ALUMINUM-LITHIUM ALLOY

Abstract: An aluminium-lithium alloy exhibiting good fracture toughness and relatively high strength, having a composition of 2.0 to 2.4% Li, 0.3 to 0.9% Mg. 2.1 to 2.9% Cu, 0.08 to 0.15% Zr, up to 0.15% Fe, up to 0.12% Si, up to 0.25% Zn, up to 0.15% Ti, up to 0.1% Cr, balance Al. A preferred composition has 2.2% Li, 0.6% Mg, 2.5% Cu, 0.12% Zr, balance Al.

Description: The present invention relates to aluminum-lithium alloys and more particularly to an aluminum-lithium alloy composition with good fracture toughness and high strength. It has been estimated that current large commercial transport aircraft may be able to save from 15 to 20 gallons of fuel per year for every pound of weight that can be saved when building the aircraft. Over the projected 20 year life of an airplane, this savings amounts to 300 to 400 gallons of fuel. At current fuel costs, a significant investment to reduce the structure weight of the aircraft can be made to improve overall economic efficiency of the aircraft. The need for improved performance in aircraft of various types can be satisfied by the use of improved engines, improved airframe design, and improved or new structural materials in the aircraft. The development of new and improved structural materials has recently received increased attention and is expected to yield significant gains in performance. Materials have always played an important role in dictating aircraft structural concepts. In the early part of this century, aircraft structure was composed of wood, primarily spruce, and fabric. Because shortages of spruce developed in the early part of the century, lightweight metal alloys began to be used as aircraft structural materials. At about the same time, improvements in design brought about the development of the all metal cantilevered wing. It was not until the 1930's, however, that the metal skin wing design became standard, and firmly established metals, primarily aluminum alloys, as the major airframe structural material. Since that time, aircraft structural materials have remained remarkably consistent with aluminum structural materials being used primarily in the wind, body and empennage, and with steel comprising the material for the landing gear and certain other speciality applications requiring very high strength materials. Several new materials are currently being developed for incorporation into aircraft structure. These include new metallic materials, metal matrix composites and resin matrix composites. It is believed that improved aluminum alloys and carbon fiber composites will dominate aircraft structural materials in the coming decades. While composites will be used in increased percentages as aircraft structural materials, new low-density aluminum alloys, and especially aluminum-lithium alloys show great promise for extending the use of aluminum alloys in aerospace structures. Heretofore, aluminum-lithium alloys have been used only sparsely in aircraft structure. The relatively low use has been caused by casting difficulties associated with aluminum-lithium alloys and by their relatively low fracture toughness compared to other more conventional aluminum alloys. Lithium additions to aluminum alloys, however, provide a substantial lowering of the density as compared to conventional aluminum alloys ,which has been found to be very important in decreasing the overall structural weight of aircraft. Lithium additions are also effective in achieving a relatively high strength to weight ratio. While substantial strides have been made in improving the aluminum-lithium processing technology, a major challenge still outstanding is an ability to obtain a good blend of fracture toughness and high strength in an aluminum-lithium alloy. In Aluminum-Lithium Alloys II, Proceedings of the 2nd International Aluminum Conference, 12th-14th April 1983, pages 255-285, the alloy NOR 81 comprises a too high zirconium concentration (0.19% by wt) and further such a magnesium concentration (0.56% by wt), that this NOR 81 chemical composition falls outside the single phase region of Al-Li-Mg-Cu-alloys. Accordingly, the NOR 81 alloy would contain an appreciable volume percentage of solid intermetallic particles deteriorating the fracture toughness. Furthermore, the reported strong presence of the βAl₃Zr phase in the as-extruded and all aged conditions indicates that the zirconium content of the NOR 81 alloy is too rich in zirconium. These β-phase particles are detrimental to the fracture toughness. Finally, the presence of T₂ and/or R phase as well as the early development of the Δ phase provide additional objective indications that the NOR 81 alloy exceeded the single phase boundary. EP-A-157,600 (date of publication October 9, 1985) relates to an aluminum-lithium base alloy wrought product comprising 0.5 to 4.0 weight percent Li, 0 to 5.0 weight percent Mg, up to 5.0 weight percent Cu, 0 to 1.0 weight percent Zr, 0 to 2.0 weight percent Mn, 0 to 7.0 weight percent Zn, 0.5 weight percent max. Fe, 0.5 weight percent max. Si, the balance aluminum and incidental impurities. The product has imparted thereto prior to an aging step, a working effect equivalent to stretching in an amount greater than 3% in order to improve strength and fracture toughness combinations. Example II discloses an aluminum alloy consisting of by weight 2.0% Li, 2.7% Cu, 0.65% Mg and 0.12% Zr, the balance essentially aluminum and impurities. Inventive specimens stretched for 6% were compared with a single comparative example stretched for 2% and aged for 72 hours at 162°C (325°F). The present invention provides processes for manufacturing an article from a novel aluminum alloy composition that can be worked and heat treated so as to provide an aluminum-lithium alloy artical with high strength, good fracture toughness, and relatively low density compared to conventional aluminum alloys such as 7XXX and 2XXX series alloys that it is intended to replace. A first process according to the invention comprises the steps of: a) preparing an alloy of the following composition: element amount (% by wt) Li 2.0 to 2.4 Mg 0.3 to 0.9 Cu 2.1 to 2.9 Zr 0.08 to 0.15 Fe 0.15 max Si 0.12 max Zn 0.25 max Ti 0.15 max Cr 0.1 max other trace elements, each 0.05 max total of other trace elements 0.15 max Al Balance b) forming an article from said alloy; c) subjecting the article to a solution heat treatment; d) quenching the article in a quenching medium; and e) subjecting the quenched article to an aging treatment, and is chacacterized in that the aging treatment comprising aging to near peak strength at an aging temperature being in the range of 162°C (325°F) to 176°C (350°F), disclaiming an aluminum-lithium alloy consisting of, by weight, 2.0% Li, 2.7% Cu, 0.65% Mg and 0.12% Zr, the balance essentially aluminum and impurities, being 2% stretched and aged for 72 hours at 162°C (325°F). A second process according to the invention comprises the steps of: a) preparing an alloy of the following composition: element amount (% by wt) Li 2.0 to 2.4 Mg 0.3 to 0.9 Cu 2.1 to 2.9 Zr 0.08 to 0.15 Fe 0.15 max Si 0.12 max Zn 0.25 max Ti 0.15 max Cr 0.1 max other trace elements, each 0.05 max total of other trace elements 0.15 max Al Balance b) forming an article from said alloy; c) subjecting the article to a solution heat treatment; d) quenching the article in a quenching medium; and e) subjecting the quenched article to an aging treatment, and is characterized in that the aging treatment comprises naturally aging. An alloy prepared in accordance with the present invention has a nominal composition on the order of 2.2 weight percent lithium, 0.6 percent magnesium, 2, 5 percent copper and 0.12 percent zirconium. Artificial aging of the alloy at a temperature in the range of 162 to 176°C (325 to 350°F) to a near-peak age condition results in high strengths comparable to those of current 7XXX-T6 alloys in combination with good toughness and resistance to stress corrosion cracking. By underaging the alloy, strength and fracture toughness levels equivalent to or better than those of existing 2XXX-T3 type alloys are obtained. An aluminum-lithium alloy formulated in accordance with the present invention can contain from about 2.0 to about 2.4 percent lithium, 0.3 to 0.9 percent magnesium, 2.1 to 2,9 percent copper, and from about 0.08 to a maximum of 0.15 percent zirconium as a grain refiner. Preferably from about 0.09 to 0.14 percent zirconium is incorporated. All percentages herein are by weight percent based on the total weight of the alloy unless otherwise indicated. The magnesium is included to increase strength without increasing density. Preferred amounts of magnesium range from about 0.4 to 0.8 percent, with 0.6 percent being most preferred. The copper adds strength to the alloy. Iron and silicon can each be present in maximums up to a total of 0.3 percent. It is preferred that these impurities be present only in trace amounts, limiting the iron to a maximum of 0.15 percent and the silicon to a maximum of 0.12 percent, and preferably to maximums of 0.10 and 0.10 percent, respectively. The element zinc may be present in amounts up to but not exceeding 0.25 percent of the total. Titanium and chromium should not exceed 0.15 percent and 0.10 percent, respectively. Other elements such as manganese must each be held to levels of 0.05 percent or below, and the total amount of such other trace elements must be held to a maximum 0.15 percentage. If the foregoing maximums are exceeded the desired properties of the aluminum-lithium alloy will tend to deteriorate. The trace elements sodium and hydrogen are also thought to be harmful to the properties (fracture toughness in particular) of aluminum-lithium alloys and should be held to the lowest levels practically attainable, for example on the order of 15 to 30 ppm (0.0015-0.0030 wt.%) or less for the sodium and less than 15 ppm (0.0015 wt.%) and preferably less than 1.0 ppm (0.0001 wt. %) for the hydrogen. The balance of the alloy, of course, comprises aluminum. An aluminum-lithium alloy formulated in the proportions set forth in the foregoing two paragraphs is processed into an article utilizing known techniques. The alloy is formulated in molten form and cast into an ingot. The ingot is then homogenized at temperatures ranging from 496 to 543°C (925 to 1010°F) or higher. Thereafter, the alloy is converted into a usable article by conventional mechanical formation techniques such as rolling, extrusion, or the like. Once an article is formed, the alloy is normally subjected to a solution treatment at temperatures ranging from 510 to 543°C (950 to 1010°F), followed by quenching in a quenching medium such as water that is maintained at a temperature on the order of 21 to 65°C (70 to 150°F). If the alloy has been rolled or extruded, it is generally stretched on the order of 1 to 3 percent of its original length to relieve internal stresses, and to provide improved age-hardening response. The alumium alloy can then be further worked and formed by secondary operations into the various shapes for its final application. Additional heat treatments such as solution heat treatment and/or aging can be employed if desired after such forming operations. For example, sheet products after stretch forming to the desired shapes may be re-solution heat treated at a temperature on the order of 535°C (995°F) for 10 minutes to one hour. The article is normally then quenched in a quenching medium held at temperatures ranging from about 21 to 65°C (70 to 150°F). Thereafter, in accordance with the present invention, the alloy is subjected to an aging treatment at moderately low temperatures on the order of from 162 to 176°C (325 to 350°F). When this alloy is intended to replace conventional 7XXX series type alloys, the alloy can be aged for a period of time that will allow it to achieve near peak strength, preferably about 95 percent, and most preferably about 95 to 97 percent, of its peak strength. This level of strength can be achieved by aging for about 4 to 120 hours, and preferably for about 24 to 96 hours. When this alloy is intended to replace conventional 2XXX series alloys, the alloy is naturally aged, after quenching, for periods of four to seven days. Ultimate strength values of approximately 448 MPa (65 ksi) are developed in plate products together with yield strengths of about 379 MPa (55 ksi). In addition, outstanding fracture toughness and ductility properties are developed in both longitudinal and transverse grain directions.

Claims: A process of manufacturing a product of an aluminum-lithium alloy comprising the steps of: a) preparing an alloy of the following composition: element amount (% by wt) Li 2.0 to 2.4 Mg 0.3 to 0.9 Cu 2.1 to 2.9 Zr 0.08 to 0.15 Fe 0.15 max Si 0.12 max Zn 0.25 max Ti 0.15 max Cr 0.1 max other trace elements, each 0.05 max total of other trace elements 0.15 max Al Balance b) forming an article from said alloy; c) subjecting the article to a solution heat treatment; d) quenching the article in a quenching medium; and e) subjecting the quenched article to an aging treatment, characterized in that the aging treatment comprising aging to near peak strength at an aging temperature being in the range of 162°C (325°F) to 176°C (350°F), disclaiming an aluminum-lithium alloy consisting of, by weight, 2.0% Li, 2.7% Cu, 0.65% Mg and 0.12% Zr, the balance essentially aluminum and impurities, being 2% stretched and aged for 72 hours at 162°C (325°F). A process of manufacturing a product of an aluminum-lithium alloy comprising the steps of: a) preparing an alloy of the following composition: element amount (% by wt) Li 2.0 to 2.4 Mg 0.3 to 0.9 Cu 2.1 to 2.9 Zr 0.08 to 0.15 Fe 0.15 max Si 0.12 max Zn 0.25 max Ti 0.15 max Cr 0.1 max other trace elements, each 0.05 max total of other trace elements 0.15 max Al Balance b) forming an article from said alloy; c) subjecting the article to a solution heat treatment; d) quenching the article in a quenching medium; and e) subjecting the quenched article to an aging treatment, characterized in that the aging treatment comprises naturally aging. Process as claimed in claim 1, wherein the aging treatment comprises aging to about 95% peak strength. Process as claimed in claim 1 or 3, wherein the aging treatment comprises aging to about 95 to 97% peak strength. Process as claimed in claim 1, 3 or 4, wherein the aging treatment comprises aging for about 4 to 120 hours. Process as claimed in claim 1 or 3-5, wherein the aging treatment comprises aging for about 24 to 96 hours. Process as claimed in claim 2, wherein the aging treatment comprises naturally aging for time periods of 4 to 7 days. Process as claimed in claim 1-7, wherein the quenched article is subjected to stretching prior to the aging treatment. Process as claimed in claim 8, wherein the article is stretched in the order of 1 to 3%. Process as claimed in claim 1-9, wherein lithium is present in an amount ranging from about 2.0 to 2.2% by wt. Process as claimed in claim 1-10, wherein magnesium is present in an amount ranging from about 0.4 to 0.8% by wt. Process as claimed in claim 1-11, wherein magnesium is present in an amount of 0.6% by wt. Process as claimed in claim 1-12, wherein zirconium is present in an amount of 0.09 to 0.14% by wt. Process as claimed in claim 1-13, wherein an alloy is prepared having a nominal composition of 2.2% by wt lithium, 0.6% by wt magnesium, 2.5% by wt copper and 0.12% by wt zirconium.

Inventors: AXTER, SVEN E., CURTIS, R. EUGENE, HYATT, MICHAEL V., NARAYANAN, HARI G., QUIST, WILLIAM E.
Applicants: THE BOEING COMPANY
Application year: 1986
Main classification: C22C21
IPC codes: C22C21