFECT: higher strength, lower residual strain.

11 cl, 4 ex

FIELD: metallurgy.

SUBSTANCE: invention may be used for making critical parts operated at high loads at 150°C, e.g, those of aircraft, cars and trucks, etc. Proposed composition contains the following substances, in wt %: 5.5-6.5 Zn, 1.7-2.3 Mg, 0.4-0.7 Ni, 0.3-0.7 Fe, 0.02-0.25 Zr, 0.05-0.3 Cu. Its solvus temperature does not exceed 410°C while hardness makes, at least 150 HV.

EFFECT: higher strength and machinability.

3 cl, 3 tbl, 3 ex, 2 dwg

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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