Method for Fabricating Seal-Free Multi-Metallic Thrust Chamber Liner

• US 20190329355A1
• Filed: 04/27/2018
• Published: 10/31/2019
• Est. Priority Date: 04/27/2018
• Status: Abandoned Application

• Current Assignee: United States of America as represented by the Administrator of NASA
• Original Assignee: United States of America as represented by the Administrator of NASA
• Inventors: Gradl, Paul R., Brandsmeier, William C. C., Medina, Cory R., Protz, Christopher Stephen, Mireles, Omar
• Application Number:
  US15/965,389
• Publication Number:
  US 20190329355A1
• Time in Patent Office: Days
• CPC Class Codes: B22F 2005/005 Article surface comprising protrusions
• CPC Class Codes: B22F 2207/01 Composition gradients
• CPC Class Codes: B22F 2999/00 Aspects linked to processes or compositions used in powder metallurgy
• CPC Class Codes: B22F 5/106 Tube or ring forms
• CPC Class Codes: B22F 7/062 involving the connection or repairing of preformed parts
• CPC Class Codes: B22F 7/08 with one or more parts not made from powder B22F7/062 takes precedence
• CPC Class Codes: B23K 15/0053 Seam welding
• CPC Class Codes: B23K 15/0093 characterised by the properties of the materials to be welded
• CPC Class Codes: B23K 2101/001 Turbines
• CPC Class Codes: B23K 2101/34 Coated articles , e.g. plated or painted; Surface treated articles
• CPC Class Codes: B23K 2103/05 Stainless steel
• CPC Class Codes: B23K 2103/12 Copper or alloys thereof
• CPC Class Codes: B23K 2103/18 Dissimilar materials
• CPC Class Codes: B23K 26/0006 taking account of the properties of the material involved B23K26/32, B23K26/40 take precedence
• CPC Class Codes: B23K 26/08 Devices involving relative movement between laser beam and workpiece
• CPC Class Codes: B23K 26/282 of tube sections
• CPC Class Codes: B23K 26/323 involving parts made of dissimilar metallic material
• CPC Class Codes: B23K 26/342 Build-up welding
• CPC Class Codes: B23P 15/008 Rocket engine parts, e.g. nozzles, combustion chambers
• CPC Class Codes: B29C 70/682 Preformed parts characterised by their structure, e.g. form
• CPC Class Codes: B33Y 10/00 : Processes of additive manufacturing
• CPC Class Codes: B33Y 40/00 : Auxiliary operations or equipment, e.g. for material handling
• CPC Class Codes: B33Y 80/00 : Products made by additive manufacturing
• CPC Class Codes: C23C 24/106 : Coating with metal alloys or metal elements only
• CPC Class Codes: C23C 28/00 : Coating for obtaining at least two superposed coatings either by methods not provided for in a single one of groups C23C2/00 - C23C26/00 or by combinations of methods provided for in subclasses C23C and C25C or C25D
• CPC Class Codes: C23C 28/021 : including at least one metal alloy layer
• CPC Class Codes: C23C 28/023 : only coatings of metal elements only
• CPC Class Codes: F02K 9/972 : Fluid cooling arrangements for nozzles F02K9/64 takes precedence
• CPC Class Codes: F02K 9/974 : Nozzle- linings; Ablative coatings
• CPC Class Codes: F05D 2230/31 : Layer deposition

Abstract

A method for fabricating a thrust chamber liner for a rocket engine commences with a ring made from a first material on a build plate. A base layer of a second material in powder form is deposited on the exposed axial end of the ring. A laser beam is directed towards the base layer and the ring such that energy associated with the laser beam melts the base layer and a portion of the ring adjacent to the base layer. A melted portion of the base layer intermixes with a melted portion of the ring. Following this step, additional layers of the second material are deposited on the base layer. The first axial end of the ring is then exposed and additional layers of the first material are deposited on the first axial end of the ring.

Claims

1. A method for fabricating a thrust chamber liner for a rocket engine, comprising the steps of:

positioning a ring made from a first material on a build plate, wherein a first axial end of said ring rests on said build plate and a second axial end of said ring is exposed;

depositing a base layer of a second material in powder form on said second axial end of said ring;

directing a laser beam towards said base layer and said ring, wherein energy associated with said laser beam melts said base layer and a portion of said ring adjacent to said base layer, and wherein a melted portion of said base layer intermixes with a melted portion of said ring;

depositing, following said step of directing, additional layers of said second material on said base layer;

exposing said first axial end of said ring; and

depositing additional layers of said first material on said first axial end of said ring.

2. A method according to claim 1, wherein said base layer and said additional layers of said second material comprise a main combustion chamber liner for a rocket engine.

3. A method according to claim 1, wherein said ring and said additional layers of said first material comprise a nozzle liner for a rocket engine.

4. A method according to claim 1, wherein said first material is selected from the group consisting of stainless steel and a superalloy.

5. A method according to claim 1, wherein said second material comprises a copper-alloy.

6. A method according to claim 1, further comprising the step of wrapping, following said steps of depositing said additional layers of said second material and depositing said additional layers of said first material, a composite material on an outer surface of said first material and an outer surface of said second material.

7. A method according to claim 6, wherein said composite material is selected from the group consisting of carbon fiber composites, fiber-reinforced polymer composites, metal matrix composites, and ceramic matrix composites.

8. A method for fabricating a thrust chamber liner for a rocket engine, comprising the steps of:

providing a ring made from a first material on a build plate, wherein a first axial end of said ring rests on said build plate and a second axial end of said ring is exposed, said first material being selected from the group consisting of stainless steel and a superalloy;

depositing a base layer of a second material in powder form on said second axial end of said ring, said second material comprising a copper-alloy;

directing a laser beam towards said base layer and said ring, wherein energy associated with said laser beam melts said base layer and a portion of said ring adjacent to said base layer, and wherein an integrated region is generated from a melted portion of said base layer intermixed with a melted portion of said ring, said integrated region having a gradient function associated therewith;

depositing, following said step of directing, additional layers of said second material on said base layer;

exposing said first axial end of said ring; and

depositing additional layers of said first material on said first axial end of said ring.

9. A method according to claim 8, wherein said base layer and said additional layers of said second material comprise a main combustion chamber liner for a rocket engine.

10. A method according to claim 8, wherein said ring and said additional layers of said first material comprise a nozzle liner for a rocket engine.

11. A method according to claim 8, further comprising the step of wrapping, following said steps of depositing said additional layers of said second material and depositing said additional layers of said first material, a composite material on an outer surface of said first material and an outer surface of said second material.

12. A method according to claim 11, wherein said composite material is selected from the group consisting of carbon fiber composites, fiber-reinforced polymer composites, metal matrix composites, and ceramic matrix composites.

13. A method for fabricating a thrust chamber liner for a rocket engine, comprising the steps of:

providing a nozzle inlet made from a first material on a build plate, wherein a first axial end of said nozzle inlet rests on said build plate and a second axial end of said nozzle inlet is exposed;

depositing a base layer of a second material in powder form on said second axial end of said nozzle inlet;

directing a laser beam towards said base layer and said nozzle inlet, wherein energy associated with said laser beam melts said base layer and a portion of said nozzle inlet adjacent to said base layer, and wherein an integrated region is generated from a melted portion of said base layer and a melted portion of said nozzle inlet;

building, following said step of directing, a main combustion chamber liner on said base layer, said main combustion chamber liner made from said second material;

exposing said first axial end of said nozzle inlet; and

building a nozzle liner on said first axial end of said nozzle inlet, said nozzle liner made from said first material.

14. A method according to claim 13, wherein said first material is selected from the group consisting of stainless steel and a superalloy.

15. A method according to claim 13, wherein said second material comprises a copper-alloy.

16. A method according to claim 13, further comprising the step of wrapping, following said steps of building, a composite material on an outer surface of said main combustion chamber liner and an outer surface of said nozzle liner.

17. A method according to claim 16, wherein said composite material is selected from the group consisting of carbon fiber composites, fiber-reinforced polymer composites, metal matrix composites, and ceramic matrix composites.