Al-si-mg aluminum alloy aircraft structural component production method

Abstract

An aircraft structural component, particularly a fuselage component, production method using rolled, extruded or forged products made of aluminum alloy with the following composition (% by weight): Si: 0.7-1.3; Mg: 0.6-1.1; Cu: 0.5-1.1; Mn: 0.3-0.8; Zn<1; Fe<0.30; Zr<0.20; Cr<0.25; other elements <0.05 each and <0.15 in total; remainder aluminum. The product is treated by steps of: solution heat treatment of the product between 540 and 570° C.; quenching the solution heat treated product; producing the structural component by forming of the product, and welding if required; ageing of the structural component, in one or more stages, for which the total equivalent time at 175° C. expressed in hours is between (−160+57γ) and (−184+69γ), γ being the sum of Si+2Mg+2Cu contents in % by weight. The invention results in an improvement in tolerance to damage with no loss of other operating parameters.

What is claimed is:  
     
         1 . A method for production of aircraft structural components using rolled, extruded or forged products made of an aluminum alloy, comprising the steps of: 
 casting a blank with a composition consisting essentially of, in % by weight:    Si: 0.7-1.3; Mg: 0.6-1.1; Cu: 0.5-1.1; Mn: 0.3-0.8; Zn<1; Fe<0.30; Zr<0.20; Cr<0.25; other elements <0.05 each and <0.15 in total; remainder aluminum;    hot, and possibly cold, transforming of said blank to obtain a product;    solution heat treating the product between 540 and 570° C.;    quenching the solution heat treated product;    forming and optionally welding the quenched product to produce the structural component;    ageing the structural component, in one or more stages, for a total equivalent time at 175° C. expressed in hours between (−160+57γ) and (−184+69γ), γ being Si+2Mg+2Cu in % by weight.    
     
     
         2 . Method according to  claim 1 , wherein the blank is homogenized at a temperature between 540 and 570° C.  
     
     
         3 . Method according to  claim 1 , wherein the quenched product is subjected, before ageing, to cold-working resulting in a permanent elongation between 1 and 15%.  
     
     
         4 . Method according to  claim 3 , wherein the cold working results in a permanent elongation between 2 and 10%.  
     
     
         5 . Method according to  claim 1 , wherein total equivalent time at 175° C., in hours is between (−150+57γ) and (−184+69γ).  
     
     
         6 . Method according to  claim 1 , wherein the product has a composition, in % by weight: 
 Si: 0.7-1.1; Mg: 0.6-0.9; Cu: 0.5-0.7; Mn: 0.3-0.8; Zr<0.2; Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg: 2.0-2.6; other elements <0.05 each and <0.15 in total; remainder alumnium.    
     
     
         7 . Method according to  claim 6 , wherein Si+2Mg is between 2.3 and 2.6.  
     
     
         8 . Method according to  claim 6 , wherein total equivalent ageing time at 175° C. is between 40 and 65 hours.  
     
     
         9 . Method according to  claim 3 , additionally comprising assembly of sheets and profiles produced by the method, wherein the profiles undergo, before said assembly and said ageing, an additional cold-working step in relation to the sheets, such that resistance to intercrystalline corrosion of the profiles is at the same level as the sheets.  
     
     
         10 . Aircraft fuselage component, produced using a rolled, extruded or forged product made of an aluminum alloy consisting essentially of, in % by weight: 
 Si: 0.7-1.1; Mg: 0.6-0.9; Cu: 0.5-0.7; Mn: 0.3-0.8; Zr<0.2; Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg: 2-2.6; other elements <0.05 each and <0.15 in total; remainder aluminum;    subjected to a solution heat treatment, quenching, shaping and ageing in a T78 temper with a total equivalent time at 175° C. between 40 and 65 hours.    
     
     
         11 . Fuselage component according to  claim 10 , wherein Si+2Mg is between 2.3 and 2.6.  
     
     
         12 . Fuselage component according to  claim 10 , having, in TL direction, a yield strength R 0.2 >330 MPa, an ultimate tensile strength R m >360 MPa and an elongation A>8%.  
     
     
         13 . Fuselage component according to  claim 10 , having a plane strain fracture toughness in T-L direction, with at least one of the properties: 
 K R  (Δa=20 mm)>90 MPa{square root}m;    K R  (Δa=40 mm)>115 MPa{square root}m;    K c0 >80 MPa{square root}m;    K c >110 MPa{square root}m.    
     
     
         14 . Fuselage component according to  claim 10 , having a plane strain fracture toughness in L-T direction such that: 
 K c0 >90 MPa{square root}m or K c >130 MPa{square root}m.    
     
     
         15 . Fuselage component according to  claim 10 , having a crack growth rate da/dn, measured in T-L direction for R=0.1, less than: 
 2 10 −3  mm/cycle for ΔK=20 MPa{square root}m;    4 10 −3  mm/cycle for ΔK=25 MPa{square root}m;    8 10 −3  mm/cycle for ΔK=30 MPa{square root}m.    
     
     
         16 . Fuselage component produced using a rolled, extruded or forged product made of alloy consisting essentially of, in % by weight: 
 Si: 0.7-1.1; Mg: 0.6-0.9; Cu: 0.5-0.7; Mn: 0.3-0.8; Zr<0.2; Fe<0.2; Zn<0.5; Cr<0.25; Mg/Si<1; Si+2 Mg: 2-2.6; other elements <0.05 each and <0.15 in total; the remainder aluminum;    subjected to a solution heat treatment, quenching, shaping and ageing in T6 temper.    
     
     
         17 . Fuselage component according to  claim 16 , having in TL direction, a yield strength R 0.2 >350 MPa, an ultimate tensile strength R m >380 MPa and an elongation A>6%.  
     
     
         18 . Fuselage component according to  claim 16 , having a plane strain fracture toughness in T-L direction, with at least one of the properties: 
 K R  (Δa=20 mm)>95 MPa{square root}m;    K R  (Δa=40 mm)>120 MPa{square root}m;    K c0 >85 MPa{square root}m;    K c >115 MPa{square root}m.    
     
     
         19 . Fuselage component according to  claim 16 , having a plane strain fracture toughness in L-T direction such that: 
 K c0 >100 MPa{square root}m or K c >150 MPa{square root}m.    
     
     
         20 . Fuselage component according to  claim 16 , having a crack growth rate da/dn, measured in T-L direction for R=0.1, less than: 
 2 10 −3  mm/cycle for ΔK=20 MPa{square root}m;    4 10 −3  mm/cycle for ΔK=25 MPa{square root}m;    8 10 −3  mm/cycle for ΔK=30 MPa{square root}m.