FR2789405A1 - New quenched and stretched aluminum-copper-magnesium alloy product, for aircraft wing intrados skin and wing or fuselage intrados strut manufacture has a large plastic deformation range - Google Patents

Abstract

A quenched and stretched wrought aluminum-copper-magnesium alloy product, having a specified composition and having a large plastic deformation range, is new. A rolled, extruded or forged AlCuMg alloy product, which has been solution treated, quenched and cold stretched and which is used for manufacture of aircraft structural elements, contains (by wt.) less than 0.15% Fe, less than 0.15% Si, 3.8-4.4(preferably 4.0-4.3)% Cu, 1-1.5% Mg, 0.5-0.8% Mn, 0.08-0.15% Zr and less than 0.05% each (less than 0.15% total) of other elements and has a Rm(L)/R0.2(L) ratio of greater than 1.25, preferably greater than 1.30. An Independent claim is also included for a process for producing the above product

Description

AICuMg alloy product for aircraft structural element.

  TECHNICAL FIELD The invention relates to rolled, forged or forged products made from cold-formed, cold-formed AiCuMg alloys for the manufacture of aircraft structural elements, in particular skin panels and wing-bottom stiffeners, and presenting, with respect to the products of the prior art used for the same application, improved properties of mechanical strength, toughness and damage tolerance. The designation of the alloys and the metallurgical states corresponds to the nomenclature of the Aluminum Association, taken up by the standards

EN 515 and EN 573.

  STATE OF THE ART High-capacity commercial aircraft wings comprise an upper part (or extrados) consisting of a skin made from thick plates made of alloy 7150 in state T651, or alloy 7055 in state T7751 or 7449 in the T7951 state, and stiffeners made from profiles of the same alloy, and a lower part (or intrados) consisting of a skin made from thick 2024 alloy plates in the T351 or 2324 state. T39 state, and stiffeners made from profiles of the

  same alloy. Both parts are assembled by longitudinal members and ribs.

  The alloy 2024 according to the designation of the Aluminum Association or the standard EN 573-3 has the following chemical composition (% by weight): Si <0.5 Fe <0.5 Cu: 3.8 - 4.9 Mg : 1.2 - 1.8 Mn: 0.3 - 0.9 Cr <0.10 Zn <0.25 Ti <0.15 Various variants have been developed and deposited at the Aluminum Association under the designations 2224, 2324 and 2424, in particular with more limited levels of silicon and iron. The alloy 2324 in the T39 state has been the subject of patent EP 0038605 (US 4294625) to Boeing. in which the improvement of the yield strength is obtained by cold-working using a cold rolling pass after quenching. This work-hardening tends to reduce the tenacity and, to compensate for the decrease in toughness, the contents are reduced. Fe, Si, Cu and Mg. Boeing also developed alloy 2034 with composition: Si <0.10 Fe <0.12 Cu: 4.2 - 4.8 Mg: 1.3 - 1.9 Mn: 0.8-1.3 Cr < 0.05 Zn <0.20 Ti <0.15 Zr: 0.08-0.15 This alloy was the subject of patent EP 0031605 (= US 4336075). Compared to 2024 in the T351 state, it has a better specific yield stress due to the increase in manganese content and the addition of another antirecrystallizer

  (Zr), as well as improved toughness and fatigue resistance.

  EP 0473122 (= US Pat. No. 5,213,639) to Alcoa discloses an alloy, registered in the Aluminum Association as 2524, of composition: Si <0.1 Fe <0.12 Cu: 3.8 - 4.5 Mg: 1 , 2 - 1.8 Mn: 0.3 - 0.9 possibly containing another antirecrystallizer (Zr, V, Hf, Cr, Ag or Sc). This alloy is intended more particularly for thin sheets for fuselage and has a tenacity

  and improved crack propagation resistance over 2024.

  The patent application EP 0731185 of the Applicant relates to an alloy, subsequently recorded under No. 2024A, of composition: Si <0.25 Fe <0.25 Cu: 3.5 - 5 Mg: 1 - 2 Mn <0.55 with the relation: 0 <Mn - 2Fe <0.2 The thick plates in this alloy have both improved toughness and

  reduced level of residual stresses, without loss on the other properties.

  Issue For the construction of new high capacity commercial aircraft, it is imperative to limit the weight, so that manufacturers' specifications impose higher typical requirements for wing panels, resulting in higher minimum values. for mechanical characteristics

  static and damage tolerance of aluminum alloy products used.

  In addition, larger parts must be machined without distortion in thicker sheets, which implies better control of the level of residual stresses. The object of the invention is therefore to provide AlCuMg alloy products in the quenched and cold-deformed state, intended for the manufacture of airline wing bottom panels, and having, compared to similar 2024 alloy products. , improved static mechanical characteristics, reduced crack growth rate, higher toughness and reduced residual stress, without degradation of

other properties of use.

  OBJECT OF THE INVENTION The subject of the invention is a laminated, spun or forged product made of AlCuMg alloy, treated by solution treatment, quenching and cold deformation, intended for the manufacture of aircraft structural elements, of composition ( % by weight): Fe + Si <0.20 (preferably <0.15) Cu: 3.6 - 4.4 Mg: 1 - 1.5 Min: 0.5 - 0.8 Zr: 0.08 - 0.15 other items: <0.05 each and <0.15 at

total.

  It also relates to a laminated product (a sheet) of the same composition with a thickness of between 6 and 60 mm and having in the quenched and cold-deformed state (T351 state) at least one of the following groups of properties; a) Breaking strength RmcL)> 475 MPa and yield strength Ro, 2 (L)> 370 MPa b) Critical intensity factor (L-T direction) K,> 170 MPa / m and Koc> 120 MPa'Jm

  (measured according to ASTM E561 on notched specimens taken at quarter

  thickness with parameters B = 5 mm, W = 500 and 2ao = 165 mm) c) Crack propagation velocity (LT) da / dn, measured according to ASTM E 647 on notched specimens taken at quarter thickness with W = 200 mm and B = 5mm <10-4 mm / cycle for AK = 12 MPa / m <2.5 10-4 mm / cycle for AK = 15 MPax / m and <5 10-4 mm / cycle for AK = This sheet also has a level of residual stresses such that the arrow f measured in the L and TL directions after machining at half thickness of a bar resting on two supports distant by a length I is such that: f <( 0.14 12) / ef being measured in microns, the thickness e of the sheet and the length I

being expressed in mm.

  The subject of the invention is also a process for the manufacture of a rolled, forged or forged product comprising the following steps: casting of a plate or a billet of the indicated composition, homogenization of this plate or billet between 450 and 500.degree. C., heat conversion to the desired product, optionally cold processing, solution setting at a temperature between 480 and 505.degree. C., quenching with cold water, cold deformation with at least 1 , 5% permanent deformation,

  - natural aging at ambient temperature.

Description of the invention

  The chemical composition of the product differs from that of the usual 2024 in reduced iron and silicon content, higher manganese content and zirconium addition. Compared to 2034, there is a lower manganese content and a slightly reduced copper content. Surprisingly, this narrow domain of composition (particularly with regard to manganese), associated with modifications of the manufacturing range, leads, compared with 2024 and 2034, to a significant improvement in the compromise between the resistance mechanical and damage tolerance. In addition, and quite unexpectedly, we observe, for thick products, a low rate of residual stresses, allowing a

  machining without distortion of large parts.

  The manufacturing process comprises the casting of plates, in the case where the product to be manufactured is a rolled sheet, or billets in the case in the case of a spun section or a forged part. The plate or the billet is scalped, then homogenized between 450 and 500 C. The hot transformation is then carried out by rolling, spinning or forging. This transformation is preferably carried out at a higher temperature than the temperatures usually used, the outlet temperature being greater than 420 ° C. and preferably 440 ° C., so as to obtain a little recrystallized structure on the treated product, with a recrystallization rate. at a quarter thickness less than 20%, and preferably 10%. The laminated, spun or forged half-product is then dissolved between 480 and 505 ° C., so that this dissolution is as complete as possible, that is to say that the maximum of potentially soluble phases, in particular the precipitates AI2Cu and Al2CuMg, effectively in solid solution. The quality of the dissolution can be assessed by differential enthalpy analysis (AED) by measuring the specific energy using the area of the peak on the thermogram. This specific energy must be

preferably less than 2 J / g.

  Then quenching with cold water, and cold deformation (eg controlled pulling) leading to a permanent elongation of at least 1.5%. The

  product finally undergoes natural aging at room temperature.

  The products according to the invention have significantly improved static mechanical characteristics with respect to the 2024-T351 alloy, currently used for aircraft wing-backs, and only slightly lower than that of 2034-T351. The tenacity, measured by the critical stress factors K and K is greater than 10% greater than that of 2024 and 2034, and the crack propagation rate da / dn is significantly improved compared to these two alloys, in particular for the high values of AK The durations of life in fatigue, measured on notched specimens taken at mid-thickness in the direction L, are also improved by more than 20% with respect to the 2024 and the 2034. Finally , the level of residual stresses, measured by the arrow f after machining at mid-thickness of a bar resting on two supports distant by a length 1, is rather low, whereas one could have expected on the contrary with a fiber structure. This arrow, measured in microns, is always less than the quotient (0.14 12) / e, the length I and

  the thickness e of the sheet being expressed in mm.

  All these properties mean that the products according to the invention are particularly well suited to the manufacture of aircraft structural elements, including wing bottoms, but also sections for wing box, for sill insoles and assembled ribs and skins and fuselage stiffeners.

Examples

  Three plates 1450 mm wide and 446 mm thick were cast in alloy 2024, 2034 and alloy according to the invention respectively. The chemical compositions (% by weight) of the alloys are given in Table 1:

Table I

alloy Si Fe Cu Mg "Mn Zr

  2024 0.12 0.20 4.06 1.36 0.54 0, 002

  2034 0.05 0.07 4.30 1.34 0.98 0, 104

  invention 0.06 0.08 4.14 1.26 0.65 0, 102 The plates were scalped and then homogenized under the following conditions For 2024, 2h at 495 C then 5h at 460 C For 2034, 5 h at 497 ° C. For the alloy according to the invention, mounted in 12 hours and maintained for 6 hours at 48 ° C. A portion of the sheets was then hot-rolled to a thickness of 40 mm in successive passes of the order of 20 ° C. mm. Another part of the sheets was hot rolled up to 15 mm. For the alloy according to the invention, the hot rolled inlet temperature was 467 ° C., the outlet temperature at 40 mm was 465 ° C.

mm of 444 C.

  The sheets were put in solution under the following conditions 3h and 6h at 497 C for the sheets in 2024 of thickness 15 and 40 mm respectively, 2h and 5h at 499 C for the sheets in 2034 of thickness 15 and 40 mm

  9h to 497 C for the sheets according to the invention.

  After quenching with cold water, all the sheets were then subjected to controlled traction at

2% permanent elongation.

  The static mechanical characteristics in the L and TL directions were measured on the plates, namely the breaking strength Rm (in MPa), the conventional yield strength at 0.2% Ro0.2 (in MPa) and elongation at break A (in% O). The results are summarized in Table 2:

Table 2

Alloy Thickness Rm R0.2 A

2024 40 L 468 362 20.0

2024 40 TL 469 330 17.4

2024 15 L 462 360 21.2

2024 15 TL 467 325 17.6

2034 40 L 534 416 11.2

2034 40 TL 529 393 12.0

2034 15 L 548 431 13.8

2034 15 TL 531 395 14.6

  The invention also relates to the toughness by the critical stress factors Kc and Kc. Kc (in MPa '/ m) in the LT direction, according to ASTM E 561, on CCT specimens, taken at quarter-thickness, of width W = 500 mm, thickness B = 5 mm, and a central notch Machined by electroerosion 2ao = 165 mm, enlarged by fatigue test up to 170 mm. The results are given in Table 3

Table 3

Thickness alloy KI Ka

2024 40 143.4 105.2

2034 40 128.8 97.8

Invention 40 179.7 122

2034 15 136.4 103.7

  The fatigue crack propagation speed da / dn in the LT direction (in mm / cycle) for various AK values (in MPaJm) according to ASTM E 647 was also measured. For this purpose, two CCT test pieces of width W = 200 mm and thickness B = 5 mm taken at a quarter thickness of sheet metal in the LT direction are used. The length of the central notch machined by electroerosion is 30 mm, and this notch is enlarged by fatigue test to 40 mm. The crack velocity measurement test is performed on an MTS machine with a stress in R = 0.05 and a stress of 40 MPa, calculated to obtain an AK value of 10 MPaVm for the starting slit length of 40 mm. The results are given in Table 4

Table 4

  Alloy Ep. AK = 10 AK = 12 AK = 15 AK = 20 AK = 25

  2024 40 9 10-, 1.5 10-4 33.010-4 6 10-4 9 10-3

  2034 40 8 10-, 1.5 10-4 3 10-4 5.7 10-4 1.7 10-3

  invention 40 5.5 10-5 1.7 104 2.0 10-4 4.0 10-4 7.8 10-4

  2034 15 8 0-5 1.5 10-4 3 104 5.2 10 -'4 2.10 0-4

  invention 4.9 10- '6.0 10- 1.3 10-4 2.5 1-4 5.4 10-4 Fatigue tests according to Airbus specification AITM 1-0011 were carried out on test specimens hole length 230 mm, width 50 mm and thickness 7.94

  mm, taken at mid-thickness of the L-shaped plate. The diameter of the hole is 7.94 mm.

  A mean full specimen stress of 80 MPa was applied with 4 alternating stress levels: 85 MPa, 55 MPa, 45 MPa and 35 MPa for 40

  mm, 110, 85, 55 and 45 MPa for 15 mm sheets, with 2 test pieces per level.

  The average values of the service life (in number of cycles) are indicated in table 5. It is noted that, for test pieces with a kerf factor of K. = 2.5, the duration

  fatigue life is improved by more than 20% compared to alloy 2024.

Table 5

  alloy Thickness 80 + 85 80 + 55 80 + 45 80 + 35 mm MPa MPa MPa MPa

2024 40 36044 159721

2034 40 30640 125565 340126 839340

  invention 42933 219753 392680 1018240

2034 15 41040 204038 352957

  Finally, the arrows f in the direction L and TL were measured, as was the recrystallization rate (in%) at the surface, at quarter-thickness and at mid-thickness, determined by

  image analysis after etching of the sample.

  The arrow f is measured as follows. Two bars are taken from the plate of thickness e, one called a bar L, of length b in the direction of the length of the sheet (direction L), of width 25 mm in the direction of the width of the sheet (TL direction) and thickness e according to the full thickness of the sheet (TC direction), the other, called bar

  TL direction, having 25 mm in the L direction, b in the TL direction and e in the TC direction.

  Each bar is machined to mid-thickness and the boom is measured at the end of the bar. This arrow is representative of the level of internal stresses of the sheet and its ability not to deform at machining. The distance 1 between the supports was 180 mm and the length b of the bars 200 mm. Machining is a machining

  Progressive mechanical with passes of about 2 mm. The measurement of the arrow at half

  length is achieved using a comparator with a resolution of one micron. Results for arrows and recrystallization rates are given in Table 6. o10

Table 6

  alloy Thickness fL (gm) fTL (jtm) Surf. / 4 ep. / 2 ep.

2024 40 210 120 79 58 30

2034 40 147 129 12 0 0

invention 40 86 75 46 5 2 He

Claims (12)

  1. Rolled, spun or forged product made of AICuMg alloy, treated by dissolving, quenching and cold forming, intended for the manufacture of aircraft structural elements, of composition (% by weight): Fe + Si <0 , Cu: 3.6 - 4.4 Mg: 1 - 1.5 Mn: 0.5 - 0.8 Zr: 0.08-0.15 other elements: <0.05 each and <0.15 in total.   2. Product according to claim 1, characterized in that Fe + Si <0, 15%.   3. laminated product with a thickness of 6 to 60 mm according to one of claims I or 2,   having, in the quenched and cold-deformed state, a breaking strength Rpm)> 475 MPa and a yield strength RO, 2 (L)> 370 MPa   4. laminated product with a thickness of 6 to 60 mm according to one of claims 1 to 3,   having, in the quenched and cold-deformed state, a critical intensity factor (LT) K [> 170 MPa · lm and K> 120 MPaqm, measured according to ASTM E 561 on notched specimens taken at quarter thickness with W parameters = 500 mm, B = 5 mm and 2ao = 165mm.   5. laminated product according to one of claims 1 to 4, characterized in that   has a critical intensity factor (L-T direction) K, or K <0 increased by at least   % relative to alloy 2024 under the same conditions.   6. laminated product with a thickness of 6 to 60 mm according to one of claims 1 to 5,   having a crack propagation rate (LT) da / dn in the T351 state, measured according to ASTM E 647, on notched specimens taken at quarter thickness with W = 200 mm and B = 5mm: <10-4 mm / cycle for AK = 12 MPa / m <2.5 10-4 mm / cycle for AK = 15 MPa / m and <5 10-4 mm / cycle for AK = 20 MPaVm 7. The laminated product according to one of claims 1 to 6, characterized in that it has an arrow f measured in the L and TL directions after machining at half the thickness of a bar resting on two supports distant a length 1 less than (0.14 12) / e, f being measured in microns, the thickness e of the sheet and the length I being expressed in mm   8. laminated product according to one of claims 3 to 7, having a shelf life   average fatigue, measured on a notched specimen taken at half thickness   L direction, increased by more than 20% with respect to alloy 2024.   9. A method of manufacturing a product according to one of claims 1 to 8,   comprising the following steps: - casting of a plate of the indicated composition - homogenization of this plate between 450 and 500 C, hot transformation by rolling, spinning or forging to the desired product, - optionally cold processing, - setting solution at a temperature between 480 and 505 C, - quenching with cold water, cold deformation to more than 1.5% of permanent deformation,   - natural aging at ambient temperature.   10. Process according to claim 9, characterized in that the hot transformation   is done with an outlet temperature> 420 C, and preferably> 440 C.   11. Use of sheets according to one of claims 3 to 8 for the manufacture of skin of lower wing of airplane wing.   12. Use of profiles according to one of claims 1 or 2 for the manufacture of   wing or aircraft fuselage underside stiffeners.