Method for producing a high strength Al-Zn-Mg-Cu alloy

Abstract

The present invention relates to a method for producing a high strength Al—Zn—Cu—Mg alloy with an improved fatigue crack growth resistance and a high damage tolerance, comprising the steps of casting an ingot with the following composition (in weight percent) Zn 5.5-9.5, Cu 1.5-3.5, Mg 1.5-3.5, Mn<0.25, Zr<0.25, Cr<0.10, Fe<0.25, Si<0.25, Ti<0.10, Hf and/or V<0.25, other elements each less than 0.05 and less than 0.15 in total, balance aluminum, homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working into a worked product of more than 50 mm thickness, solution heat treating, quenching the heat treated product, and artificially ageing the worked and heat-treated product, wherein the ageing step comprises a first heat treatment at a temperature in a range of 105° C. to 135° C. for more than 2 hours and less than 8 hours and a second heat treatment at a higher temperature than 135° C. but below 170° C. for more than 5 hours and less than 15 hours. The invention concerns a weldable plate product of such high strength Al—Zn—Cu—Mg having a thickness of more than 50 mm and an aircraft structural member produced from such alloy.

1 . Method for producing a high strength Al—Zn—Cu—Mg alloy with a high damage tolerance and an improved corrosion resistance, comprising the steps of: 
 a) casting an ingot with the following composition (in weight percent):                                                Zn   5.5 to 9.5         Cu   1.5 to 3.5         Mg   1.5 to 3.5         Mn   <0.25         Zr   <0.25         Cr   <0.10         Fe   <0.25         Si   <0.25         Ti   <0.10                                                      Hf and/or V<0.25     other elements each less than 0.05 and less than 0.15 in total, balance aluminum,    b) homogenizing and/or pre-heating the ingot after casting,    c) hot-working the ingot and optionally cold working into a worked product of more than 50 mm thickness,    d) solution heat-treating,    e) quenching the solution heat treated product, and    f) artificially ageing the worked and heat-treated product, wherein the ageing step comprises a first heat treatment at a temperature in a range of 105° C. to 135° C. for more than 2 hours and less than 8 hours and a second heat treatment at a higher temperature than 135° C. but below 170° C. for more than 5 hours and less than 15 hours to achieve a product with a compression yield strength in L-direction at S/4 of at least 475 MPa, an ultimate tensile strength of at least 510 MPa and an ST elongation at S/2 of at least 3.0%.    
     
     
         2 . Method according to  claim 1 , wherein the ageing step consists of two heat treatments, the first heat treatment is performed for 2 to 5 hours at temperatures in the range of 105° C. to 135° C., and the second heat treatment is performed for 5 to 15 hours at temperatures in the range of 155° C. to 169° C.  
     
     
         3 . Method according to  claim 1 , wherein the first heat treatment is performed at temperatures in the range 115° C. to 125° C.  
     
     
         4 . Method according to  claim 1 , wherein the first heat treatment is performed for 2 to 5 hours at about 120° C.  
     
     
         5 . Method according to  claim 1 , wherein the second heat treatment is performed at temperatures in the range 161° C. to 167° C.  
     
     
         6 . Method according to  claim 1 , wherein the second heat treatment is performed for about 13 hours.  
     
     
         7 . Method according to  claim 1 , wherein the improved corrosion resistance has exfoliation properties (“EXCO”) of EB or better according to ASTM G34.  
     
     
         8 . Method according to  claim 1 , wherein the amount of Mg is in a range of 1.5 to 2.5.  
     
     
         9 . Method according to  claim 8 , wherein the amount of Mg is in a range of 1.6 to 2.3.  
     
     
         10 . Method according to  claim 8 , wherein the amount of Mg is in a range of 1.90 to 2.10.  
     
     
         11 . Method according to  claim 1 , wherein the amount of Cu is in a range of 1.5 to 2.5.  
     
     
         12 . Method according to  claim 11 , wherein the amount of Cu is in a range of 1.6 to 2.3.  
     
     
         13 . Method according to  claim 11 , wherein the amount of Cu is in a range of 1.85 to 2.1 0.  
     
     
         14 . Method according to  claim 1 , wherein the amount of Mg depends on the amount of Zn as follows: [Mg] is in between 2.4-0.1[Zn] and 1.5+0.1[Zn].  
     
     
         15 . Method according to  claim 1 , wherein the amount of Zn is in a range of 5.9 to 6.2.  
     
     
         16 . Method according to  claim 1 , wherein the amount of Zn is in a range of 6.8 to 7.1.  
     
     
         17 . Method according to  claim 1 , wherein the amount of Zn is in a range of 7.8 to 8.1.  
     
     
         18 . Method according to  claim 1 , wherein the amount of Fe is less than 0.15.  
     
     
         19 . Method according to  claim 1 , wherein the amount of Fe is 0.08 or less.  
     
     
         20 . Method according to  claim 1 , wherein the amount of Si is less than 0.10.  
     
     
         21 . Method according to  claim 1 , wherein the amount of Si is 0.04 or less.  
     
     
         22 . Method according to  claim 1 , wherein the amount of Zr is in a range of 0.06 to 0.16.  
     
     
         23 . Method according to  claim 1 , wherein the amount of Mn is 0.08 or less.  
     
     
         24 . Method according to  claim 1 , wherein the amount of Mn is 0.02 or less.  
     
     
         25 . Method according to  claim 1 , wherein after the step e) the quenched solution heat-treated product is stretched or compressed or otherwise cold worked to relieve stresses prior to the ageing practice of the step f).  
     
     
         26 . Method according to  claim 1 , wherein the product is hot-worked by means of rolling.  
     
     
         27 . Method according to  claim 1 , wherein the product is cold-worked by means of rolling.  
     
     
         28 . Method according to  claim 1 , wherein after homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working the ingot into a worked product of more than 60 mm.  
     
     
         29 . Method according to  claim 1 , wherein after homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working the ingot into a worked product of more than 110 mm.  
     
     
         30 . Method according to  claim 1 , wherein after homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working the ingot into a worked product of not more than 220 mm.  
     
     
         31 . Method according to  claim 1 , wherein after homogenizing and/or pre-heating the ingot after casting, hot working the ingot and optionally cold working the ingot into a worked product of not more than 160 mm.  
     
     
         32 . Method according to  claim 1 , wherein said high strength Al—Zn—Cu—Mg alloy is selected from the group consisting of AA7010, AA7×50, AA7040, AA7020, AA7×75, AA7349, AA7×55, and AA7×85.  
     
     
         33 . A plate product of high strength Al—Zn—Cu—Mg alloy produced in accordance with a method as defined in  claim 1  and having a thickness of more than 50 mm.  
     
     
         34 . A plate product of high strength Al—Zn—Cu—Mg alloy produced in accordance with a method as defined in  claim 1  and having a thickness in a range of 110 to 220  
     
     
         35 . A plate product of high strength Al—Zn—Cu—Mg alloy produced in accordance with a method as defined in  claim 1  and having a thickness of more than 60 mm.  
     
     
         36 . A plate product according to  claim 33 , wherein said plate product is a structural member of an aircraft.  
     
     
         37 . A plate product according to  claim 34 , wherein said plate product is a structural member of an aircraft.  
     
     
         38 . A plate product according to  claim 33 , wherein said plate product is a bar or a spar of a wing of an aircraft.  
     
     
         39 . A plate product according to  claim 34 , wherein said plate product is an upper-wing member of an aircraft.  
     
     
         40 . An aircraft structural member produced from a high strength Al—Zn—Cu—Mg alloy produced in accordance with a method as defined in  claim 1 .  
     
     
         41 . An aircraft structural member having a thickness of at least 50 mm and manufactured from a rolled product made of an alloy with a composition, consisting of, in % by weight: 
 Zn 5.5to9.5    Cu 1.5 to 3.5    Mg 1.5 to 3.5    Mn<0.25    Zr<0.25    Cr<0.10    Fe<0.25    Si<0.25    Ti<0.10    Hf and/or V<0.25    other elements each less than 0.05 and less than 0.15 in total, balance aluminum,    and treated by solution heat treating, quenching, and ageing practice consisting essentially of a first heat treatment at a temperature in a range of 105° C. to 135° C. for more than 2 hours and less than 8 hours and a second heat treatment at a higher temperature than 135° C. but below 170° C. for more than 5 hours and less than 15 hours, the product having a compression yield strength in L-direction at S/4 of at least 475 MPa, an ultimate tensile strength of at least 510 MPa and an ST elongation at S/2 of at least 3.0%.    
     
     
         42 . An aircraft structural member according to  claim 41 , wherein the aircraft structural member has a thickness in a range of 50 to 220 mm.  
     
     
         43 . An aircraft structural member according to  claim 41 , wherein the aircraft structural member has a thickness in a range of 60 to 160 mm.  
     
     
         44 . An aircraft structural member according to  claim 41 , wherein the aircraft structural member has a thickness in a range of 110 to 160 mm.  
     
     
         45 . An aircraft structural member according to  claim 41 , forming a part of an aircraft upper wing.  
     
     
         46 . An aircraft structural member according to  claim 41 , forming a spar of an aircraft wing.  
     
     
         47 . An aircraft structural member according to  claim 41 , forming a bar of an aircraft wing.  
     
     
         48 . An aircraft structural member according to  claim 41 , obtained by machining.  
     
     
         49 . An aircraft structural member according to  claim 41 , wherein the ageing practice consists of two heat treatments, the first heat treatment is performed for 2 to 5 hours at temperatures in the range of 105° C. to 135° C., and the second heat treatment is performed for 5 to 15 hours at temperatures in the range of 155° C. to 169° C.  
     
     
         50 . An aircraft structural member according to  claim 41 , wherein the first heat treatment of the ageing practice is performed at temperatures in the range 115° C. to 125° C.  
     
     
         51 . An aircraft structural member according to  claim 41 , wherein the first heat treatment of the ageing practice is performed for 2 to 5 hours at about 120° C.  
     
     
         52 . An aircraft structural member according to  claim 41 , wherein the second heat treatment of the ageing practice is performed at temperatures in the range 161° C. to 167° C.  
     
     
         53 . An aircraft structural member according to  claim 41 , wherein the second heat treatment of the ageing practice is performed for about 13 hours.  
     
     
         54 . An aircraft structural member according to  claim 41 , wherein the improved corrosion resistance has exfoliation properties (“EXCO”) of EB or better according to ASTM G34.  
     
     
         55 . An aircraft structural member according to  claim 40 , wherein in the amount of Mg is in a range of 1.5 to 2.5.  
     
     
         56 . An aircraft structural member according to  claim 55 , wherein in the amount of Mg is in a range of 1.6 to 2.3.  
     
     
         57 . An aircraft structural member according to  claim 55 , wherein in the amount of Mg is in a range of 1.90 to 2.10.  
     
     
         58 . An aircraft structural member according to  claim 41 , wherein the amount of Cu is in a range of 1.5 to 2.5.  
     
     
         59 . An aircraft structural member according to  claim 58 , wherein the amount of Cu is in a range of 1.6 to 2.3.  
     
     
         60 . An aircraft structural member according to  claim 58 , wherein the amount of Cu is in a range of 1.85 to 2.10.  
     
     
         61 . An aircraft structural member according to  claim 41 , wherein the amount of Mg depends on the amount of Zn as follows: [Mg] is in between 2.4-0.1[Zn] and 1.5+0.1[Zn].  
     
     
         62 . An aircraft structural member according to  claim 41 , wherein the amount of Zn is in a range of 5.9 to 6.2.  
     
     
         63 . An aircraft structural member according to  claim 41 , wherein the amount of Zn is in a range of 6.8 to 7.1.  
     
     
         64 . An aircraft structural member according to  claim 41 , wherein the amount of Zn is in a range of 7.8 to 8.1.  
     
     
         65 . An aircraft structural member according to  claim 41 , wherein the amount of Fe is less than 0.15.  
     
     
         66 . An aircraft structural member according to  claim 41 , wherein the amount of Si is less than 0.10.  
     
     
         67 . An aircraft structural member according to  claim 41 , wherein the amount of Zr is in a range of 0.06 to 0.16.  
     
     
         68 . An aircraft structural member according to  claim 41 , wherein the amount of Mn is in a range of 0.08 or less.  
     
     
         69 . An aircraft structural member according to  claim 41 , wherein after the quenching following the solution heat-treating the product is stretched or compressed or otherwise cold worked to relieve stresses prior to the ageing practice.