Nano-metal coated vane component for gas turbine engines and method of manufacturing same

1. A method of manufacturing a vane component of a gas turbine engine, the method comprising:

forming, using a rapid manufacturing process, an internal space frame structure out of a substrate material and forming, using the rapid manufacturing process, a core of the vane component with a cavity therein by applying the substrate material to the space frame structure such that the core is at least partially hollow; and

applying a coating of nanocrystalline metal onto the core, the nanocrystalline metal fully encapsulating the core and defining an outer structural surface of the vane component.

2. The method of claim 1 , wherein the step of forming the core further comprises forming the core out of at least one of a composite material and a polymer material using the rapid manufacturing process.

3. The method of claim 1 , wherein the rapid manufacturing process includes at least one of rapid freeform fabrication technologies, stereolithography, 3D inkjet printing, 3D printing, laser metal deposition, selective laser sintering, direct metal deposition, direct laser powder deposition.

4. The method of claim 1 , wherein the step of applying the coating further comprises plating the nanocrystalline metal coating onto the core.

5. The method of claim 4 , further comprising electroplating the nanocrystalline metal coating onto the core.

6. The method of claim 4 , wherein the step of plating the nanocrystalline metal coating includes plating a single layer of nano-scale nickel base, ferrous base or cobalt base metal onto the core.

7. A method of manufacturing a vane component of a gas turbine engine, the method comprising the steps of:

i) using a computer-controlled rapid manufacturing machine to produce a pre-form core of said vane component out of a core substrate material, the core defining a cavity therein such that the core is at least partially hollow, the core substrate material comprising composite, polymer or metal, the computer-controlled rapid manufacturing machine being controlled by a computer readable information source representative of the core of the vane component;

ii) applying the core substrate material to fabricate a space frame structure within the cavity using the computer-controlled rapid manufacturing machine to form the pre-form core; and

iii) plating said pre-form core produced by the computer-controlled rapid manufacturing machine with a nano-grain size metallic coating to encapsulate the pre-form core with said nano-grain size metallic coating and define an outer structural surface of the vane component.

8. The method of claim 7 , further comprising, prior to said step (i), producing the computer readable information source representative of the core of the vane component, the information source comprising a two-dimensional or a three-dimensional image of said vane component.

9. The method of claim 7 , wherein said vane component is a vane assembly of said gas turbine engine, the vane assembly including at least one of a hollow vane, a hollow vane shroud, a hollow vane shroud pack and a vane ring assembly.

10. The method of claim 7 , wherein said step (iii) further comprises using at least one of electroplating and laser metal deposition to apply said nano-grain size metallic coating onto said pre-form core of said vane component.

11. The method of claim 7 , wherein the step (ii) further comprises varying at least one of a space frame density or cell size, space frame geometry, space frame arm thickness, core airfoil skin thickness, and shape of the pre-form core cavity depending on dynamic requirements of said vane component.

12. The method of claim 7 , further comprising varying at least one of metal composition, grain size, density and thicknesses of said nano-grain size metal coating, in order to optimize at least one of the characteristics of strength, stiffness, ductibility, galvanic potential, thermal expansion, corrosion resistance, erosion resistance, impact resistance, and weight of said vane component.

13. The method of claim 7 , wherein said vane component includes a hollow vane airfoil and wherein said cavity is connected through openings in shrouds of the pre-form core to provide for hot air deicing of the hollow vane airfoil.

14. The method of claim 13 , further comprising the step of closing said cavity with a closure member secured within the openings in said cavity of said vane airfoil by adhesive bonding or laser welding.

15. The method as claimed in claim 7 , further comprising forming the space frame structure of the pre-form core with a hexadron or tetrahedron structural geometrical shape.