Aluminum alloy products having improved property combinations and method for artificially aging same
10450640 ยท 2019-10-22
Assignee
Inventors
- Dhruba J. Chakrabarti (Woodside, NY, US)
- John Liu (Murrysville, PA, US)
- Jay H. Goodman (Murrysville, PA, US)
- Gregory B. Venema (Bettendorf, IA, US)
- Ralph R. Sawtell (Gibsonia, PA)
- Cynthia M. Krist (Bettendorf, IA, US)
- Robert W. Westerlund (Bettendorf, IA)
Cpc classification
C22F1/053
CHEMISTRY; METALLURGY
International classification
B22D17/22
PERFORMING OPERATIONS; TRANSPORTING
C22F1/053
CHEMISTRY; METALLURGY
Abstract
Aluminum alloy products, such as plate, forgings and extrusions, suitable for use in making aerospace structural components like integral wing spars, ribs and webs, comprises about: 6 to 10 wt. % Zn; 1.2 to 1.9 wt. % Mg; 1.2 to 2.2 wt. % Cu, with Mg(Cu+0.3); and 0.05 to 0.4 wt. % Zr, the balance Al, incidental elements and impurities. Preferably, the alloy contains about 6.9 to 8.5 wt. % Zn; 1.2 to 1.7 wt. % Mg; 1.3 to 2 wt. % Cu. This alloy provides improved combinations of strength and fracture toughness in thick gauges. When artificially aged per the 3-stage method of preferred embodiments, this alloy also achieves superior SCC performance, including under seacoast conditions.
Claims
1. A method comprising: (a) casting an alloy that consists essentially of: 6.9 to 9 wt. % Zn; 1.3 to 1.68 wt. % Mg; 1.2 to 1.9 wt. % Cu, with wt. % Mg(wt. % Cu+0.3), wherein wt. % Cu+wt. % Mg3.5 wt. %; 0.05 to 0.3 wt. % Zr, up to 0.3 wt. % Mn and up to 0.1 wt, % Cr, wherein at least one of the Mn and the Cr is purposefully included in the alloy, the balance Al, incidental elements and impurities; (b) homogenizing and then hot forming the alloy into a workpiece by one or more methods selected from the group consisting of: rolling, extruding and forging; (i) wherein the homogenizing comprises heating to a first temperature and then holding at the first temperature for at least for 4 hours, and then heating to a second higher temperature and holding at the second temperature for at least 4 hours; (ii) wherein the first temperature is at least 800 F.; (iii) wherein the second temperature is at least 890 F.; (c) solution heat treating the workpiece; (d) quenching the solution heat treated workpiece; and (e) artificially aging the quenched workpiece, wherein, after the artificial aging, the workpiece is unrecrystallized having less than 50% recrystallization.
2. The method of claim 1, comprising: age forming the workpiece into a structural component shape.
3. The method of claim 1 wherein the quenched workpiece is 3 to 12 inches at its thickest cross sectional point.
4. The method of claim 1 wherein the alloy contains less than 8 wt. % Zn and less than 1.8 wt. % Cu.
5. The method of claim 1 wherein artificial aging step (e) comprises: (i) a first aging stage within 200 to 275 F.; and (ii) a second aging stage within 300 to 335 F.
6. The method of claim 1 wherein artificial aging step (e) comprises: (i) a first aging stage within 200 to 275 F.; (ii) a second aging stage within 300 to 335 F.; and (iii) a third aging stage within 200 to 275 F.
7. The method of claim 6 wherein the first aging stage (i) proceeds within 230 to 260 F.
8. The method of claim 6 wherein the first aging stage (i) proceeds for 2 to 12 hours.
9. The method of claim 6 wherein the first aging stage (i) proceeds for 6 or more hours within 235 to 255 F.
10. The method of claim 6 wherein the second aging stage (ii) proceeds for 4 to 18 hours within 310 to 325 F.
11. The method of claim 10 wherein the second aging stage (ii) proceeds for 6 to 15 hours within 300 to 315 F.
12. The method of claim 10 wherein the second aging stage (ii) proceeds for 7 to 13 hours within 310 to 325 F.
13. The method of claim 6 wherein the third aging stage (iii) proceeds within 230 to 260 F.
14. The method of claim 6 wherein one or more of the first, second and third aging stages includes an integration of multiple temperature aging effects.
Description
DESCRIPTION OF THE DRAWINGS
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PREFERRED EMBODIMENTS
(15) Mechanical properties of importance for the thick plate, extrusion or forging for aircraft structural products, as well as other non-aircraft structural applications, include strength, both in compression as for the upper wing skin and in tension for the lower wing skin. Also important are fracture toughness, both plane-strain and plane-stress, and corrosion resistance performance such as exfoliation and stress corrosion cracking resistance, and fatigue, both smooth and open-hole fatigue life (S/N) and fatigue crack growth (FCG) resistance.
(16) As described above, integral wing spars, ribs, webs, and wing skin panels with integral stringers, can be machined from thick plates or other extruded or forged product forms which have been solution heat treated, quenched, mechanically stress relieved (as needed) and artificially aged. It is not always feasible to solution heat treat and rapidly quench the finished structural component itself because the rapid cooling from quenching may induce residual stress and cause dimensional distortions. Such quench-induced residual stresses can also cause stress corrosion cracking. Likewise, dimensional distortions due to rapid quenching may necessitate re-working to straighten parts that have become so distorted as to render standard assembly impracticably difficult. Other representative aerospace parts/products that can be made from this invention include, but are not limited to: large frames and fuselage bulkheads for commercial jet airliners, hog out plates for the upper and lower wing skins of smaller, regional jets, landing gear and floor beams for various jet aircraft, even the bulkheads, fuselage components and wing skins of fighter plane models. In addition, the alloy of this invention can be made into miscellaneous small forged parts and other hogged out structures of aircraft that are currently made from alloy 7050 or 7010 aluminum.
(17) While it is easier to obtain better mechanical properties in thin cross sections (because the faster cooling of such parts prevents unwanted precipitation of alloying elements), rapid quenching can cause excessive quench distortion. To the extent practical, such parts may be mechanically straightened and/or flattened while residual stress relief practices are performed thereon after which these parts are artificially aged.
(18) As indicated above, in solution heat treating and quenching thick sections, the quench sensitivity of the aluminum alloy is of great concern. After solution heat treating, it is desirable to quickly cool the material for retaining various alloying elements in solid solution rather than allowing them to precipitate out of solution in coarse form as otherwise occurs via slow cooling. The latter occurrence produces coarse precipitates and results in a decline in mechanical properties. In products with thick cross sections, i.e. over 2 inches thick at its greatest point, and more particularly, about 4 to 8 inches thick or more, the quenching medium acting on exterior surfaces of such workpieces (either plate, forging or extrusion) cannot efficiently extract heat from the interior including the center (or mid-plane (T/2)) or quarter-plane (T/4) regions of that material. This is due to the physical distance to the surface and the fact that heat extracts through the metal by a distance dependent conduction. In thin product cross sections, quench rates at the mid-plane are naturally higher than quench rates for a thicker product cross sections. Hence, an alloy's overall quench sensitivity property is often not as important in thinner gauges as it is for thicker gauged parts, at least from the standpoint of strength and toughness.
(19) The present invention is primarily focused on increasing the strength-toughness properties in a 7XXX series aluminum alloy in thicker gauges, i.e. greater than about 1.5 inches. The low quench sensitivity of the invention alloy is of extreme importance. In thicker gauges, the less quench sensitivity the better with respect to that material's ability to retain alloying elements in solid solution (thus avoiding the formation of adverse precipitates, coarse and others, upon slow cooling from SHT temperatures) particularly in the more slowly cooling mid- and quarter-plane regions of said thick workpiece. This invention achieves its desired goal of lowering quench sensitivity by providing a carefully controlled alloy composition which permits quenching thicker gauges while still achieving superior combinations of strength-toughness and corrosion resistance performance.
(20) To illustrate the invention, twenty-eight, 11-inch diameter ingots were direct chill (or DC) cast, homogenized and extruded into 1.254 inch wide rectangular bars. Those bars were all solution heat treated before being quenched at different rates to simulate cooling conditions for thin sections as well as for approximating conditions for the mid-plane of 6- and 8-inch thick workpiece sections. These rectangular test bars were then cold stretched by about 1.5% for residual stress relief. The compositions of alloys studied are set forth in Table 2 below, in which Zn contents ranged from about 6.0 wt. % to slightly in excess of 11.0 wt. %. For these same test specimens, Cu and Mg contents were each varied between about 1.5 and 2.3 wt. %.
(21) TABLE-US-00002 TABLE 2 SAMPLE Invention Alloy Composition (wt. %) No. Y/N Cu Mg Zn 1 Y 1.57 1.55 6.01 2 N 1.64 2.29 5.99 3 N 2.45 1.53 5.86 4 N 2.43 2.26 6.04 5 N 1.95 1.94 6.79 6 Y 1.57 1.51 7.56 7 N 1.59 2.30 7.70 8 N 2.45 1.54 7.71 9 N 2.46 2.31 7.70 10 N 2.05 1.92 8.17 11 Y 1.53 1.52 8.65 12 N 1.57 2.35 8.62 13 N 2.32 1.45 8.25 14 N 2.04 2.19 8.33 15 N 1.86 1.93 10.93 16 N 1.98 2.09 11.28 17 N 1.97 1.86 9.04 18 Y 1.48 1.50 9.42 19 N 1.75 2.29 9.89 20 N 2.48 1.52 9.60 21 N 2.19 2.19 9.74 22 N 1.68 1.55 11.38 23 N 1.65 2.28 11.04 24 N 2.38 1.53 11.08 25 N 2.22 1.97 9.04 26 N 1.79 2.00 10.17 27 N 2.23 2.28 6.62 28 N 2.48 1.98 8.31 For all alloys other than the controls: Target Si = 0.03, Fe = 0.05, Zr = 0.12, Ti = 0.025 For 7150 Control (Sample # 27): Target Si = 0.05, Fe = 0.10, Zr = 0.12, Ti = 0.025 For 7055 Control (Sample # 28): Target Si = 0.07, Fe = 0.11, Zr = 0.12, Ti = 0.025
(22) Different quenching approaches were explored to obtain, at the mid-plane of a 1.25-inch thick extruded bar, a cooling rate simulating that at the mid-plane of a 6-inch thick plate spray quenched in 75 F. water as would be the case in full-scale production. A second set of data involved simulating, under identical circumstances, a bar cooling rate corresponding to that of an 8-inch thick plate.
(23) The aforesaid quenching simulation involved modifying the heat transfer characteristics of quenching medium, as well as the part surface, by immersion quenching extruded bars via the simultaneous incorporation of three known quenching practices: (i) a defined warm water temperature quench; (ii) saturation of the water with CO.sub.2 gas; and (iii) chemically treating the bars to render a bright etch surface finish to lower surface heat transfer.
(24) For simulating the 6-inch thick plate cooling condition: the water temperature for immersion quenching was held at about 180 F.; and the solubility level of CO.sub.2 in the water kept at about 0.20 LAN (a measure of dissolved CO.sub.2 concentration, LAN=standard volume of CO.sub.2/volume of water). Also, the sample surface was chemically treated to have a standard, bright etch finish.
(25) For the 8-inch thick plate cooling simulation, the water temperature was raised to about 190 F. with a CO.sub.2 solubility reading varying between 0.17 and 0.20 LAN. Like the 6-inch samples above, this thicker plate was chemically treated to have a standard bright etch surface finish.
(26) The cooling rates were measured by thermocouples inserted into the mid-plane of each bar sample. For benchmark reference, the two calculated cooling curves to approximate the mid-plane cooling rates under spray quenching at plant-made 6- and 8-inch thick plates were plotted per accompanying
(27) After solution heat treating and quenching, artificial aging behaviors were studied using multiple aging times to obtain acceptable electrical conductivity (EC) and exfoliation corrosion resistance (EXCO) readings. The first two-step aging practice for the invention alloy consisted of: a slow heat-up (for about 5 to 6 hours) to about 250 F., a 4- to 6-hour soak at about 250 F., followed by a second step aging at about 320 F. for varying times ranging from about 4 to 36 hours.
(28) Tensile and compact tension plane-strain fracture toughness test data were then collected on samples given the different minimum aging times required to obtain a visual EXCO rating of EB or better (EA or pitting only) for acceptable exfoliation corrosion resistance performance, and an electrical conductivity EC minimum value of at or above about 36% IACS (International Annealed Copper Standard), the latter value being used to indicate degree of necessary over-aging and provide some indication of corrosion resistance performance enhancement as is known in the art. All tensile tests were performed according to the ASTM Specification E8, and all plane-strain fracture toughness per ASTM specification E399, said specifications being well known in the art.
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(30) Substantial improvements in strength and toughness can also be seen when the aforementioned alloy performances are compared against two control alloys 7150 aluminum (Sample #27 above) and 7055 aluminum (Sample #28) both of which were processed in an identical manner (including temper). In
(31) Also included in the
(32) A similar set of results are graphically depicted in accompanying
(33) Thus, unlike past teachings, some of the highest strength-toughness properties were obtained at some of the leanest Cu and Mg levels used thus far for current commercial aerospace alloys. Concomitantly, the Zn levels at which these properties were most optimized correspond to levels much higher than those specified for 7050, 7010 or 7040 aluminum plate products.
(34) It is believed that a good portion of the improvement in strength and toughness properties observed for thick sections of the invention alloy are due to the specific combination of alloy ingredients. For instance, the accompanying
(35) In accompanying
(36) Also noteworthy is the performance of Sample #7 above, which according to Table 2 contained 1.59 wt. % Cu, 2.30 wt. % Mg and 7.70 wt. % Zn, (so that its Mg content exceeded Cu content). From
(37) It is desirable to achieve optimum and/or balanced fracture toughness (K.sub.Q) and strength (TYS) properties in the alloys of this invention. As can be best seen and appreciated by comparing the compositions of Table 2 with their corresponding fracture toughness and strength values plotted in
(38) The upper limit of Zn content appears to be important in achieving the proper balance between toughness and strength properties. Those samples which exceeded about 11.0 wt. %, such as Sample Nos. 24 (11.08 wt. % Zn) and 22 (11.38 wt. % Zn), failed to achieve the minimum combined strength and fracture toughness levels set forth above for alloys of the invention.
(39) The preferred alloy compositions herein thus provide high damage tolerance in thick aerospace structures resulting from its enhanced, combined fracture toughness and yield strength properties. With respect to some of the property values reported herein, one should note that K.sub.Q values are the result of plane strain fracture toughness tests that do not conform to the current validity criteria of ASTM Standard E399. In the current tests that yield K.sub.Q values, the validity criteria that were not precisely followed were: (1) P.sub.MAX/P.sub.Q<1.1 primarily, and (2) B (thickness)>2.5 (K.sub.Q/.sub.YS).sup.2 occasionally, where K.sub.Q, .sub.YS, P.sub.MAX, and P.sub.Q are as defined in ASTM Standard E399-90. These differences are a consequence of the high fracture toughnesses observed with the invention alloy. To obtain valid plane-strain K.sub.Ic results, a thicker and wider specimen would have been required than is facilitated with an extruded bar (1.25-inch thick4-inch wide). A valid K.sub.Ic is generally considered a material property relatively independent of specimen size and geometry. K.sub.Q, on the other hand, may not be a true material property in the strictest academic sense because it can vary with specimen size and geometry. Typical K.sub.Q values from specimens smaller than needed are conservative with respect to K.sub.Ic however. In other words, reported fracture toughness (K.sub.Q) values are generally lower than standard K.sub.Ic values obtained when the sample size related, validity criteria of ASTM Standard E399-90 are satisfied. The K.sub.Q values were obtained herein using compact tension test specimens per ASTM E399 having a thickness B of 1.25 inches and a width that varied between 2.5 to 3.0 inches for different specimens. Those specimens were fatigue pre-cracked to a crack length A of 1.2 to 1.5 inch (A/W=0.45 to 0.5). The tests on plant trial material, discussed below, which did satisfy the validity criterion of ASTM Standard E399 for K.sub.Ic were conducted using compact tension specimens with a thickness, B=2.0 inch, and width, W=4.0 inch. Those specimens were fatigue pre-cracked to a crack length of 2.0 inch (A/W=0.5). All cases of comparative data between varying alloy compositions were made using results from specimens of the same size and under similar test conditions.
Example 1: Plant TrialPlate
(40) A plant trial was conducted using a standard, full-size ingot cast with the following invention alloy composition: 7.35 wt. % Zn, 1.46 wt. % Mg, 1.64 wt. % Cu, 0.04 wt. % Fe, 0.02 wt. % Si and 0.11 wt. % Zr. That ingot was scalped, homogenized at 8850 to 890 f for 24 hours, and hot rolled to 6-inch thick plate. The rolled plate was then solution heat treated at 885 to 890 F. for 140 minutes, spray quenched to ambient temperature, and cold stretched from about 1.5 to 3% for residual stress relief. Sections from that plate were subjected to a two-step aging practice that consisted of a 6-hour/250 f first step aging followed by a second step aging at 320 F. for 6, 8 and 11 hours, respectively designated as times T1, T2 and T3 in the table that follows. Results from the tensile, fracture toughness, alternate immersion SCC, EXCO and electrical conductivity tests are presented in Table 3 below.
(41) TABLE-US-00003 TABLE 3 Properties of Plant Processed, 6-inch Thick Plate Samples of the Invention Alloy SCC Stress Aging (ASTM G44) Time at L-UTS L-TYS EL L-CYS L-T K.sub.IC EC (20 d-Pass) 320 F. (T/4) (T/4) (T/4) (T/4) (T/4) EXCO (T/4) (T/2) (Hrs.) (ksi) (ksi) (%) (ksi) (ksiin) (T/4) (% IACS) (ksi) 6 (T1) 77.1 74.9 6.8 73.2 33.6 EB 40.5 35 8 (T2) 75.6 72.5 7.3 71.0 35.2 EB 41.3 40 11 (T3) 71.9 67.2 8.6 65.6 40.5 EA 42.7 45
Example 2: Plant TrialForging
(42) A die forged evaluation of the invention alloy was performed in a plant-trial using two full-size production sheet/plate ingots, designated COMP 1 and COMP2, as follows: COMP 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 wt. % Ti; COMP 2: 7.39 wt. % Zn, 1.48 wt. % Mg, 1.91 wt. % Cu, 0.11 wt. % Zr, 0.036 wt. % Fe, 0.024 wt. % Si, 0.02 wt. % Ti.
A standard 7050 ingot was also run as a control. All of the aforesaid ingots were homogenized at 885 F. for 24 hours and sawed to billets for forging. A closed die, forged part was produced for evaluating properties at three different thicknesses, 2-inch, 3-inch and 7-inch. The fabrication steps conducted on these metals included: two pre-forming operations utilizing hand forging, followed by a blocker die operation and a final finish die operation using a 35,000-ton press. The forging temperatures employed therefore were between about 725-750 F. All the forged pieces were then solution heat treated at 880 to 890 F. for 6 hours, quenched and cold worked 1 to 5% for residual stress relief. The parts were next given a T74 type aging treatment for enhancing SCC performance. The aging treatment consisted of 225 F. for 8 hours, followed by 250 F. for 8 hours, then 350 F. for 8 hours. Results from the tensile tests performed in longitudinal, long-transverse and short-transverse directions are presented in accompanying
(43) The present invention clearly runs counter to conventional 7XXX series alloy design philosophies which indicate that higher Mg contents are desirable for high strength. While that may still be true for thin sections of 7XXX aluminum, it is not the case for thicker product forms because higher Mg actually increases quench sensitivity and reduces the strength of thick sections.
(44) Although the primary focus of this invention was on thick cross sectioned product quenched as rapidly as practical, those skilled in the art will recognize and appreciate that another application hereof would be to take advantage of the invention's low quench sensitivity and use an intentionally slow quench rate on thin sectioned parts to reduce the quench-induced residual stresses therein, and the amount/degree of distortion brought on by rapid quenching but without excessively sacrificing strength or toughness.
(45) Another potential application arising from the lower quench sensitivities observed with this invention alloy is for products having both thick and thin sections such as die forgings and certain extrusions. Such products should suffer less from yield strength differences between thick and thin cross sectioned areas. That, in turn, should reduce the chances of bowing or distortion after stretching.
(46) Generally, for any given 7XXX series alloy, as further artificial aging is progressively applied to a peak strength, T6-type tempered product (i.e. overaging), the strength of that product has been known to progressively and systematically decrease while its fracture toughness and corrosion resistance progressively and systematically increase. Hence, today's part designers have learned to select a specific temper condition with a compromise combination of strength, fracture toughness and corrosion resistance for a specific application. Indeed, such is the case for the alloy of the invention, as demonstrated in the cross plot of L-T plane strain fracture toughness K.sub.Ic and L tensile yield strength, in
(47) It is further understood by those skilled in the art that, within limits, for a specific 7XXX series alloy, the strength-fracture toughness trend line can be interpolated and, to some extent, extrapolated to combinations of strength and fracture toughness beyond the three examples of invention alloy given above and plotted at
(48) While the invention has been described largely with respect to aerospace structural applications, it is to be understood that its end use applications are not necessarily limited to same. On the contrary, the invention alloy and its preferred 3-stage aging practice herein are believed to have many other, non-aerospace related end use applications as relatively thick cast, rolled plate, extruded or forged product forms, especially in applications that would require relatively high strengths in a slowly quenched condition from SHT temperatures. An example of one such application is mold plate, which must be extensively machined into molds of various shapes for the shaping and/or contouring processes of numerous other manufacturing processes. For such applications, desired material characteristics are both high strength and low machining distortion. When using 7XXX alloys as mold plates, a slow quench after solution heat treatment would be necessary to impart a low residual stress, which might otherwise cause machining distortions. Slow quenching also results in lowered strength and other properties for existing 7XXX series alloys due to their higher quench sensitivity. It is the unique very low quench sensitivity for this invention alloy that permits a slow quench following SHT while still retaining relatively high strength capabilities that makes this alloy an attractive choice for such non-aerospace, non-structural applications as thick mold plate. For this particular application, though, it is not necessary to perform the preferred 3-step aging method described hereinbelow. Even a single step, or standard 2-step, aging practice should suffice. The mold plate can even be a cast plate product.
(49) The instant invention substantially overcomes the problems encountered in the prior art by providing a family of 7000 Series aluminum alloy products which exhibits significantly reduced quench sensitivity thus providing significantly higher strength and fracture toughness levels than heretofore possible in thick gauge aerospace parts or parts machined from thick products. The aging methods described herein then enhance the corrosion resistance performance of such new alloys. Tensile yield strength (TYS) and electrical conductivity EC measurements (as a % IACS) were taken on representative samples of several new 7XXX alloy compositions and comparative aging processes practiced on the present invention. The aforesaid EC measurements are believed to correlate with actual corrosion resistance performance, such that the higher the EC value measured, the more corrosion resistant that alloy should be. As an illustration, commercial 7050 alloy is produced in three increasingly corrosion resistant tempers: T76 (with a typical SCC minimum performance, or guarantee, of about 25 ksi and typical EC of 39.5% IACS); T74 (with a typical SCC guarantee of about 35 ksi and 40.5% IACS); and T73 (with it typical SCC guarantee of about 45 ksi and 41.5% IACS).
(50) In aerospace, marine or other structural applications, it is quite customary for a structural and materials engineer to select materials for a particular component based on the weakest link failure mode. For example, because the upper wing alloy of an aircraft is predominantly subjected to compressive stresses, it has relatively lower requirements for SCC resistance involving tensile stresses. As such, upper wing skin alloys and tempers are usually selected for higher strength albeit with relatively low short-transverse SCC resistance. Within that same aerospace wing box, the spar members are subjected to tensile stresses. Although the structural engineer would desire higher strengths for this application in the interest of component weight reduction, the weakest link is the requirement of high SCC resistance for those component parts. Today's spar parts are thus traditionally manufactured from a more corrosion resistant, but lower strength alloy temper such as T74. Based on the observed EC increase at the same strength, and the AI SCC test results described above, the preferred, new 3-stage aging methods of this invention can offer these structural/materials engineers and aerospace part designers a method of providing the strength levels of 7050/7010/7040-T76 products with near T74 corrosion resistance levels. Alternatively, this invention can offer the corrosion resistance of a T76 tempered material in combination with significantly higher strength levels.
EXAMPLES
(51) Three representative compositions of the new 7xxx alloy family were cast to target as large, commercial scale ingots with the following compositions:
(52) TABLE-US-00004 TABLE 4 Alloy wt. % Zn wt. % Cu wt. % Mg wt. % Fe wt. % Si wt. % Zr wt. % Ti A 7.3 1.6 1.5 0.04 0.02 0.11 0.02 B 6.7 1.9 1.5 0.05 0.02 0.11 0.02 C 7.4 1.9 1.5 0.04 0.02 0.11 0.02
Those cast ingot materials, of course after working, i.e. rolling to 6-inch finish gauge plate, solution heat treating, etc., were subjected to the comparative aging practice variations set forth in Table 5 below. Actually, two different first stages were compared in this 3-stage evaluation, one having a single exposure at 250 F. with the other broken into two sub-stages: 4 hours @ 225 F., followed by a second sub-stage of 6 hours @ 250 F. This two sub-stage procedure is referred to herein as first a first stage treatment, i.e., prior to the second stage treatment at about 310 F. In any event, no noticeable difference in properties was observed between these two types of first stages, the lone treatment at 250 F. versus the split treatments at both 225 and 250 F. Hence, referring to any stage herein embraces such variants.
(53) TABLE-US-00005 TABLE 5 First Step/Time Second Step/Time Third Step/Time Two-Step 250 F./6 hrs. 310 F./~5 Aging to 15 hrs. Three-Step 250 F./6 hrs. 310 F./~5 250 F./24 hrs. Aging to ~15 hrs. 225 F./4 hrs. + 310 F./~5 250 F./24 hrs. 250 F./6 hrs. to ~15 hrs.
Specimens from each six inch thick plate were then tested, with the averages for the two- and three-step aged properties being measured as follows:
(54) TABLE-US-00006 TABLE 6 Average TYS & EC Properties Tensile Yield 2-step Age EC, 3-step Age EC, Alloy (T/4) ksi % IACS % IACS A 74.4 38.5 40.0 B 74.6 38.5 39.8 C 75.3 38.5 39.7
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(56) AI SCC studies were performed per ASTM Standard D-1141, by alternate immersion, in a specified synthetic ocean water (or SOW) solution, which is more aggressive than the more typical 3.5% NaCl salt solution required by ASTM Standard G44. Table 7 shows the results on various Alloy A, B and C samples (all in an ST direction) with just two aging steps, the second step comprising various times (6, 8 and 11 hours) at about 320 F.3
(57) TABLE-US-00007 TABLE 7 Results of SCC Test by Alternate Immersion of Plant Processed 6 Plates of Alloys A, B and C Receiving 2-Stage Aging after 121 Days Exposure to Synthetic Ocean Water Stress Stress Stress EC TYS 6 Hours @ 250 F. (ksi) Days To (ksi) Days To (ksi) Days To (% IACS) (ksi) (1.sup.st stage) plus: (T/2) F/N(1) Failure (T/2) F/N(1) Failure (T/2) F/N(1) Failure (Surf) (T/4) Alloy A-T7X 6 Plate 6 Hr/320 F. 25 1/5 77 d 35 4/5 10, 12, 21, 70 d 40 5/5 6, 7, 7, 27, 91 d 41.2 74.9 4 OK @ 121 d 1 OK @ 121 d 8 Hr/320 F. 25 0/5 5 OK @ 121 d 35 2/5 100, 100 d 40 3/5 13, 13, 50 d 41.6 72.5 3 OK @ 121 d 2 OK @ 121 d 11 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 0/5 5 OK @ 121 d 42.9 67.2 Alloy B-I7X 6 Plate 6 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 0/5 5 OK @ 121 d 41.3 74.8 8 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 0/5 5 OK @ 121 d 41.7 73.1 11 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 0/5 5 OK @ 121 d 42.2 69.2 Alloy C-T7X 6 Plate 6 Hr/320 F. 25 1/5 13 d 35 0/5 5 OK @ 121 d 40 3/5 23, 26, 34 d 40.9 75.3 4 OK @ 121 d 2 OK @ 121 d 8 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 3/5 13, 19, 35 d 41.2 73.9 2 OK @ 121 d 11 Hr/320 F. 25 0/5 5 OK @ 121 d 35 0/5 5 OK @ 121 d 40 0/5 5 OK @ 121 d 42.2 69.2 Note: F/N(1) = Number of specimens failed over the number exposed
From this data, several SCC failures were observed following exposure for 121 days, primarily as a function of short transverse (ST) applied stress, aging time and/or alloy.
(58) Comparative Table 8 lists SCC results for just Alloys A and C (applied stress in the same ST direction) after having been aged for an additional 24 hours at 250 F., that is for a total aging practice that comprises: (1) 6 hours at 250 F.; (2) 6, 8 or 11 hours at 320 F.; and (3) 24 hours at 250 F.
(59) TABLE-US-00008 TABLE 8 Results of SCC Test by Alternate Immersion of Plant Processed 6 Plates of Alloys A and C Receiving 3-Stage Aging after 93 Days Exposure to Synthetic Ocean Water by Alternate Immersion ASTM D-1141-90 Stress Stress Stress EC TYS 6 Hours @ 250 F. (ksi) Days To (ksi) Days To (ksi) Days To (% IACS) (ksi) (1.sup.st stage) plus: (T/2) F/N(1) Failure (T/2) F/N(1) Failure (T/2) F/N(1) Failure (T/10) (T/4) Alloy A-T7X Plate 6 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 39.7 74.2 8 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 40.4 72.1 11 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 41.5 67.4 Alloy C-T7X Plate 6 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 39.5 75.3 8 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 40.0 72.8 11 Hr/320 F. + 24 h/250 F. 25 0/3 3 OK @ 93 d 35 0/3 3 OK @ 93 d 45 0/3 3 OK @ 93 d 41.0 68.8 Note: F/N(1) = Number of specimens failed over the number exposed.
(60) Quite remarkably, no sample failures were observed under identical test conditions after the first 93 days of exposure. Thus, the new 3-step aging approach of this invention is believed to confer unique strength/SCC advantages surpassing those achievable through conventional 2-step aging while promising to develop better property attributes in new products and confer further property combination improvements in still other, current aerospace product lines.
(61) The value of comparing Table 7 data to that in Table 8 is to underscore that while 2-stage/step aging may be practiced on the alloy according to this invention, the preferred 3-stage aging method herein described actually imparts a measurable SCC test performance improvement. Tables 6 and 7 also include SCC performance indicator data, EC values (as a % IACS), along with correspondingly measured TYS (T/4) values. That data must not be compared, side-by-side, for determining the relative value of a two versus 3-step aged products, however, as the EC testing was performed at different areas of the product, i.e. Table 7 using surface measured values versus the T/10 measurements of Table 8 (it being known that EC indicator values generally decrease when measuring from the surface going inward on a given test specimen). The TYS values cannot be used as a true comparison either as lot sizes varied as well as testing location (laboratory versus plant). Instead, the relative data of
(62) Seacoast SCC test data confirms the significant improvements in corrosion resistance realized by imparting a novel three-step aging method to the aforementioned new family of 7XXX alloys. For the alloy composition identified as Alloy A in above Table 4, SCC testing extended over a 568-day period for 2-stage aged versus a 328-day test period for the 3-stage aged, with the comparative 2- versus 3-stage aged SCC performances mapped per following Table 9 (The latter (3-stage) testing was started after the former (2-stage) tests had commenced; hence, the longer test times observed for 2-stage aged specimens).
(63) TABLE-US-00009 TABLE 9 Comparison of Short-Transverse Seacoast SCC Performance from 2- versus 3-Step aging Practices with 320 F. 2.sup.nd Step Aging for Alloy A Days Survived until Failure Aging Practice 2-Step Aging 3-Step Aging Aging Time at 320 F. 6 Hrs 8 Hrs 7 hrs 9 hrs L-TYS 74.9 ksi 72.5 ksi 73.3 ksi 71.0 ksi Short- 23 ksi + + + + + + Transverse 25 ksi 39, 39 507, 39 46, 39, 46, 39, 46 + + + + + + Applied 27 ksi + + + + + + Stress 29 ksi + + + + + + 31 ksi + + + + + + 33 ksi + + + + + + 35 ksi 39, 39, 39, 39, 39 39, 39, 39, 39, 39 + + + + + + 37 ksi 314 + + + + + 39 ksi + + + + + + 40 ksi 39, 39, 39, 39, 39 39, 39, 39, 59, 39 41 ksi + + + 265 + + 43 ksi 167 + 167 + + + 45 ksi 39, 39, 39, 39, 39 39, 39, 39, 39, 39 + 272, 328 + + + 47 ksi 167, 153 + + + + 49 ksi 187, 265, 90 293 + 237 51 ksi 251, 97, 160 + + +
Specimen surviving 568 Days + Specimens surviving 328 Days Note: 2-stage aging comprised: 6 hours @ 250 F.; and 6 or 8 hours @ 320 F. 3-stage aging comprised: 6 hours @ 250 F.; 7 or 9 hours @ 320 F.; and 24 hours @ 250 F.
This data is graphically summarized in accompanying
(64) A second composition, Alloy C in Table 4 (with its 7.4 wt. % Zn, 1.5 wt. % Mg, 1.9 wt. % Cu, and 0.11 wt. % Zr), was subjected to the comparative 2- versus 3-step agings as was Alloy A above. The long term results from those Seacoast SCC tests are summarized in Table 10 below.
(65) TABLE-US-00010 TABLE 10 Comparison of Short-Transverse Seacoast SCC Performance from 2-versus 3-Step aging Practices with 320 F. 2.sup.nd Step Aging for Alloy C Days Survived until Failure Aging Practice 2-Step Aging 3-Step Aging Aging Time at 320 F. 6 Hrs 8 Hrs 7 Hrs 9.5 Hrs L-TYS 75.3 ksi 73.9 ksi 74.3 ksi 72.8 ksi Short- 23 ksi + + + + + + Transverse 25 ksi
39
39
59
+ + + + + + Applied 27 ksi + + + + + + Stress 29 ksi + + + + + + 31 ksi + + + + + + 33 ksi + + + + + + 35 ksi 39, 39, 39, 39, 39 59, 39, 67, 73, 39 + + + + + + 37 ksi + + + + + + 39 ksi + + + + + + 40 ksi 39, 39, 67, 39, 39 39, 39, 39, 46, 67 41 ksi + + + + + + 43 ksi + + + + + + 45 ksi 39, 39, 39, 39, 39 39, 53, 39, 39, 39 + + 244 + + + 47 ksi + + + + + + 49 ksi + 272 + + + + 51 ksi 181 + + + 265 +
Specimen surviving 568 Days + Specimens surviving 328 Days
(66) Graphically, this Table 10 data is shown in accompanying
(67) With respect to the 3-stage aging, preferred particulars for the aforementioned alloy compositions, one must note that: the first stage age should preferably take place within about 200 to 275 F., more preferably between about 225 or 230 to 260 F., and most preferably at or about 250 F. And while about 6 hours at the aforesaid temperature or temperatures is quite satisfactory, it must be noted that in any broad sense, the amount of time spent for first step aging should be a time sufficient for producing a substantial amount of precipitation hardening. Thus, relatively short hold times, for instance of about 2 or 3 hours, at a temperature of about 250 F., may be sufficient (1) depending on part size and shape complexity; and (2) especially when the aforementioned shortened treatment/exposure is coupled with a relatively slow heat up rate of several hours, for instance 4 to 6 or 7 hours, total.
(68) The preferred second stage aging practice to be imparted on the preferred alloy compositions of this invention can be purposefully ramped up directly from the aforementioned first step heat treatment. Or, there may be a purposeful and distinct time/temperature interruption between first and second stages. Broadly stated, this second step should take place within about 290 or 300 to 330 or 335 F. Preferably, this second step age is performed within about 305 and 325 F. Preferably, second step aging takes place between about 310 to 320 or 325 F. The preferred exposure times for this critical second step processing depend somewhat inversely on the actual temperature(s) employed. For instance, if one were to operate substantially at or very near 310 F., a total exposure time from about 6 to 18 hours, preferably for about 7 to 13, or even 15 hours would suffice. More preferably, second step agings would proceed for about 10 or 11, even 13, total hours at that operating temperature. At a second aging stage temperature of about 320 F., total second step times can range between about 6 to 10 hours with about 7 or 8 to 10 or 11 hours being preferred. There is also a preferred target property aspect to second step aging time and temperature selection. Most notably, shorter treatment times at a given temperature favor higher strength values whereas longer exposure times favor better corrosion resistance performance.
(69) Finally, with respect to the preferred, third aging practice stage, it is better to not ramp slowly down from the second step for performing this necessary third step on such thick workpieces unless extreme care is exercised to coordinate closely with the second step temperature and total time duration so as to avoid exposures at second aging stage temperatures for too long a time. Between the second and third aging steps, the metal products of this invention can be purposefully removed from the heating furnace and rapidly cooled, using fans or the like, to either about 250 F. or less, perhaps even fully back down to room temperature. In any event, the preferred time/temperature exposures for the third aging step of this invention closely parallel those set forth for the first aging step above.
(70) In accordance with the invention, the invention alloy is preferably made into a product, suitably an ingot derived product, suitable for hot rolling. For instance, large ingots can be semi-continuously cast of the aforesaid composition and then can be scalped or machined to remove surface imperfections as needed or required to provide a good rolling surface. The ingot may then be preheated to homogenize and solutionize its interior structure and a suitable preheat treatment is to heat to a relatively high temperature for this type of composition, such as 900 F. In doing so, it is preferred to heat to a first lesser temperature level such as heating above 800 F., for instance about 820 F. or above, or 850 F. or above, preferably 860 F. or more, for instance around 870 F. or more, and hold the ingot at about that temperature or temperatures for a significant time, for instance, 3 or 4 hours. Next the ingot is heated the rest of the way up to a temperature of around 890 F. or 900 F. or possibly more for another hold time of a few hours. Such stepped or staged heat ups for homogenizing have been known in the art for many years. It is preferred that homogenizing be conducted at cumulative hold times in the neighborhood of 4 to 20 hours or more, the homogenizing temperatures referring to temperatures above about 880 to 890 F. That is, the cumulative hold time at temperatures above about 890 F. should be at least 4 hours and preferably more, for instance 8 to 20 or 24 hours, or more. As is known, larger ingot size and other matters can suggest longer homogenizing times. It is preferred that the combined total volume percent of insoluble and soluble constituents be kept low, for instance not over 1.5 vol. %, preferably not over 1 vol. %. Use of the herein described relatively high preheat or homogenization and solution heat treat temperatures aid in this respect, although high temperatures warrant caution to avoid partial melting. Such cautions can include careful heat-ups including slow or step-type heating, or both.
(71) The ingot is then hot rolled and it is desirable to achieve an unrecrystallized grain structure in the rolled plate product. Hence, the ingot for hot rolling can exit the furnace at a temperature substantially above about 820 F., for instance around 840 to 850 F. or possibly more, and the rolling operation is carried out at initial temperatures above 775 F., or better yet, above 800 F., for instance around 810 or even 825 F. This increases the likelihood of reducing recrystallization and it is also preferred in some situations to conduct the rolling without a reheating operation by using the power of the rolling mill and heat conservation during rolling to maintain the rolling temperature above a desired minimum, such as 750 F. or so. Typically, in practicing the invention, it is preferred to have a maximum recrystallization of about 50% or less, preferably about 35% or less, and most preferably no more than about 25% recrystallization, it being understood that the less recrystallization achieved, the better the fracture toughness properties.
(72) Hot rolling is continued, normally in a reversing hot rolling mill, until the desired thickness of the plate is achieved. In accordance with the invention, plate product intending to be machined into aircraft components such as integral spars can range from about 2 to 3 inches to about 9 or 10 inches thick or more. Typically, this plate ranges from around 4 inches thick for relatively smaller aircraft, to thicker plate of about 6 or 8 inches to about 10 or 12 inches or more. In addition to the preferred embodiments, it is believed this invention can be used to make the lower wing skins of small, commercial jet airliners. Still other applications can include forgings and extrusions, especially thick sectioned versions of same. In making extrusion, the invention alloy is extruded within around 600 to 750 F., for instance, at around 700 F., and preferably includes a reduction in cross-sectional area (extrusion ratio) of about 10:1 or more. Forging can also be used herein.
(73) The hot rolled plate or other wrought product is solution heat treated (SHT) by heating within around 840 or 850 F. to 880 or 900 F. to take into solution substantial portions, preferably all or substantially all, of the zinc, magnesium and copper soluble at the SHT temperature, it being understood that with physical processes which are not always perfect, probably every last vestige of these main alloying ingredients may not be fully dissolved during the SHT (solutionizing). After heating to the elevated temperature as just described, the product should be quenched to complete the solution heat treating procedure. Such cooling is typically accomplished either by immersion in a suitably sized tank of cold water or by water sprays, although air chilling might be usable as supplementary or substitute cooling means for some cooling. After quenching, certain products may need to be cold worked, such as by stretching or compression, so as to relieve internal stresses or straighten the product, even possibly in some cases, to further strengthen the plate product. For instance, the plate may be stretched or compressed 1 or 1 or possibly 2% or 3% or more, or otherwise cold worked a generally equivalent amount. A solution heat treated (and quenched) product, with or without cold working, is then considered to be in a precipitation hardenable condition, or ready for artificial aging according to preferred artificial aging methods as herein described or other artificial aging techniques. As used herein, the term solution heat treat, unless indicated otherwise, shall be meant to include quenching.
(74) After quenching, and cold working if desired, the product (which may be a plate product) is artificially aged by heating to an appropriate temperature to improve strength and other properties. In one preferred thermal aging treatment, the precipitation hardenable plate alloy product is subjected to three main aging steps, phases or treatments as described above, although clear lines of demarcation may not exist between each step or phase. It is generally known that ramping up to and/or down from a given or target treatment temperature, in itself, can produce precipitation (aging) effects which can, and often need to be, taken into account by integrating such ramping conditions and their precipitation hardening effects into the total aging treatment.
(75) It is also possible to use aging integration in conjunction with the aging practices of this invention. For instance, in a programmable air furnace, following completion of a first stage heat treatment of 250 F. for 24 hours, the temperature in that same furnace can be gradually progressively raised to temperature levels around 310 or so over a suitable length of time, even with no true hold time, after which the metal can then be immediately transferred to another furnace already stabilized at 250 F. and held for 6 to 24 hours. This more continuous, aging regime does not involve transitioning to room temperature between first-to-second and second-to-third stage aging treatments. Such aging integration was described in more detail in U.S. Pat. No. 3,645,804, the entire content of which is fully incorporated by reference herein. With ramping and its corresponding integration, two, or on a less preferred basis, possibly three, phases for artificially aging the plate product may be possible in a single, programmable furnace. For purposes of convenience and ease of understanding, however, preferred embodiments of this invention have been described in more detail as if each stage, step or phase was distinct from the other two artificial aging practices imposed hereon. Generally speaking, the first of these three steps or stages is believed to precipitation harden the alloy product in question; the second (higher temperature) stage then exposes the invention alloy to one or more elevated temperatures for increasing its resistance to corrosion, especially stress corrosion cracking (SCC) resistance under both normal, industrial and seacoast-simulated atmospheric conditions. The third and final stage then further precipitation hardens the invention alloy to a high strength level while also imparting further improved corrosion properties thereto.
(76) The low quench sensitivity of the invention alloy can offer yet another potential application in a class of processes generally described as press quenching by those skilled in the art. One can illustrate the press quenching process by considering the standard manufacturing flow path of an age hardenable extrusion alloy such as one that belongs to the 2XXX, 6XXX, 7XXX or 8XXX alloy series. The typical flow path involves: Direct Chill (DC) ingot casting of billets, homogenization, cooling to ambient temperature, reheating to the extrusion temperature by furnaces or induction heaters, extrusion of the heated billet to final shape, cooling the extruded part to ambient temperature, solution heat treating the part, quenching, stretching and either naturally aged at room temperature or artificially aged at elevated temperature to the final temper. The press quenching process involves controlling the extrusion temperature and other extrusion conditions such that upon exiting the extrusion die, the part is at or near the desired solution heating temperature and the soluble constituents are effectively brought to solid solution. It is then immediately and directly continuously quenched as the part exits the extrusion press by either water, pressurized air or other media. The press quenched part can then go through the usual stretching, followed by either natural or artificial aging. Hence, as compared to the typical flow path, the costly separate solution heat treating process is eliminated from this press quenched variation, thereby significantly lowering overall manufacturing costs, and energy consumption as well.
(77) For most alloys, especially those belonging to the relatively quench sensitive 7XXX alloy series, the quench provided by the press quenching process is generally not as effective as compared to that provided by the separate solution heat treatment, such that significant degradation of certain material attributes such as strength, fracture toughness, corrosion resistance and other properties can result from press quenching. Since the invention alloy has very low quench sensitivity, it is expected that the property degradation during press quenching is either eliminated or significantly reduced to acceptable levels for many applications.
(78) For the mold plate embodiments of this invention where SCC resistance is not as critical, known single or two-stage artificial aging treatments may also be practiced on these compositions instead of the preferred three-step aging method described herein.
(79) When referring to a minimum (for instance, strength or toughness property value), such can refer to a level at which specifications for purchasing or designating materials can be written or a level at which a material can be guaranteed or a level that an airframe builder (subject to safety factor) can rely on in design. In some cases, it can have a statistical basis wherein 99% of the product conforms or is expected to conform with 95% confidence using standard statistical methods. Because of an insufficient amount of data, it is not statistically accurate to refer to certain minimum or maximum values of the invention as true guaranteed values. In those instances, calculations have been made from currently available data for extrapolating values (e.g. maximums and minimums) therefrom. See, for example, the Currently Extrapolated Minimum S/N values plotted for plate (solid line A-A in
(80) Fracture toughness is an important property to airframe designers, particularly if good toughness can be combined with good strength. By way of comparison, the tensile strength, or ability to sustain load without fracturing, of a structural component under a tensile load can be defined as the load divided by the area of the smallest section of the component perpendicular to the tensile load (net section stress). For a simple, straight-sided structure, the strength of the section is readily related to the breaking or tensile strength of a smooth tensile coupon. This is how tension testing is done. However, for a structure containing a crack or crack-like defect, the strength of a structural component depends on the length of the crack, the geometry of the structural component, and a property of the material known as the fracture toughness. Fracture toughness can be thought of as the resistance of a material to the harmful or even catastrophic propagation of a crack under a load.
(81) Fracture toughness can be measured in several ways. One way is to load in tension a test coupon containing a crack. The load required to fracture the test coupon divided by its net section area (the cross-sectional area less the area containing the crack) is known as the residual strength with units of thousands of pounds force per unit area (ksi). When the strength of the material as well as the specimen geometry are constant, the residual strength is a measure of the fracture toughness of the material. Because it is so dependent on strength and specimen geometry, residual strength is usually used as a measure of fracture toughness when other methods are not as practical as desired because of some constraint like size or shape of the available material.
(82) When the geometry of a structural component is such that it does not deform plastically through the thickness when a tension load is applied (plane-strain deformation), fracture toughness is often measured as plane-strain fracture toughness, K.sub.IC. This normally applies to relatively thick products or sections, for instance 0.6 or preferably 0.8 or 1 inch or more. The ASTM has established a standard test using a fatigue pre-cracked compact tension specimen to measure K.sub.IC which has the units ksiin. This test is usually used to measure fracture toughness when the material is thick because it is believed to be independent of specimen geometry as long as appropriate standards for width, crack length and thickness are met. The symbol K, as used in K.sub.IC, is referred to as the stress intensity factor.
(83) Structural components which deform by plane-strain are relatively thick as indicated above. Thinner structural components (less than 0.8 to 1 inch thick) usually deform under plane stress or more usually under a mixed mode condition. Measuring fracture toughness under this condition can introduce variables because the number which results from the test depends to some extent on the geometry of the test coupon. One test method is to apply a continuously increasing load to a rectangular test coupon containing a crack. A plot of stress intensity versus crack extension known as an R-curve (crack resistance curve) can be obtained this way. The load at a particular amount of crack extension based on a 25% secant offset in the load vs. crack extension curve and the effective crack length at that load are used to calculate a measure of fracture toughness known as K.sub.R25. At a 20% secant, it is known as K.sub.R20. It also has the units of ksiin. Well known ASTM E561 concerns R-curve determination, and such is generally recognized in the art.
(84) When the geometry of the alloy product or structural component is such that it permits deformation plastically through its thickness when a tension load is applied, fracture toughness is often measured as plane-stress fracture toughness which can be determined from a center cracked tension test. The fracture toughness measure uses the maximum load generated on a relatively thin, wide pre-cracked specimen. When the crack length at the maximum load is used to calculate the stress-intensity factor at that load, the stress-intensity factor is referred to as plane-stress fracture toughness K.sub.c. When the stress-intensity factor is calculated using the crack length before the load is applied, however, the result of the calculation is known as the apparent fracture toughness, Kapp, of the material. Because the crack length in the calculation of K.sub.c is usually longer, values for K.sub.c are usually higher than K.sub.app for a given material. Both of these measures of fracture toughness are expressed in the units ksiin. For tough materials, the numerical values generated by such tests generally increase as the width of the specimen increases or its thickness decreases as is recognized in the art. Unless indicated otherwise herein, plane stress (K.sub.c) values referred to herein refer to 16-inch wide test panels. Those skilled in the art recognize that test results can vary depending on the test panel width, and it is intended to encompass all such tests in referring to toughness. Hence, toughness substantially equivalent to or substantially corresponding to a minimum value for K, or K.sub.app in characterizing the invention products, while largely referring to a test with a 16-inch panel, is intended to embrace variations in K.sub.c or K.sub.app encountered in using different width panels as those skilled in the art will appreciate.
(85) The temperature at which the toughness is measured can be significant. In high altitude flights, the temperature encountered is quite low, for instance, minus 65 F., and for newer commercial jet aircraft projects, toughness at minus 65 F. is a significant factor, it being desired that the lower wing material exhibit a toughness K.sub.IC level of around 45 ksiin at minus 65 F. or, in terms of K.sub.R20, a level of 95 ksiin, preferably 100 ksiin or more. Because of such higher toughness values, lower wings made from these alloys may replace today's 2000 (or 2XXX Series) alloy counterparts with their corresponding property (i.e. strength/toughness) trade-offs. Through the practice of this invention, it may also be possible to make upper wing skins from same, alone or in combination with integrally formed components, like stiffeners, ribs and stringers.
(86) The toughness of the improved products according to the invention is very high and in some cases may allow the aircraft designer's focus for a material's durability and damage tolerance to emphasize fatigue resistance as well as fracture toughness measurement. Resistance to cracking by fatigue is a very desirable property. The fatigue cracking referred to occurs as a result of repeated loading and unloading cycles, or cycling between a high and a low load such as when a wing moves up and down. This cycling in load can occur during flight due to gusts or other sudden changes in air pressure, or on the ground while the aircraft is taxing. Fatigue failures account for a large percentage of failures in aircraft components. These failures are insidious because they can occur under normal operating conditions without excessive overloads, and without warning. Crack evolution is accelerated because material inhomogeneities act as sites for initiation or facilitate linking of smaller cracks. Therefore, process or compositional changes which improve metal quality by reducing the severity or number of harmful inhomogeneities improve fatigue durability.
(87) Stress-life cycle (S-N or S/N) fatigue tests characterize a material resistance to fatigue initiation and small crack growth which comprises a major portion of total fatigue life. Hence, improvements in S-N fatigue properties may enable a component to operate at higher stresses over its design life or operate at the same stress with increased lifetime. The former can translate into significant weight savings by downsizing, or manufacturing cost saving by component or structural simplification, while the latter can translate into fewer inspections and lower support costs. The loads during fatigue testing are below the static ultimate or tensile strength of the material measured in a tensile test and they are typically below the yield strength of the material. The fatigue initiation fatigue test is an important indicator for a buried or hidden structural member such as a wing spar which is not readily accessible for visual or other examination to look for cracks or crack starts.
(88) If a crack or crack-like defect exists in a structure, repeated cyclic or fatigue loading can cause the crack to grow. This is referred to as fatigue crack propagation. Propagation of a crack by fatigue may lead to a crack large enough to propagate catastrophically when the combination of crack size and loads are sufficient to exceed the material's fracture toughness. Thus, performance in the resistance of a material to crack propagation by fatigue offers substantial benefits to aerostructure longevity. The slower a crack propagates, the better. A rapidly propagating crack in an airplane structural member can lead to catastrophic failure without adequate time for detection, whereas a slowly propagating crack allows time for detection and corrective action or repair. Hence, a low fatigue crack growth rate is a desirable property.
(89) The rate at which a crack in a material propagates during cyclic loading is influenced by the length of the crack. Another important factor is the difference between the maximum and the minimum loads between which the structure is cycled. One measurement including the effects of crack length and the difference between maximum and minimum loads is called the cyclic stress intensity factor range or K, having units of ksiin, similar to the stress intensity factor used to measure fracture toughness. The stress intensity factor range (K) is the difference between the stress intensity factors at the maximum and minimum loads. Another measure affecting fatigue crack propagation is the ratio between the minimum and the maximum loads during cycling, and this is called the stress ratio and is denoted by R, a ratio of 0.1 meaning that the maximum load is 10 times the minimum load. The stress, or load, ratio may be positive or negative or zero. Fatigue crack growth rate testing is typically done in accordance with ASTM E647-88 (and others) well known in the art. As used herein, Kt refers to a theoretical stress concentration factor as described in ASTM E1823.
(90) The fatigue crack propagation rate can be measured for a material using a test coupon containing a crack. One such test specimen or coupon is about 12 inches long by 4 inches wide having a notch in its center extending in a cross-wise direction (across the width; normal to the length). The notch is about 0.032 inch wide and about 0.2 inch long including a 600 bevel at each end of the slot. The test coupon is subjected to cyclic loading and the crack grows at the end(s) of the notch. After the crack reaches a predetermined length, the length of the crack is measured periodically. The crack growth rate can be calculated for a given increment of crack extension by dividing the change in crack length (called a) by the number of loading cycles (N) which resulted in that amount of crack growth. The crack propagation rate is represented by a/N or da/dN and has units of inches/cycle. The fatigue crack propagation rates of a material can be determined from a center cracked tension panel. In a comparison using R=0.1 tested at a relative humidity over 90% with K ranging from around 4 to 20 or 30, the invention material exhibited relatively good resistance to fatigue crack growth. However, the superior performance in S-N fatigue makes the invention material much better suited for a buried or hidden member such as a wing spar.
(91) The invention products exhibit very good corrosion resistance in addition to the very good strength and toughness and damage tolerance performance. The exfoliation corrosion resistance for products in accordance with the invention can be EB or better (meaning EA or pitting only) in the EXCO test for test specimens taken at either mid-thickness (T/2) or one-tenth of the thickness from the surface (T/10) (T being thickness) or both. EXCO testing is known in the art and is described in well known ASTM Standard No. G34. An EXCO rating of EB is considered good corrosion resistance in that it is considered acceptable for some commercial aircraft; EA is still better.
(92) Stress corrosion cracking resistance across the short transverse direction is often considered an important property especially in relatively thick members. The stress corrosion cracking resistance for products in accordance with the invention in the short transverse direction can be equivalent to that needed to pass a -inch round bar alternate immersion test for 20, or alternately 30, days at 25 or 30 ksi or more, using test procedures in accordance with ASTM G47 (including ASTM G44 and G38 for C-ring specimens and G49 for -inch bars), said ASTM G47, G44, G49 and G38, all well known in the art.
(93) As a general indicator of exfoliation corrosion and stress corrosion resistance, the plate typically can have an electrical conductivity of at least about 36, or preferably 38 to 40% or more of the International Annealed Copper Standard (% IACS). Thus, the good exfoliation corrosion resistance of the invention is evidenced by an EXCO rating of EB or better, but in some cases other measures of corrosion resistance may be specified or required by airframe builders, such as stress corrosion cracking resistance or electrical conductivity. Satisfying any one or more of these specifications is considered good corrosion resistance.
(94) The invention has been described with some emphasis on wrought plate which is preferred, but it is believed that other product forms may be able to enjoy the benefits of the invention, including extrusions and forgings. To this point, the emphasis has been on stiffener-type, fuselage or wing skin stringers which can be J-shaped, Z- or S-shaped, or even in the shape of a hat-shaped channel. The purpose of such stiffeners is to reinforce the plane's wing skin or fuselage, or any other shape that can be attached to same, while not adding a lot of weight thereto. While in some cases it is preferred for manufacturing economies to separately fasten stringers, such can be machined from a much thicker plate by the removal of the metal between the stiffener geometries, leaving only the stiffener shapes integral with the main wing skin thickness, thus eliminating all the rivets. Also the invention has been described in terms of thick plate for machining wing spar members as explained above, the spar member generally corresponding in length to the wing skin material. In addition, significant improvements in the properties of this invention render its use as thickly cast mold plate highly practical.
(95) Because of its reduced quench sensitivity, it is believed that when an alloy product according to the invention is welded to a second product, it will exhibit in its heat affected, welding zone an improved retention of its strength, fatigue, fracture toughness and/or corrosion resistance properties. This applies regardless of whether such alloy products are welded by solid state welding techniques, including friction stir welding, or by known or subsequently developed fusion techniques including, but not limited to, electron beam welding and laser welding. Through the practice of this invention, both welded parts may be made from the same alloy composition.
(96) For some parts/products made according to the invention, it is likely that such parts/products may be age formed. Age forming promises a lower manufacturing cost while allowing more complex wing shapes to be formed, typically on thinner gauge components. During age forming, the part is mechanically constrained in a die at an elevated temperature usually about 250 F. or higher for several to tens of hours, and desired contours are accomplished through stress relaxation. Especially during a higher temperature artificial aging treatment, such as a treatment above about 320 F., the metal can be formed or deformed into a desired shape. In general, the deformations envisioned are relatively simple such as including a very mild curvature across the width of a plate member together with a mild curvature along the length of said plate member. It can be desirable to achieve the formation of these mild curvature conditions during the artificial aging treatment, especially during the higher temperature, second stage artificial aging temperature. In general, the plate material is heated above around 300 F., for instance around 320 or 330 F., and typically can be placed upon a convex form and loaded by clamping or load application at opposite edges of the plate. The plate more or less assumes the contour of the form over a relatively brief period of time but upon cooling springs back a little when the force or load is removed. The expected springback is compensated for in designing the curvature or contour of the form which is slightly exaggerated with respect to the desired forming of the plate to compensate for springback. Most preferably, the third artificial aging stage at a low temperature such as around 250 F. follows age forming. Either before or after its age forming treatment, the plate member can be machined, for instance, such as by tapering the plate such that the portion intended to be closer to the fuselage is thicker and the portion closest to the wing tip is thinner. Additional machining or other shaping operations, if desired, can also be performed either before or after age forming. High capacity aircrafts may require a relatively thicker plate and a higher level of forming than previously used on a large scale for thinner plate sections.
(97) Various invention alloy product forms, i.e. both thick plate (
(98) TABLE-US-00011 TABLE 11 Invention Alloy Compositions Zn Mg Cu Zr Fe Si Product (wt. %) (wt. %) (wt %) (wt. %) (wt. %) (wt. %) Plate D, F & G 7.25 1.45 1.54 0.11 0.03 0.007 and Forging D Plate E and 7.63 1.42 1.62 0.11 0.04 0.007 Forging E
(99) For these open hole fatigue life evaluations, in the L-T orientation, specific test parameters for both plate and forged product forms included: a K.sub.t value of 2.3, Frequency of 30 Hz, R value=0.1 and Relative Humidity (RH) greater than 90%. The plate test results were then graphed in accompanying
(100) Referring now to
(101) TABLE-US-00012 TABLE 12 Currently Extrapolated Minimum S/N Plate Values (L-T) Applied Maximum Stress (ksi) Minimum Cycles to Failure 47.0 6,000 42.3 10,000 32.4 30,000 25.1 100,000 21.8 300,000 19.5 1,000,000
(102) Solid line (A-A) was then drawn on
(103) From the open hole fatigue life (S/N) data for various sized (i.e. 4-inch, 6-inch, and 8-inch) forgings, a dotted line was drawn for mathematically representing the mean values of 6-inch thick comp E and 8-inch thick comp D forgings. Note, several samples tested did not fracture during these tests; they are grouped together in a circle to the right of
(104) TABLE-US-00013 TABLE 13 Currently Extrapolated Minimum S/N Forging Values (L-T) Applied Maximum Stress (ksi) Minimum Cycles to Failure 42.0 8,000 39.4 10,000 30.8 30,000 25.1 100,000 21.8 300,000 19.2 1,000,000
(105) Solid line (B-B) was then drawn on
(106) In
(107) TABLE-US-00014 TABLE 14 Currently Extrapolated Maximum L-T, FCG Values K (ksiin) Max. da/dN (in./cycle) 10 0.000025 15 0.000047 20 0.00009 25 0.0002 30 0.0005 34 0.0014
(108) A currently extrapolated maximum FCG value, solid curve line (C-C) for thick plate and forgings per the invention was drawn, against which one jetliner manufacturer's specified FCG values for 7040/7050-T7451 (3 to 8.7 inch thick) plate was overlaid, said values being taken in both the L-T and T-L orientation.
(109) Plate product forms of the invention have also been subjected to hole crack initiation tests, involving the drilling of a preset hole (less than 1 inch diameter) into a test specimen, inserting into that drilled hole a split sleeve, then pulling a variably oversized mandrel through said sleeve and pre-drilled hole. Under such testing, the 6- and 8-inch thick plate product of this invention did not have any cracks initiate from the drilled holes thereby showing very good performance.
(110) Having described the presently preferred embodiments, it is to be understood that the invention may be otherwise embodied within the scope of the appended claims.