Method for manufacturing a structural component for a turbomachine

Abstract

A method for producing engine components with a complex geometric structure particularly for an aircraft engine, includes providing a first component that has a first sintered state with a first degree of compaction; providing at least one second component that has a second sintered state with a second degree of compaction; wherein the second degree of compaction is lower than the first degree of compaction; and wherein the at least one second component has an opening, into which a portion of the first component is introducible; and wherein the opening is dimensioned to close substantially completely during a sintering joining process, to form a force-fitting sintered joint with the first component; introducing the first component into the opening; and carrying out the sintering joining process.

1 . A method for producing a part for a turbomachine, in particular for an aircraft engine, having the steps of:
 providing a first component that has
 a first sintered state with a first degree of compaction; 
   providing at least one second component that has
 a second sintered state with a second degree of compaction; 
   wherein the second degree of compaction is lower than the first degree of compaction; and   wherein the at least one second component has an opening, into which a portion of the first component is introducible; and   wherein the opening is dimensioned in such a way as to close substantially completely during a sintering joining process, in order to form a force-fitting sintered joint with the first component;   introducing the first component into the opening; and   carrying out the sintering joining process.   
     
     
         2 . The method according to  claim 1 , wherein, before the sintering joining process is carried out, a slot is provided, in particular in the region of the opening, between the first component and second component, and wherein the slot is closed substantially completely after the sintering joining process has been carried out. 
     
     
         3 . The method according to  claim 1 , wherein the second component shrinks more than the first component while the sintering joining process is being carried out, with the result that the opening or the slot closes substantially completely. 
     
     
         4 . The method according to  claim 1 , wherein a contact pressure between the first component and second component while the sintering joining process is being carried out is settable by way of a suitable dimensioning, choice and/or setting of the opening or of the slot. 
     
     
         5 . The method according to  claim 1 , wherein a directed force action of the contact pressure is settable by way of a suitable dimensioning, choice and/or setting of the opening or of the slot. 
     
     
         6 . The method according to  claim 1 , wherein a joining paste is provided between the first component and the second component. 
     
     
         7 . The method according to  claim 1 , wherein the first component and the second component have a shape such that, after the sintering joining process has been carried out, a form-fitting joint is producible. 
     
     
         8 . The method according to  claim 1 , wherein the first component and the second component are formed from a material from the group consisting of a ceramic material, a metal material, a material containing nickel, a material containing cobalt, IN713LC alloy, IN718 alloy, CM247 alloy, Haynes25 alloy and Hastelloy X alloy. 
     
     
         9 . The method according to  claim 1 , wherein the first component and the second component have a combination of sintered states from the group consisting of brown part and brown part, brown part and pre-sintered component, pre-sintered component and pre-sintered component, pre-sintered component and fully sintered component, and fully sintered component and fully sintered component. 
     
     
         10 . The method according to  claim 1 , wherein the part is selected from the group consisting of blade airfoil, blade airfoil with tip shroud, a pair of blade airfoils, part for the nozzle guide vane and liner for the combustion chamber. 
     
     
         11 . A part for a turbomachine, in particular for a gas turbine engine for an aircraft, and more particularly an aircraft engine, produced by the method according to  claim 1 . 
     
     
         12 . Gas turbine engine for an aircraft, having at least one part according to  claim 11 .