CA1267797A - Aluminum-lithium alloy - Google Patents
Aluminum-lithium alloyInfo
- Publication number
- CA1267797A CA1267797A CA000507264A CA507264A CA1267797A CA 1267797 A CA1267797 A CA 1267797A CA 000507264 A CA000507264 A CA 000507264A CA 507264 A CA507264 A CA 507264A CA 1267797 A CA1267797 A CA 1267797A
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- alloy
- percent
- aged
- aluminum
- lithium
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Abstract
ALUMINUM-LITHIUM ALLOY
Abstract An aluminum-lithium alloy exhibiting good fracture toughness and relatively high strength has a nominal composition of 2.2 percent lithium, 0.6 percent magnesium, 2.5 percent copper, 0.12 percent zirconium with the balance being aluminum and trace elements.
Abstract An aluminum-lithium alloy exhibiting good fracture toughness and relatively high strength has a nominal composition of 2.2 percent lithium, 0.6 percent magnesium, 2.5 percent copper, 0.12 percent zirconium with the balance being aluminum and trace elements.
Description
1;~6~7~ 7 ALUMINUM-LITHIUM ALLOY
Back~round of the Invention The present invention relates to aluminum-lithium alloys and more particularly to an aluminum-lithium alloy composition with good ~racture toughness and high strength.
It has been estimated that current large commercial transport aircraft may be able to save from 15 10 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 15 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 20 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 30 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 35 1930's, however, that the metal skin wing design became standard, and firmly established metals, primarily aluminum alloys, as the major airframe structural 7';~9'7 material. Since that time, aixcraft structural materials have remained remarkably consistent with aluminum structural materials being used primarily in the wing, 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.
10 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. Nhile composites will be used in increasecl 15 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 25 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 30 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 35 aluminum-lithium alloy.
12f~7'~9~7 Summary of the Invention The present invention provides a novel aluminum alloy composition that can be worked and heat treated so as to provide an aluminum-lithium alloy with high strength, good fracture toughness, and relatively low density compared to conventional aluminum alloys such as 7XXX and 2XX~ series alloys that it is intended to replace. An alloy prepared in accordance with the present 10 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 250 to 350F
to a near-peak age condition results in high strengths lS comparable to those of current 7XX~-T6 alloys in combination with good toughness and resistance to stress corrosion cracking. By underaging the alloy, strength and fracture toughness levels equivalen~ to or better than those of existing 2XXX-T3 type alloys are obtained.
20 Considerable improvement in the already excellent combination of fracture toughness and high strength is achieved for both 7XXX and 2XXX applications by underaging the alloy within the relatively low temperature range of 250 to 300F.
Detailed Description of the Invention An aluminum-lithium alloy formulated in accordance with the present invention can contain from about 2.0 to 30 about 2.4 percent lithium, 0.3 to 0.9 percent magnesium,
Back~round of the Invention The present invention relates to aluminum-lithium alloys and more particularly to an aluminum-lithium alloy composition with good ~racture toughness and high strength.
It has been estimated that current large commercial transport aircraft may be able to save from 15 10 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 15 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 20 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 30 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 35 1930's, however, that the metal skin wing design became standard, and firmly established metals, primarily aluminum alloys, as the major airframe structural 7';~9'7 material. Since that time, aixcraft structural materials have remained remarkably consistent with aluminum structural materials being used primarily in the wing, 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.
10 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. Nhile composites will be used in increasecl 15 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 25 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 30 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 35 aluminum-lithium alloy.
12f~7'~9~7 Summary of the Invention The present invention provides a novel aluminum alloy composition that can be worked and heat treated so as to provide an aluminum-lithium alloy with high strength, good fracture toughness, and relatively low density compared to conventional aluminum alloys such as 7XXX and 2XX~ series alloys that it is intended to replace. An alloy prepared in accordance with the present 10 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 250 to 350F
to a near-peak age condition results in high strengths lS comparable to those of current 7XX~-T6 alloys in combination with good toughness and resistance to stress corrosion cracking. By underaging the alloy, strength and fracture toughness levels equivalen~ to or better than those of existing 2XXX-T3 type alloys are obtained.
20 Considerable improvement in the already excellent combination of fracture toughness and high strength is achieved for both 7XXX and 2XXX applications by underaging the alloy within the relatively low temperature range of 250 to 300F.
Detailed Description of the Invention An aluminum-lithium alloy formulated in accordance with the present invention can contain from about 2.0 to 30 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 35 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, 1;26~797 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 10 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 15 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 20 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 25 (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 30 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 925 to 1010F or higher. Thereafter, the alloy is converted into a usable article by conventional 35 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 950 to 1010F, followed by 1~ ~7.--1~ ~7 5~
quenching in a quenching medium such as water that is maintained at a ~emperature on the order of 70 to 15QF.
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 aluminum alloy can then be further worked and formed by secondary operations into the various shapes 10 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 15 temperature on the order of 995F for 10 minutes to one hour. The article is normally then quenched in a quenching medium held at temperatures ranging from ahout 70 to 150F.
Thereafter, in accordance with the present invention, the alloy is subjected to an aging treatment at moderately low temperatures on the order of from 250 to 350F.
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.
30 Preferred aging temperatures for this purpose range from 275 to 325F. Within these temperature ranges, 95 to 97 percent peak age 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 can be aged to achieve moderately high strength in conjunction with high fracture toughness. Preferred aging temperatures for this -6- ~ Z ~ 7'79~
purpose range from about 250 to 300F, more preferably from about 250 to 275F, and most preferably on the order of 25nF. Within these temperature ranges, moderately high strength in conjunction with high fracture toughness can be achieved by aging for about 4 to 48 hours, and preferably about 4 to 24 hours.
An alternative way to achieve 2XXX series alloy properties is to naturally age the alloy, after quenching, 10 for periods of four to seven days. Ultimate strength values of approximately 65 ksi are developed in plate products together with yield strengths of about 55 ksi.
In addition, outstanding fracture toughness and ductility properties are developed in both longitudinal and 15 transverse grain directions.
Example The following example is presented to illustrate 20 the significantly improved and unexpected characteristics of an aluminum-lithium alloy formulated and manufactured in accordance with the parameters of the present invention. Moreover, it is intended to illustrate the superior characteristics of this aluminum-lithium alloy 25 when aged in accordance with the present invention and to assist one of ordinary skill in making and using the present invention. The following example is not intended in any way to otherwise limit the scope of this disclosure or the protection granted by Letters Patent hereon.
An aluminum alloy containing 2.2 percent lithium, 0.62 percent magnesium, 2.5 percent copper, 0.09 percent zirconium with the balance being aluminum was formulated. The trace elements present in the formulation 35 constituted less than 0.25 percent of the total. The iron and silicon present in the formulation constituted 0.10 percent each. The alloy was cast and homogenized at about 975F. Thereafter, the alloy was extruded to a thickness 1~7'797 of 0.75 inches. The resulting extrusion was then solution treated at about 975F for about 90 minutes. The extrusion was then quenched in water maintained at about 70F. Thereafter, the extrusion was subjected to a stretch Gf about 1 1/2 percent of its initial length. The material was then cut into specimens for fracture toughness (precracked Charpy impact test) and tensile strength testing. The precracked Charpy impact test specimens were machined to a final specimen size of 0.394 10 X 0.394 X 2.16 inches. The specimens prepared for tensile strength tests were standard round specimens having a gauge section diameter of 0.25 inches. A plurality of specimens were then artificially aged for up to 72 hours at 300F to a near-peak age condition. In addition, a 15 second set of tensile and Charpy specimens were aged at 275F for eight hours to a considerably underaged condition. The specimens were then subjected to tensile strength and precracked Charpy impact tests in accordance with standard A~TM and industry testing procedures. The 20 specimens aged at 300F to a near-peak age condition exhibited ultimate strengths ranging on the order of 82 to 92 ksi with fracture toughness values on the order of 220 to 350 in-lbs/in2. By comparison, the specimens underaged at 275 developed a very high fracture toughness, on the 25 order of 650 to 850 in-lbs/in2, in combination with ultimate strengths in the range of 70 to 75 ksi.
The present invention has been described in relation to various embodiments, including the preferred 30 formulation and processing parameters. One of ordinary skill after reading the foregoing specification will be able to effect various changes, substitutions of equivalents, and other alterations without departing from the broad concepts set forth herein. It is therefore 35 intended that the scope of the Letters Patent granted hereon will be limited only by the definition contained in the appended claims and equivalents thereof.
Preferably from about 0.09 to 0.14 percent zirconium is incorporated. All percentages herein are by weight 35 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, 1;26~797 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 10 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 15 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 20 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 25 (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 30 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 925 to 1010F or higher. Thereafter, the alloy is converted into a usable article by conventional 35 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 950 to 1010F, followed by 1~ ~7.--1~ ~7 5~
quenching in a quenching medium such as water that is maintained at a ~emperature on the order of 70 to 15QF.
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 aluminum alloy can then be further worked and formed by secondary operations into the various shapes 10 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 15 temperature on the order of 995F for 10 minutes to one hour. The article is normally then quenched in a quenching medium held at temperatures ranging from ahout 70 to 150F.
Thereafter, in accordance with the present invention, the alloy is subjected to an aging treatment at moderately low temperatures on the order of from 250 to 350F.
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.
30 Preferred aging temperatures for this purpose range from 275 to 325F. Within these temperature ranges, 95 to 97 percent peak age 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 can be aged to achieve moderately high strength in conjunction with high fracture toughness. Preferred aging temperatures for this -6- ~ Z ~ 7'79~
purpose range from about 250 to 300F, more preferably from about 250 to 275F, and most preferably on the order of 25nF. Within these temperature ranges, moderately high strength in conjunction with high fracture toughness can be achieved by aging for about 4 to 48 hours, and preferably about 4 to 24 hours.
An alternative way to achieve 2XXX series alloy properties is to naturally age the alloy, after quenching, 10 for periods of four to seven days. Ultimate strength values of approximately 65 ksi are developed in plate products together with yield strengths of about 55 ksi.
In addition, outstanding fracture toughness and ductility properties are developed in both longitudinal and 15 transverse grain directions.
Example The following example is presented to illustrate 20 the significantly improved and unexpected characteristics of an aluminum-lithium alloy formulated and manufactured in accordance with the parameters of the present invention. Moreover, it is intended to illustrate the superior characteristics of this aluminum-lithium alloy 25 when aged in accordance with the present invention and to assist one of ordinary skill in making and using the present invention. The following example is not intended in any way to otherwise limit the scope of this disclosure or the protection granted by Letters Patent hereon.
An aluminum alloy containing 2.2 percent lithium, 0.62 percent magnesium, 2.5 percent copper, 0.09 percent zirconium with the balance being aluminum was formulated. The trace elements present in the formulation 35 constituted less than 0.25 percent of the total. The iron and silicon present in the formulation constituted 0.10 percent each. The alloy was cast and homogenized at about 975F. Thereafter, the alloy was extruded to a thickness 1~7'797 of 0.75 inches. The resulting extrusion was then solution treated at about 975F for about 90 minutes. The extrusion was then quenched in water maintained at about 70F. Thereafter, the extrusion was subjected to a stretch Gf about 1 1/2 percent of its initial length. The material was then cut into specimens for fracture toughness (precracked Charpy impact test) and tensile strength testing. The precracked Charpy impact test specimens were machined to a final specimen size of 0.394 10 X 0.394 X 2.16 inches. The specimens prepared for tensile strength tests were standard round specimens having a gauge section diameter of 0.25 inches. A plurality of specimens were then artificially aged for up to 72 hours at 300F to a near-peak age condition. In addition, a 15 second set of tensile and Charpy specimens were aged at 275F for eight hours to a considerably underaged condition. The specimens were then subjected to tensile strength and precracked Charpy impact tests in accordance with standard A~TM and industry testing procedures. The 20 specimens aged at 300F to a near-peak age condition exhibited ultimate strengths ranging on the order of 82 to 92 ksi with fracture toughness values on the order of 220 to 350 in-lbs/in2. By comparison, the specimens underaged at 275 developed a very high fracture toughness, on the 25 order of 650 to 850 in-lbs/in2, in combination with ultimate strengths in the range of 70 to 75 ksi.
The present invention has been described in relation to various embodiments, including the preferred 30 formulation and processing parameters. One of ordinary skill after reading the foregoing specification will be able to effect various changes, substitutions of equivalents, and other alterations without departing from the broad concepts set forth herein. It is therefore 35 intended that the scope of the Letters Patent granted hereon will be limited only by the definition contained in the appended claims and equivalents thereof.
Claims (15)
1. An aluminum-lithium alloy exhibiting good fracture toughness consisting essentially of:
2. The alloy of Claim 1 wherein said zirconium is present in amounts from about 0.09 to 0.14 percent.
3. The alloy of Claim 1 wherein said magnesium is present in an amount ranging from about 0.4 to 0.8 percent.
4. The alloy of Claim 1 having a nominal composition of 2.2 percent lithium, 0.6 percent magnesium, 2.5 percent copper, and 0.12 percent zirconium.
5. The alloy of Claim 1 wherein said alloy has been aged at a moderately low temperature to near peak strength.
6. The alloy of Claim 1 wherein said alloy has been aged at a temperature in the range of from 250° to 350°F.
7. The alloy of Claim 6 aged at a temperature in the range of from 275° to 325°F.
8. The alloy of Claim 6 aged for 4 to 120 hours.
9. The alloy of Claim 7 aged for 24 to 96 hours.
10. The alloy of Claim 6 aged at a temperature in the range of 250° to 300°F.
11. The alloy of Claim 10 aged at a temperature in the range of 250° to 275°F.
12. The alloy of Claim 11 aged at a temperature on the order of 250°F.
13. The alloy of Claim 10 aged for about 4 to 48 hours.
14. The alloy of Claim 13 aged for about 4 to 24 hours.
15. The alloy of Claim 1 wherein said alloy has been naturally aged.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA000507264A CA1267797A (en) | 1986-04-22 | 1986-04-22 | Aluminum-lithium alloy |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CA000507264A CA1267797A (en) | 1986-04-22 | 1986-04-22 | Aluminum-lithium alloy |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1267797A true CA1267797A (en) | 1990-04-17 |
Family
ID=4132937
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA000507264A Expired - Fee Related CA1267797A (en) | 1986-04-22 | 1986-04-22 | Aluminum-lithium alloy |
Country Status (1)
| Country | Link |
|---|---|
| CA (1) | CA1267797A (en) |
-
1986
- 1986-04-22 CA CA000507264A patent/CA1267797A/en not_active Expired - Fee Related
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