Classifications

• • C—CHEMISTRY; METALLURGY
  • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
  • C22C—ALLOYS
  • C22C21/00—Alloys based on aluminium
  • C22C21/10—Alloys based on aluminium with zinc as the next major constituent
• • C—CHEMISTRY; METALLURGY
  • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
  • C22F—CHANGING THE PHYSICAL STRUCTURE OF NON-FERROUS METALS AND NON-FERROUS ALLOYS
  • C22F1/00—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working
  • C22F1/04—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of aluminium or alloys based thereon
  • C22F1/053—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of aluminium or alloys based thereon of alloys with zinc as the next major constituent

Definitions

• the present invention relates generally to rolled, extruded and/or forged products made of Al—Zn—Cu—Mg alloy treated by solution heat treatment, quenching, cold working and artificial aging, and particularly structural members made from such products and designed for use such as in aircraft construction.
• 7xxx type alloys are used for wing structural members (except for undersurface wing members).
• Patent application WO 02/052053 describes three Al—Zn—Cu—Mg type alloys with composition (a) Zn 7.3+Cu 1.6; (b) Zn 6.7+Cu 1.9; (c) Zn 7.4 Cu 1.9; each of these three alloys also containing Mg 1.5+Zr 0.11.
• This WO publication also describes appropriate thermomechanical treatments for making structural members for aircraft.
• a 7040 alloy with the following standardized chemical composition is known: Zn 5.7-6.7 Mg 1.7-2.4 Cu 1.5-2.3 Zr 0.05-0.12 Si ⁇ 0.10 Fe ⁇ 0.13 Ti ⁇ 0.06 Mn ⁇ 0.04 other elements ⁇ 0.05 each and ⁇ 0.15 total.
• a 7475 alloy with the following standardized chemical composition is also known: Zn 5.2-6.2 Mg 1.9-2.6 Cu 1.2-2.9 Cr 0.18-0.25 Si ⁇ 0.10 Fe ⁇ 0.12 Ti ⁇ 0.06 Mn ⁇ 0.06 other elements ⁇ 0.05 each and ⁇ 0.15 total.
• Alloys in the 2xxx series are routinely used, for example the 2324 alloy, for some structural members of civil aircraft wings such as under wing members.
• Alloys conventionally used for fuselage structural members typically belong to the 2xxx series, for example the 2024 alloy.
• a purpose of the present invention was to obtain aircraft structural members, and particularly fuselage members made of Al—Zn—Cu—Mg alloy, with a higher mechanical strength than is possible in prior alloys, with comparable damage tolerance and sufficient formability.
• Another purpose was to obtain aircraft structural members, and particularly members for the under surface wings of aircraft, or for machining integral structures made of Al—Zn—Cu—Mg alloy, with a better compromise between mechanical strength, toughness and fatigue strength properties, than is possible to achieve with prior materials.
• a work-hardened product (preferably a rolled, extruded and/or forged product) of an alloy comprising (% by weight):
• a structural member suitable for aeronautical construction and particularly for an aircraft fuselage, or for members of the under surface of an aircraft wing, or an integral structural member for an aircraft, made from such a work hardened product, and particularly from such a rolled or extruded product.
• static mechanical characteristics in other words the ultimate tensile strength Rm, the tensile yield strength R p02 and the elongation at fracture A, are determined by a tensile test according to standard EN 10002-1, the location at which the pieces are taken and their direction being defined in standard EN 485-1.
• the fatigue strength is determined by a test according to ASTM E 466, and the fatigue crack propagation rate (using the da/dn test) according to ASTM E 647.
• the R curve is determined according to ASTM standard 561.
• the critical strength intensity factor K C in other words the intensity factor that makes the crack unstable, is calculated starting from the R curve.
• the stress intensity factor K CO is also calculated by assigning the initial crack length to the critical load, at the beginning of the monotonous load. These two values are calculated for a test piece of the required shape. K app denotes the K CO factor corresponding to the test piece that was used to make the R curve test. Unless otherwise mentioned, the crack size at the end of the fatigue precracking stage is W/4 for test pieces of the M(T) type, and W/2 for test pieces of the CT type, wherein W is the width of the test piece as defined in standard ASTM E561.
• extruded product includes so-called “drawn” products, in other words, products produced by extrusion followed by drawing.
• structural member in this specification refers to a mechanical part used in mechanical construction, for which failure could endanger the safety of the said construction and/or its users, or others.
• these structure members include particularly members making up the fuselage (such as the fuselage skin, fuselage stiffeners or stringers, bulkheads, fuselage circumferential frames, wings (such as wing skin), stringers or stiffeners, ribs and spars and the tail fin composed particularly of horizontal and vertical stabilisers, and floor beams, seat tracks and doors.
• integral structure means the structure of part of an aircraft that was designed to achieve material continuity over the largest possible size in order to reduce the number of mechanical assembly points.
• An integral structure may be made either by in-depth machining, or by the use of shaped parts for example obtained by extrusion, forging or casting, or by welding of structural members made of weldable alloys.
• the result is large single-piece structure members, with no mechanical assembly or with a small number of mechanical assembly points compared with an assembled structure in which the thin or thick plates depending on the destination of the structure member (for example fuselage member or wing member) are fixed, usually by riveting, onto stiffeners and/or frames (that can be made by machining from extruded or rolled products).
• the present invention can advantageously be applied to an aluminium alloy containing from about 6.7% to about 7.3% of zinc.
• the zinc content should preferably be high enough to achieve good mechanical properties, but if it too high, the sensitivity of the alloy to quenching may increase, which in particular introduces a risk for thick products of degrading the compromise between target properties.
• the product is a plate thinner than about 20 mm. In another advantageous embodiment of the invention, the product is a thick plate, thicker than about 20 mm.
• the chemical composition of the Al—Zn—Cu—Mg alloy was chosen such that the Mg/Cu ratio of the alloy according to the invention is preferably below about 1. Preferably, this ratio is kept at a value less than 0.9. A value less than 0.85, or even about 0.8 is preferred.
• zirconium content from about 0.07 to about 0.13% made it possible to achieve a better compromise between R P0.2 , toughness (at ambient temperature or when cold) and fatigue strength (particularly the propagation rate of fatigue cracks), for this composition of major elements Al—Zn—Cu—Mg. If the content of Zr exceeds about 0.12%, there may be a significant risk of primary Al 3 Zr type phases being formed, unless cooling is fast enough; in the case of semi-continuous casting, such a sufficient rate can be achieved particularly when billets are being cast.
• the Zr content for rolled products is preferably less than about 0.12%, and advantageous results have been obtained with a content of from about 0.07 to about 0.09%.
• a zirconium content of up to about 0.13% can be suitable for billets in some embodiments.
• silicon and iron contents should preferably each be kept below about 0.15% and particularly preferably below about 0.10% to have good toughness.
• the iron content preferably does not exceed about 0.07%, and the silicon content preferably does not exceed about 0.06%.
• An alloy according to the instant invention can be cast according to one of the techniques known to those skilled in the art to obtain unwrought products such as an extrusion billet, or a rolling ingot.
• This unwrought product possibly after scalping, is then homogenized, typically for a duration of 15 to 16 hours at a temperature of preferably from about 470 to about 485° C.
• the unwrought product is then transformed hot into extruded products (particularly bars, tubes or sections), hot rolled plates or forged parts.
• extruded products particularly bars, tubes or sections
• hot rolled plates or forged parts the inventors have observed that surprisingly, thick products according to the invention could be hot rolled at a temperature of about 350° C., which is much lower than the temperature usually used for this type of product (which is about 415 to 440° C.) without affecting the required compromise between properties for thick products used in aircraft structures.
• Products obtained are then preferably solution heat treated.
• This solution heat treatment may be made in any appropriate furnace such as an air furnace (horizontal or vertical) or a salt bath furnace.
• this solution heat treatment is carried out at a temperature from about 470 to 480° C. and preferably from about 475 to about 480° C. preferably for at least 4 hours.
• leading to thin products ( ⁇ 10 mm) the solution heat treatment is carried out at preferably from about 470° C. to about 475° C., and the duration of the solution heat treatment, for which the optimum value depends on the product thickness, is typically at least about one hour.
• the products are then quenched, preferably in a liquid medium such as water, the liquid preferably being at a temperature of not more than about 40° C.
• the products are then usually subjected to controlled stretching with a permanent set of the order of preferably from about 1 to 5%, and particularly preferably from about 1.5 to 3%.
• R p0.2 (L) of at least about 520 MPa
• products according to the present invention advantageously replace structural members made of alloys known as 2 ⁇ 24 alloys, for example a 2024 or 2324 alloy.
• rolled products according to the present invention may be thinner than about 10 mm and thus be used, for example, as a fuselage skin. They may also be thicker than about 10 mm and thus be used as structural members such as for lower wing structures.
• Rolled products more than about 40 mm thick may be used, for example, for the manufacture of structural members by integral machining as described below.
• Rolled products with a thickness of more than about 60 mm can be used, for example, for manufacturing stiffeners or frames, particularly for large capacity aircraft.
• Products according to the present invention may be clad on at least one face thereof if desired for any reason using methods and with alloys conventionally used to clad products made of Al—Zn—Cu—Mg type alloys. This is particularly attractive for plates used for manufacturing aircraft fuselage members that have to resist corrosion.
• One exemplary cladding alloy that can be used is 7072.
• One particularly advantageous use of products according to the instant invention is related to the concept of the integral structure in aeronautical construction.
• a large proportion of aircraft structures are sized as a function of a compromise between damage tolerance and resistance to static loads.
• Requirements for damage tolerance are, for example specified in the article “Damage Tolerance Certification of Commercial Aircraft” by T. Swift, ASM Handbook vol. 19 (1996), pp 566-576.
• Sizing under static loads is explained for example in the book “Airframe Stress Analysis and Sizing” by M. Niu, Hong Kong Commilit Press Ltd, 1999, particularly pages 607 to 654. From the material point of view, it is known that the damage tolerance of alloys in the 7xxx series, and particularly their toughness, generally decreases when their yield strength increases.
• x is typically preferably from about 15 to about 30%.
• a weight saving of the same order of magnitude as the improvement in toughness namely about 10%
• a product according to the invention with a yield strength R P0.2(L) at mid-thickness equal to at least about 540 MPa and a toughness K app(LT) measured on an M(T) type specimen with a width W of 16 inches (about 406 mm) equal to at least about 140 Mpa ⁇ square root ⁇ m, can be used to make structural members for aircraft such as a wing skin member with a weight saving equal to at least 10% compared with the same part with the same shape and size made from a 7475 alloy according to the state of the art and typically having an R P0.2(L) at mid-thickness equal to 475 MPa, and an K app(LT) measured on an M(T) type specimen with a width W of 16 inches (about 406 mm) equal to
• the inventors have observed that refining the grain to a lower level than is accepted in normal practice during casting can give a particularly attractive compromise between properties, particularly for toughness.
• a refining agent made of TiC for example addition of an A13% Ti0.15% C wire
• the solidification germ obtained with this approach having a different compromise between germination and growth than is possible with germs obtained for example by refining with A15% Ti1% B (in other words a TiB 2 type germ).
• the level of this refining may be quantified by the quantity of C added, since it indirectly corresponds to the quantity of added solidification germs and the quantity of free Ti (not combined with C) into the alloy.
• the stoechiometry of the germ is not definitively quantified, it can be considered that the germ comprises TiC, each C atom combining with a Ti atom to form the said germs.
• the quantity of added germs is preferably proportional to the quantity of refining agent (in kilograms) added per ton of liquid metal multiplied by y/o, in other words proportional to A (number of kilograms of refining agent added per ton of metal) ⁇ y %.
• the addition of germs can be quantified by specifying 3 g/t of added C (2 ⁇ 0.0015 kg/t).
• a refining agent comprising titanium and carbon is also added such that the added carbon quantity is preferably between 0.4 and 3 g/t of carbon, more preferably between 0.6 and 2 g/t and such that the total content of Ti in the final product is between 50 and 500 ppm (by weight) and preferably between 150 and 300 ppm.
• An alloy N was made for which the chemical composition complies with the invention.
• the liquid metal was treated firstly in a holding furnace by injecting gas using an IRMA® type of rotor, and then in an Alpur® type of ladle, these two trademarks belonging to the inventors. Refining was done in line, in other words in the channel between the holding furnace and the Alpur® ladle, with 1.1 kg/tonne of Al-3% Ti-0.15% C wire (9.5 mm diameter).
• An industrial sized rolling ingot was cast. It was relaxed for 10 h at 350° C.
• the product thus cast was homogenised after scalping for 15 hours at a temperature between 471° C. and 482° C. (between 880° F. and 900° F.) and then hot rolled to a thickness of 5 mm (0.2 inches).
• the rolling start temperature was 450° C. (840° F.) and the rolling end temperature was 349° C. (660° F.). Plates with width 178 mm (7 inches) and length 508 mm (20 inches) were sampled.
• These coupons were solution heat treated in a salt bath furnace for 1 hour at 472° C. and then quenched in water and tensioned to obtain a permanent deformation of 2%.
• the coupons thus obtained were then subjected to a two-step artificial aging treatment, the first step being 6 hours at 105° C., the second step being 18 hours at 155° C., in order to reach the peak of mechanical properties.
• the alloy was cast firstly by treating the liquid metal in a holding furnace by injecting gas using an IRMA® type of rotor, and then in an Alpur® type of ladle. Refining was done in line, in other words in the channel between the holding furnace and the Alpur® ladle, with 0.7 kg/tonne of AT5B wire (9.5 mm diameter).
• the cast rolling ingots were stress relieved for 10 hours at 350° C. These rolling ingots were then homogenised for 12 hours at 500° C., then hot rolled (end of rolling temperature between 230 and 255° C.) to a thickness of 6 mm.
• a solution heat treatment was then carried out in a salt bath furnace for 1 hour at 500° C. on the 600 mm by 200 mm coupons. This operation was followed by quenching in cold water at about 20° C. and stretching with a permanent set of 2% (temper T351).
• Rolling ingots made of a 7xxx alloy according to prior art were also cast (reference G), in the same founding device as plates made with 2xxx alloy described above.
• the resulting rolling ingot was homogenised for 24 hours at 470° C. and then 24 hours at 495° C., then hot rolled (end of rolling temperature between 230 and 255° C.) to a thickness of 6 mm.
• a solution heat treatment of 1 hour was then carried out at 450° C. in a salt bath furnace on a 600 mm by 200 mm coupon. This operation was followed by quenching in water and stretched with a permanent set of 2%.
• the coupon was then subjected to artificial aging treatment for 5 hours at 100° C. then 6 hours at 155° C., in order to achieve the peak mechanical properties (temper T6).
• a rolling ingot made of an AA7475 alloy was also cast (reference H) according to conventional processes according to prior art.
• the rolling ingot thus obtained was homogenised for 9 hours at 480° C., and then hot co-rolled at a temperature of about 270° C. with a 7072 cladding plate, until a sheet with a thickness of 4.5 mm was obtained.
• the 7072 cladding accounts for about 2% of the final thickness.
• the product thus obtained was solution heat treated in a salt bath furnace for 45 minutes at 478° C., then quenched in water at a temperature of about 20° C., and then stretched with a permanent set of 2%. It was then subjected to a two-step artificial aging operation for 4 hours at 120° C., then 24 hours at 162° C. (temper T76).
• the ultimate tensile strength R m (in MPa), the tensile yield strength at 0.2% elongation R P0.2 (in MPa) and the elongation at failure A (in %) were measured using a tensile test according to EN 10002-1.
• the ultimate tensile strength and tensile yield strength of the plate according to the invention in both measured directions is very much higher than the corresponding values for plates made of a 2xxx alloy.
• the elongation of the plate according to the invention is lower than that of plate E, but is sufficient for the target applications.
• the alloy according to the invention has a significantly improved ultimate strength and yield strength for a comparable elongation.
• K app for the plate according to the invention is much greater than the value for plates made of 7xxx alloy according to prior art, and is of the same order of magnitude as for plates made of 2xxx alloy.
• the fatigue behaviour was also tested according to ASTM standard E 647, measuring the crack propagation rate in plate N in comparison to plates E, F and G.
• the test pieces used were of the C(T) type, where W is 76.2 mm (3 inches).
• the plate according to the invention behaves just is well in fatigue as plates according to prior art.
• the plates thus obtained were then subjected to a solution heat treatment at 479° C. for 4 hours (total time, about 1 ⁇ 3 of which is spent in the temperature increase), and were then quenched and tensioned such that the resulting permanent deformation is 2%.
• the plates were then subjected to an artificial aging treatment for 8 hours at 160° C.
• Alloy I (AA2324) was subjected to a conventional procedure to obtain a plate made of AA 2324 alloy, 25.4 mm thick in the T39 temper, in other words a homogenisation step followed by a hot rolling step, then solution heat treatment and quenching, followed by cold working of about 9%, and controlled stretch with a permanent set of between 1.5 and 3%.
• TABLE 6 Static mechanical properties L direction Thickness Rm R P0.2(L) Plate [mm] [MPa] [MPa] A [%] M 25.4 570 540 12.3 I 25.4 490 470 14
• the alloy according to the invention had better toughness than the conventional alloy I under all conditions. And also surprisingly, the alloy according to the invention had a value of K app(LT) that was of the same order at ⁇ 54° C. as it is at ambient temperature.
• the fatigue behaviour was also tested according to ASTM standard E 647, by measuring the crack propagation rate in plate M in comparison with plate I.
• the test pieces used were of the C(T) type where B is equal to 9.52 mm (0.375 inches) and W is equal to 101.6 mm (4 inches).
• exfoliation corrosion behaviour of plates in this test was evaluated according to ASTM standard G34; this test was done on the surface and at mid-thickness under conditions adapted to 7xxx alloys for plate M according to the invention, and under conditions adapted to 2xxx alloys for plate I.
• Sample M according to the invention was classified EA, both at the surface and at mid-thickness, while sample I according to prior art was classified EA at the surface and EB at mid-thickness. Therefore, the performance of the plate according to the invention in terms of exfoliation corrosion is at least as good, if not better, than the plate according to prior art.
• plate M is better for static mechanical characteristics, K app , fatigue resistance and for the crack propagation rates.
• An alloy P similar to alloy M in example 2 was produced.
• a manufacturing procedure similar to that for example 2 was used to make thick integrally hot rolled plates from this alloy (input temperature 420-440° C.), with a thickness of 75 mm.

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