US20020014070A1 - Rocket engine having a transition attachment between a combustion chamber and an injector - Google Patents
Rocket engine having a transition attachment between a combustion chamber and an injector Download PDFInfo
- Publication number
- US20020014070A1 US20020014070A1 US09/772,196 US77219601A US2002014070A1 US 20020014070 A1 US20020014070 A1 US 20020014070A1 US 77219601 A US77219601 A US 77219601A US 2002014070 A1 US2002014070 A1 US 2002014070A1
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- Prior art keywords
- combustion chamber
- injector
- annular
- wall
- attachment
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Abandoned
Links
- 238000002485 combustion reaction Methods 0.000 title claims abstract description 59
- 230000007704 transition Effects 0.000 title claims abstract description 17
- 239000010955 niobium Substances 0.000 claims description 28
- GUCVJGMIXFAOAE-UHFFFAOYSA-N niobium atom Chemical compound [Nb] GUCVJGMIXFAOAE-UHFFFAOYSA-N 0.000 claims description 28
- 238000000034 method Methods 0.000 claims description 13
- RTAQQCXQSZGOHL-UHFFFAOYSA-N Titanium Chemical compound [Ti] RTAQQCXQSZGOHL-UHFFFAOYSA-N 0.000 claims description 12
- 239000010936 titanium Substances 0.000 claims description 12
- 229910052719 titanium Inorganic materials 0.000 claims description 12
- 229910052702 rhenium Inorganic materials 0.000 claims description 9
- WUAPFZMCVAUBPE-UHFFFAOYSA-N rhenium atom Chemical compound [Re] WUAPFZMCVAUBPE-UHFFFAOYSA-N 0.000 claims description 9
- 238000003466 welding Methods 0.000 claims description 7
- BASFCYQUMIYNBI-UHFFFAOYSA-N platinum Chemical compound [Pt] BASFCYQUMIYNBI-UHFFFAOYSA-N 0.000 claims description 4
- 238000005219 brazing Methods 0.000 claims description 3
- 238000005229 chemical vapour deposition Methods 0.000 claims description 3
- 238000005253 cladding Methods 0.000 claims description 2
- 239000002360 explosive Substances 0.000 claims description 2
- 229910052697 platinum Inorganic materials 0.000 claims description 2
- 238000000151 deposition Methods 0.000 claims 5
- 238000007750 plasma spraying Methods 0.000 claims 1
- 229910045601 alloy Inorganic materials 0.000 description 16
- 239000000956 alloy Substances 0.000 description 16
- 239000003380 propellant Substances 0.000 description 8
- 239000000446 fuel Substances 0.000 description 7
- 229910052751 metal Inorganic materials 0.000 description 7
- 239000002184 metal Substances 0.000 description 7
- 239000007800 oxidant agent Substances 0.000 description 7
- 239000000203 mixture Substances 0.000 description 6
- 239000000567 combustion gas Substances 0.000 description 5
- PXHVJJICTQNCMI-UHFFFAOYSA-N Nickel Chemical compound [Ni] PXHVJJICTQNCMI-UHFFFAOYSA-N 0.000 description 4
- KDLHZDBZIXYQEI-UHFFFAOYSA-N Palladium Chemical compound [Pd] KDLHZDBZIXYQEI-UHFFFAOYSA-N 0.000 description 4
- 230000008901 benefit Effects 0.000 description 4
- 239000007789 gas Substances 0.000 description 4
- 239000000463 material Substances 0.000 description 4
- 238000010894 electron beam technology Methods 0.000 description 3
- 229910052735 hafnium Inorganic materials 0.000 description 3
- VBJZVLUMGGDVMO-UHFFFAOYSA-N hafnium atom Chemical compound [Hf] VBJZVLUMGGDVMO-UHFFFAOYSA-N 0.000 description 3
- 238000004519 manufacturing process Methods 0.000 description 3
- 238000002156 mixing Methods 0.000 description 3
- 239000010948 rhodium Substances 0.000 description 3
- MHOVAHRLVXNVSD-UHFFFAOYSA-N rhodium atom Chemical compound [Rh] MHOVAHRLVXNVSD-UHFFFAOYSA-N 0.000 description 3
- MWUXSHHQAYIFBG-UHFFFAOYSA-N Nitric oxide Chemical compound O=[N] MWUXSHHQAYIFBG-UHFFFAOYSA-N 0.000 description 2
- 229910001260 Pt alloy Inorganic materials 0.000 description 2
- 229910000629 Rh alloy Inorganic materials 0.000 description 2
- 239000000919 ceramic Substances 0.000 description 2
- 230000007797 corrosion Effects 0.000 description 2
- 238000005260 corrosion Methods 0.000 description 2
- WFPZPJSADLPSON-UHFFFAOYSA-N dinitrogen tetraoxide Chemical compound [O-][N+](=O)[N+]([O-])=O WFPZPJSADLPSON-UHFFFAOYSA-N 0.000 description 2
- 230000003628 erosive effect Effects 0.000 description 2
- 229910052741 iridium Inorganic materials 0.000 description 2
- GKOZUEZYRPOHIO-UHFFFAOYSA-N iridium atom Chemical compound [Ir] GKOZUEZYRPOHIO-UHFFFAOYSA-N 0.000 description 2
- 229910052759 nickel Inorganic materials 0.000 description 2
- 229910052763 palladium Inorganic materials 0.000 description 2
- 229910001069 Ti alloy Inorganic materials 0.000 description 1
- 229910052782 aluminium Inorganic materials 0.000 description 1
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical compound [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 238000010586 diagram Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000010438 heat treatment Methods 0.000 description 1
- 238000002347 injection Methods 0.000 description 1
- 239000007924 injection Substances 0.000 description 1
- 238000005304 joining Methods 0.000 description 1
- 239000007788 liquid Substances 0.000 description 1
- 239000000155 melt Substances 0.000 description 1
- 238000002844 melting Methods 0.000 description 1
- 230000008018 melting Effects 0.000 description 1
- 238000012986 modification Methods 0.000 description 1
- 230000004048 modification Effects 0.000 description 1
- HDZGCSFEDULWCS-UHFFFAOYSA-N monomethylhydrazine Chemical compound CNN HDZGCSFEDULWCS-UHFFFAOYSA-N 0.000 description 1
- TWNQGVIAIRXVLR-UHFFFAOYSA-N oxo(oxoalumanyloxy)alumane Chemical compound O=[Al]O[Al]=O TWNQGVIAIRXVLR-UHFFFAOYSA-N 0.000 description 1
- RVTZCBVAJQQJTK-UHFFFAOYSA-N oxygen(2-);zirconium(4+) Chemical compound [O-2].[O-2].[Zr+4] RVTZCBVAJQQJTK-UHFFFAOYSA-N 0.000 description 1
- 239000011819 refractory material Substances 0.000 description 1
- 239000003870 refractory metal Substances 0.000 description 1
- 229910052703 rhodium Inorganic materials 0.000 description 1
- 239000007921 spray Substances 0.000 description 1
- ZCUFMDLYAMJYST-UHFFFAOYSA-N thorium dioxide Chemical compound O=[Th]=O ZCUFMDLYAMJYST-UHFFFAOYSA-N 0.000 description 1
- 229910003452 thorium oxide Inorganic materials 0.000 description 1
- 229910052720 vanadium Inorganic materials 0.000 description 1
- LEONUFNNVUYDNQ-UHFFFAOYSA-N vanadium atom Chemical compound [V] LEONUFNNVUYDNQ-UHFFFAOYSA-N 0.000 description 1
- 229910001928 zirconium oxide Inorganic materials 0.000 description 1
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
- F02K9/62—Combustion or thrust chambers
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
- F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
- F02K9/60—Constructional parts; Details not otherwise provided for
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
- F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
- F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
- F05D2230/233—Electron beam welding
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/23—Manufacture essentially without removing material by permanently joining parts together
- F05D2230/232—Manufacture essentially without removing material by permanently joining parts together by welding
- F05D2230/237—Brazing
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/10—Metals, alloys or intermetallic compounds
- F05D2300/13—Refractory metals, i.e. Ti, V, Cr, Zr, Nb, Mo, Hf, Ta, W
- F05D2300/133—Titanium
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2300/00—Materials; Properties thereof
- F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
- F05D2300/611—Coating
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
- Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
- Y10T29/00—Metal working
- Y10T29/49—Method of mechanical manufacture
- Y10T29/49346—Rocket or jet device making
Definitions
- a typical liquid-fueled rocket engine includes a generally cylindrical combustion chamber, with an injector attached to its injector end and a flared nozzle attached to its nozzle end.
- a liquid propellant including fuel and an oxidizer flows through injector ports in the injector and into the combustion chamber. The propellant is ignited in the combustion chamber. The hot gas resulting from the combustion expands through the nozzle and drives the rocket engine and the attached rocket structure in the direction opposite to that in which the nozzle is pointed.
- the present invention provides a rocket engine with an improved attachment between the injector and the combustion chamber.
- the rocket engine may be operated reliably at higher temperatures than possible with prior attachment procedures.
- the fabrication approach uses separately known techniques in a new way.
- a rocket engine comprises a combustion chamber comprising an annular wall, an injector, and an attachment between the combustion chamber and the injector.
- the attachment comprises an annular metallic deposit joined to a first region of the wall of the combustion chamber, an annular transition ring structure, a first weldment between a first welded portion of the transition ring structure and the metallic deposit, and a second weldment between a second welded portion of the transition ring structure and the injector.
- the first region of the wall of the combustion chamber, to which the metallic deposit is joined, is typically an outer surface of the chamber wall.
- the metallic deposit may be deposited by any operable technique. Suitably, it is deposited by chemical vapor deposition. The result is a good bond between the metallic deposit and the wall of the combustion chamber. Such a good bond is not easily attained for refractory metals such as the rhenium used in high-performance thrust chamber walls.
- the metallic deposit is suitably columbium.
- a metal identified generically includes both the pure metal and its alloys containing at least about 50 percent by weight of the pure metal.
- “columbium” includes both pure columbium and its alloys.
- An injector 38 is attached to the injector end 34 of the combustion chamber 22 .
- the injector 38 may be of any operable design, but is typically a plate having a plurality of injector ports 40 therein, of which only two are shown in FIGS. 1 and 2.
- the propellant is provided to the combustion chamber 22 through the injector ports 40 .
- Some of injector ports 40 a are supplied with a fuel through a fuel valve 42
- others of the injector ports 40 b are supplied with an oxidizer through an oxidizer valve 44 .
- the fuel and oxidizer flow through their respective injector ports 40 into the interior of the combustion chamber 22 and mix together.
- the mixture is hypergolic, as in the case of monomethylhydrazine (fuel) and nitrogen tetroxide/3 percent nitric oxide (oxidizer), the mixture ignites spontaneously. In other cases where the mixture does not spontaneously ignite, an ignitor (not shown) would be provided.
- the gaseous combustion products of the combustion expand rearwardly and outwardly through the nozzle 28 and drive the rocket engine 20 , and the spacecraft to which it is attached, in the opposite direction.
- a generally cylindrical annular step collar 46 is optionally but suitably fitted within the combustion chamber 22 at and adjacent to the injector end 34 thereof.
- the step collar 46 also protects the inner cylindrical surface 32 from contact with the hot combustion gas when the rocket engine is fired, along the upper end of the combustion chamber, closest to the injector 38 , where the greatest corrosion and erosion damage from the combustion product is expected.
- the step in the step collar 46 encourages downstream turbulent mixing of the injected fuel and oxidizer to achieve maximum burning efficiency and specific impulse of the engine.
- the attachment 48 further includes an annular adaptor ring 54 joined to the step collar 46 (if present), suitably by a joint 56 , which is most suitably a brazed joint.
- the adaptor ring 54 is joined by a joint 58 at an injector end 60 to the injector 38 , and is joined by a joint 62 at a chamber end 64 to the annular metallic deposit 50 .
- the joints 58 and 62 are suitably both welded joints. This structure is not altered in the absence of the step collar 46 .
- the wall 24 of the combustion chamber 22 is made of rhenium with a 0.003-0.005 inch thick coating of iridium on the inner cylindrical surface 32 .
- a metal identified generically includes both the pure metal and its alloys containing at least about 50 percent by weight of the pure metal.
- rhenium includes both pure rhenium and its alloys.
- the metallic deposit is suitably made of relatively pure columbium or C 103 alloy having a composition of 10 weight percent hafnium, 1 weight percent titanium, balance columbium.
- the step collar 46 is suitably made of a material having a high melting point and good corrosion/erosion resistance in the combustion environment of the combustion chamber 22 , such as an alloy of platinum and rhodium, an alloy of columbium, or a ceramic.
- a suitable alloy of platinum and rhodium is 90 percent by weight platinum, balance rhodium.
- a suitable alloy of columbium is 10 percent by weight hafnium, 1 percent by weight titanium, balance columbium.
- a suitable ceramic is aluminum oxide, thorium oxide, or yttria-stabilized zirconium oxide.
- the adaptor ring 54 is suitably made of columbium, most suitably C 103 alloy having a composition of 10 weight percent hafnium, 1 weight percent titanium, balance columbium.
- the injector 38 is suitably made of titanium alloy of 6 weight percent aluminum, 4 weight percent vanadium, balance titanium.
- the joint 56 is suitably brazed with a braze alloy of 60 weight percent palladium, 40 weight percent nickel. This braze alloy is compatible with corrosive combustion gas products, does not substantially diffuse into the structure on either side of the joint, and does not leak propellant gas.
- the joints 58 and 62 are suitably electron beam welds.
- the metallic deposit 50 is present to join the adaptor ring 54 to the wall 24 of the combustion chamber 22 .
- the columbium adaptor ring 54 does not readily braze or weld directly to the relatively thin rhenium wall 24 of the combustion chamber 22 with the required strength and soundness.
- the metallic deposit 50 is first deposited onto the wall 24 , and then the adaptor ring 54 is welded to the deposit 50 .
- FIG. 3 illustrates one approach to the fabrication of the rocket engine 20 .
- the adaptor ring 54 is provided, numeral 70
- the step collar 46 is provided, numeral 72 .
- the adaptor ring 54 and the step collar 46 are brazed together, numeral 74 .
- a shim of the braze alloy suitably 60 weight percent palladium, 40 weight percent nickel, is placed between the adaptor ring 54 and the step collar 46 , and additional braze alloy is provided if necessary as a wire at the end of the joint 56 .
- the assembly is heated in vacuum to a brazing temperature where the braze alloy melts; above about 1238° C. in the case of the preferred braze alloy, and then cooled to room temperature.
- the injector 38 is provided, numeral 76 .
- the combustion chamber 22 is provided, numeral 78 .
- the metallic deposit 50 is deposited in the region 52 by any operable method, suitably on the outer surface 30 , numeral 80 .
- a suitable columbium deposit 50 is deposited onto the outer surface of the wall 24 of the combustion chamber 22 by inertial welding of a deposit, explosive cladding of a deposit, chemical vapor deposition, or plasma spray into the masked-off region 52 . The result is an excellent metallurgical bond between the wall 24 and the metallic deposit 50 .
- the brazed assembly of the adaptor ring 54 and the step collar 46 is welded at its injector end 60 to the injector 38 , and welded at its chamber end 64 to the metallic deposit 50 , numeral 82 .
- Any operable welding technique may be used.
- both welds are electron beam welds.
- the two welds may be made in any order.
- the columbium adaptor ring 54 readily welds both to the columbium deposit 50 and to the titanium injector 38 .
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Pressure Welding/Diffusion-Bonding (AREA)
- Welding Or Cutting Using Electron Beams (AREA)
- Coating By Spraying Or Casting (AREA)
Abstract
A rocket engine has a combustion chamber, an injector, and an attachment between the combustion chamber and the injector. The attachment includes an annular metallic deposit joined to the chamber wall outer surface, and an annular transition ring structure. The transition ring structure has an annular step collar, and an annular adaptor ring brazed to the annular step collar. The adaptor ring is welded on one end to the injector and on the other end to the metallic deposit.
Description
- This application is a continuation of U.S. application Ser. No. 09/144,375, filed on Aug. 31, 1998, which in turn claims the benefit of U.S. Provisional Application No. 60/057,594, filed Aug. 29, 1997.
- This invention relates to the structure of a liquid-fueled rocket engine, and, more particularly, to the joining of the propellant injector and the combustion chamber.
- A typical liquid-fueled rocket engine includes a generally cylindrical combustion chamber, with an injector attached to its injector end and a flared nozzle attached to its nozzle end. A liquid propellant including fuel and an oxidizer flows through injector ports in the injector and into the combustion chamber. The propellant is ignited in the combustion chamber. The hot gas resulting from the combustion expands through the nozzle and drives the rocket engine and the attached rocket structure in the direction opposite to that in which the nozzle is pointed.
- The wall of the combustion chamber is exposed to high temperature combustion gas during service. The wall is typically made of a refractory material such as rhenium coated with iridium on the inwardly facing surface. The injector plate is much cooler, and is typically made of titanium. When the rocket engine is fired, there is a large temperature increase from room temperature and a large temperature gradient between the upper end of the combustion chamber and the adjacent injector, through the region where the two are attached.
- Recent advances in rocket engine design to allow higher-temperature combustion and the use of more powerful propellants have resulted in even greater temperatures and temperature gradients. The existing attachment structures for attaching the injector to the combustion chamber are insufficient for operation in this environment. There is therefore a need for an improved approach to the attachment of the injector to the combustion chamber. The present invention fulfills this need, and further provides related advantages.
- The present invention provides a rocket engine with an improved attachment between the injector and the combustion chamber. The rocket engine may be operated reliably at higher temperatures than possible with prior attachment procedures. The fabrication approach uses separately known techniques in a new way.
- In accordance with the invention, a rocket engine comprises a combustion chamber comprising an annular wall, an injector, and an attachment between the combustion chamber and the injector. The attachment comprises an annular metallic deposit joined to a first region of the wall of the combustion chamber, an annular transition ring structure, a first weldment between a first welded portion of the transition ring structure and the metallic deposit, and a second weldment between a second welded portion of the transition ring structure and the injector. The first region of the wall of the combustion chamber, to which the metallic deposit is joined, is typically an outer surface of the chamber wall.
- The transition ring structure suitably, but not necessarily, includes an annular step collar, an annular adaptor ring, and a braze joint between at least a portion of the step collar and a portion of the adaptor ring. The step collar protects the end of the combustion chamber inner wall adjacent the injector from damage by the combustion gas and also improves the mixing of the propellants after injection. The step collar is brazed to the adaptor ring, which in turn is welded to the injector and to the metallic deposit.
- The metallic deposit may be deposited by any operable technique. Suitably, it is deposited by chemical vapor deposition. The result is a good bond between the metallic deposit and the wall of the combustion chamber. Such a good bond is not easily attained for refractory metals such as the rhenium used in high-performance thrust chamber walls. In this case, the metallic deposit is suitably columbium. (As used herein, a metal identified generically includes both the pure metal and its alloys containing at least about 50 percent by weight of the pure metal. Thus, for example, in the Specification and in the claims, “columbium” includes both pure columbium and its alloys.)
- The injector is typically titanium (including both pure titanium and its alloys), and the adaptor ring is typically columbium (including both pure columbium and its alloys). The titanium/columbium weld between the injector and the adaptor ring, and the columbium/columbium weld between the adaptor ring and the metallic deposit, are both readily accomplished by electron beam welding.
- The present approach therefore allows the fabrication of a rocket engine from difficult-to-join materials. The joints are sound and gas tight, both at low temperatures and at the elevated temperatures achieved during service. The integrity of the joints is not lost upon the rapid heating of the joint and under imposed high thermal gradients.
- Other features and advantages of the present invention will be apparent from the following detailed description of the preferred embodiment, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. The scope of the invention is not, however, limited to the preferred embodiment.
- The foregoing aspects and many of the attendant advantages of this invention will become more readily appreciated as the same become better understood by reference to the following detailed description, when taken in conjunction with the accompanying drawings, wherein:
- FIG. 1 is a sectional view of a rocket engine;
- FIG. 2 is an enlarged detail of FIG. 1 in region2-2, showing the injector, the injector end of the combustion chamber, and the attachment therebetween; and
- FIG. 3 is a block flow diagram of a method for fabricating the rocket engine.
- While the preferred embodiment of the invention has been illustrated and described, it will be appreciated that various changes can be made therein without departing from the spirit and scope of the invention.
- FIG. 1 depicts a
rocket engine 20, and FIG. 2 shows a detail of the rocket engine. Therocket engine 20 includes acombustion chamber 22 having two major parts, a generally cylindricalannular wall 24 having acylindrical axis 26, and anexpansion nozzle 28. A narrowedthroat 29 separates the cylindricalannular wall 24 from theexpansion nozzle 28. In the illustratedrocket engine 20, thewall 24, thethroat 29, and thenozzle 28 are fabricated separately and joined together, but they may instead be formed integrally. Thecylindrical wall 24 has an outercylindrical surface 30 and an innercylindrical surface 32. Thecombustion chamber 22 has aninjector end 34 and anozzle end 36. - An
injector 38 is attached to theinjector end 34 of thecombustion chamber 22. Theinjector 38 may be of any operable design, but is typically a plate having a plurality of injector ports 40 therein, of which only two are shown in FIGS. 1 and 2. The propellant is provided to thecombustion chamber 22 through the injector ports 40. Some of injector ports 40 a are supplied with a fuel through a fuel valve 42, and others of the injector ports 40 b are supplied with an oxidizer through anoxidizer valve 44. The fuel and oxidizer flow through their respective injector ports 40 into the interior of thecombustion chamber 22 and mix together. Where the mixture is hypergolic, as in the case of monomethylhydrazine (fuel) and nitrogen tetroxide/3 percent nitric oxide (oxidizer), the mixture ignites spontaneously. In other cases where the mixture does not spontaneously ignite, an ignitor (not shown) would be provided. The gaseous combustion products of the combustion expand rearwardly and outwardly through thenozzle 28 and drive therocket engine 20, and the spacecraft to which it is attached, in the opposite direction. - To aid in the thorough mixing and combustion of the propellant fuel and oxidizer, a generally cylindrical
annular step collar 46 is optionally but suitably fitted within thecombustion chamber 22 at and adjacent to theinjector end 34 thereof. Thestep collar 46 also protects the innercylindrical surface 32 from contact with the hot combustion gas when the rocket engine is fired, along the upper end of the combustion chamber, closest to theinjector 38, where the greatest corrosion and erosion damage from the combustion product is expected. The step in thestep collar 46 encourages downstream turbulent mixing of the injected fuel and oxidizer to achieve maximum burning efficiency and specific impulse of the engine. When viewed in circumferential section as in FIG. 2, thestep collar 46, which is suitably made of a single piece of material, has a generally “L” shape, with along leg 47 of the “L” lying parallel to thecylindrical axis 26 and ashort leg 49 of the “L” lying perpendicular to thecylindrical axis 26. Theshort leg 49 of the “L” extends further radially outwardly from the location where it meets thelong leg 47. Suitably, thestep collar 46 is not joined directly to thewall 24 because in such a structure the outward thermal expansion of thestep collar 46 during service deforms thewall 24 and can lead to its failure. Instead, there is suitably asmall gap 51 between thestep collar 46 and the wall. Thegap 51 is sufficiently large that, when the engine is operated and the step collar and wall heat and expand, the outer surface of the step collar does not contact the inner surface of the wall. Thegap 51 is suitably no larger than required to prevent such contact, because a larger gap would allow hot combustion gas to flow into the gap by a backdraft effect. - An attachment48 joins the
combustion chamber 22, thestep collar 46, and theinjector 38. The attachment 48 includes several elements and several joints. The attachment 48 must mechanically join thecombustion chamber 22, thestep collar 46, and theinjector 38 to bear the loads imposed during handling and service, during the large temperature changes and gradients which are experienced during service, and also provide a seal against the leakage of hot gas at theinjector end 34 of thecombustion chamber 22. - The attachment48 includes an annular metallic deposit 50 joined in a metal-to-metal contact to a
first region 52 of thewall 24 of thecombustion chamber 22. Thefirst region 52 is suitably on the outercylindrical surface 30 of thewall 24, at theinjector end 34. The metallic deposit 50 is suitably about 0.110 thick (in the radial direction perpendicular to the cylindrical axis 26) at its thickest location and tapers toward thewall 24 with increasing distance from theinjector 38. The thick region is about 0.1 inches long in the direction parallel to thecylindrical axis 26, and the tapered region extends about another 0.1 inch, for a total length of the deposit 50 of about 0.2 inch. - The attachment48 further includes an
annular adaptor ring 54 joined to the step collar 46 (if present), suitably by a joint 56, which is most suitably a brazed joint. - The
adaptor ring 54 is joined by a joint 58 at an injector end 60 to theinjector 38, and is joined by a joint 62 at a chamber end 64 to the annular metallic deposit 50. Thejoints step collar 46. - In one embodiment, the
wall 24 of thecombustion chamber 22 is made of rhenium with a 0.003-0.005 inch thick coating of iridium on the innercylindrical surface 32. (As used herein, a metal identified generically includes both the pure metal and its alloys containing at least about 50 percent by weight of the pure metal. Thus, for example, “rhenium” includes both pure rhenium and its alloys.) The metallic deposit is suitably made of relatively pure columbium or C 103 alloy having a composition of 10 weight percent hafnium, 1 weight percent titanium, balance columbium. Thestep collar 46 is suitably made of a material having a high melting point and good corrosion/erosion resistance in the combustion environment of thecombustion chamber 22, such as an alloy of platinum and rhodium, an alloy of columbium, or a ceramic. A suitable alloy of platinum and rhodium is 90 percent by weight platinum, balance rhodium. A suitable alloy of columbium is 10 percent by weight hafnium, 1 percent by weight titanium, balance columbium. A suitable ceramic is aluminum oxide, thorium oxide, or yttria-stabilized zirconium oxide. Theadaptor ring 54 is suitably made of columbium, most suitably C 103 alloy having a composition of 10 weight percent hafnium, 1 weight percent titanium, balance columbium. Theinjector 38 is suitably made of titanium alloy of 6 weight percent aluminum, 4 weight percent vanadium, balance titanium. The joint 56 is suitably brazed with a braze alloy of 60 weight percent palladium, 40 weight percent nickel. This braze alloy is compatible with corrosive combustion gas products, does not substantially diffuse into the structure on either side of the joint, and does not leak propellant gas. Thejoints - The metallic deposit50 is present to join the
adaptor ring 54 to thewall 24 of thecombustion chamber 22. Thecolumbium adaptor ring 54 does not readily braze or weld directly to the relativelythin rhenium wall 24 of thecombustion chamber 22 with the required strength and soundness. The metallic deposit 50 is first deposited onto thewall 24, and then theadaptor ring 54 is welded to the deposit 50. - FIG. 3 illustrates one approach to the fabrication of the
rocket engine 20. Theadaptor ring 54 is provided, numeral 70, and thestep collar 46 is provided, numeral 72. Theadaptor ring 54 and thestep collar 46 are brazed together, numeral 74. In brazing, a shim of the braze alloy, suitably 60 weight percent palladium, 40 weight percent nickel, is placed between theadaptor ring 54 and thestep collar 46, and additional braze alloy is provided if necessary as a wire at the end of the joint 56. The assembly is heated in vacuum to a brazing temperature where the braze alloy melts; above about 1238° C. in the case of the preferred braze alloy, and then cooled to room temperature. - The
injector 38 is provided, numeral 76. Thecombustion chamber 22 is provided, numeral 78. The metallic deposit 50 is deposited in theregion 52 by any operable method, suitably on theouter surface 30,numeral 80. A suitable columbium deposit 50 is deposited onto the outer surface of thewall 24 of thecombustion chamber 22 by inertial welding of a deposit, explosive cladding of a deposit, chemical vapor deposition, or plasma spray into the masked-off region 52. The result is an excellent metallurgical bond between thewall 24 and the metallic deposit 50. - The brazed assembly of the
adaptor ring 54 and thestep collar 46 is welded at its injector end 60 to theinjector 38, and welded at its chamber end 64 to the metallic deposit 50,numeral 82. Any operable welding technique may be used. Suitably, both welds are electron beam welds. The two welds may be made in any order. In one combination of materials, thecolumbium adaptor ring 54 readily welds both to the columbium deposit 50 and to thetitanium injector 38. - Although a particular embodiment of the invention has been described in detail for purposes of illustration, various modifications and enhancements may be made without departing from the spirit and scope of the invention. Accordingly, the invention is not to be limited except as by the appended claims.
Claims (14)
1. A rocket engine, comprising
a combustion chamber comprising an annular wall with a chamber wall outer surface and a combustion chamber axis;
an injector; and
an attachment between the combustion chamber and the injector, the attachment comprising:
an annular metallic deposit joined to a first region of the wall of the combustion chamber,
an annular transition ring structure, wherein the transition ring structure comprises:
an annular step collar,
an annular adaptor ring, and
a braze joint between at least a portion of the step collar and a portion of the adaptor ring.
a first joint between the adaptor ring and the metallic deposit, and
a second joint between the adaptor ring and the injector.
2. The rocket engine of claim 1 , wherein the first region of the wall of the combustion chamber is a portion of the chamber wall outer surface.
3. The rocket engine of claim 1 , wherein the second joint lies substantially perpendicular to the combustion chamber axis.
4. The rocket engine of claim 1 , wherein the step collar is made of platinum.
5. The rocket engine of claim 1 , wherein
the wall of the combustion chamber is made of rhenium,
the metallic deposit is made of columbium,
the first joint portion of the transition ring structure is made of columbium,
the second joint portion of the transition ring structure is made of columbium, and
the injector is made of titanium.
6. The rocket engine of claim 1 , wherein the first joint and the second joint are each weldments.
7. A rocket engine, comprising
a combustion chamber comprising a generally cylindrical annular wall with a chamber wall outer surface, a chamber wall inner surface, and a cylindrical axis;
an injector; and
an attachment between the combustion chamber and the injector, the attachment comprising:
an annular metallic deposit joined to the chamber wall outer surface,
an annular transition ring structure including
an annular step collar,
an annular adaptor ring, and
a braze joint between at least a portion of the step collar and a portion of the adaptor ring.
a first weldment between a first welded portion of the adaptor ring structure and the metallic deposit, and
a second weldment between a second welded portion of the adaptor ring structure and the injector.
8. The rocket engine of claim 7 , wherein
the wall of the combustion chamber is made of rhenium,
the metallic deposit is made of columbium,
the first welded portion of the transition ring structure is made of columbium.
the second welded portion of the transition ring structure is made of columbium, and
the injector is made of titanium.
9. A method for fabricating a rocket engine, comprising the steps of providing a combustion chamber comprising an annular wall with a chamber wall outer surface and a combustion chamber axis;
depositing an annular metallic deposit on a first region of the wall of the combustion chamber;
providing an attachment between the combustion chamber and the injector,
the attachment comprising an annular transition ring structure comprising an annular adaptor ring and an annular step collar;
welding the attachment to the metallic deposit;
providing an injector; and
welding the injector to the attachment.
10. The method of claim 9 , wherein the step of depositing includes the step of
depositing the annular metallic deposit by a technique selected from the group consisting of inertial welding, explosive cladding, chemical vapor deposition, and plasma spraying.
11. The method of claim 9 , wherein the step of depositing includes the step of
depositing the metallic deposit onto the chamber wall outer surface.
12. The method of claim 9 , wherein the step of welding the attachment to the metallic deposit includes the step of
forming a weld joint lying substantially perpendicular to the combustion chamber axis.
13. The method of claim 9 wherein the step of providing an attachment includes the steps of
providing an annular step collar,
providing an annular adaptor ring, and
brazing together at least a portion of the step collar and a portion of the adaptor ring.
14. The method of claim 9 , wherein
the wall of the combustion chamber is made of rhenium,
the metallic deposit is made of columbium,
the first welded portion of the transition ring structure is columbium.
the second welded portion of the transition ring structure is columbium, and the injector is titanium.
Priority Applications (1)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US09/772,196 US20020014070A1 (en) | 1997-08-29 | 2001-01-30 | Rocket engine having a transition attachment between a combustion chamber and an injector |
Applications Claiming Priority (3)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
US5759497P | 1997-08-29 | 1997-08-29 | |
US09/144,375 US6381949B1 (en) | 1997-08-29 | 1998-08-31 | Rocket engine having a transition attachment between a combustion chamber and an injector |
US09/772,196 US20020014070A1 (en) | 1997-08-29 | 2001-01-30 | Rocket engine having a transition attachment between a combustion chamber and an injector |
Related Parent Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US09/144,375 Continuation US6381949B1 (en) | 1997-08-29 | 1998-08-31 | Rocket engine having a transition attachment between a combustion chamber and an injector |
Publications (1)
Publication Number | Publication Date |
---|---|
US20020014070A1 true US20020014070A1 (en) | 2002-02-07 |
Family
ID=22011575
Family Applications (2)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US09/144,375 Expired - Lifetime US6381949B1 (en) | 1997-08-29 | 1998-08-31 | Rocket engine having a transition attachment between a combustion chamber and an injector |
US09/772,196 Abandoned US20020014070A1 (en) | 1997-08-29 | 2001-01-30 | Rocket engine having a transition attachment between a combustion chamber and an injector |
Family Applications Before (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
US09/144,375 Expired - Lifetime US6381949B1 (en) | 1997-08-29 | 1998-08-31 | Rocket engine having a transition attachment between a combustion chamber and an injector |
Country Status (4)
Country | Link |
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US (2) | US6381949B1 (en) |
EP (1) | EP0899449B1 (en) |
JP (1) | JP3051377B2 (en) |
DE (1) | DE69812014T2 (en) |
Cited By (3)
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US20080264372A1 (en) * | 2007-03-19 | 2008-10-30 | Sisk David B | Two-stage ignition system |
US20130305687A1 (en) * | 2012-05-17 | 2013-11-21 | Ihi Aerospace Co., Ltd. | Thruster and Spacecraft |
CN115073200A (en) * | 2022-05-19 | 2022-09-20 | 北京控制工程研究所 | Butt-joint sealing structure and method for ceramic reaction chamber and high-temperature alloy injector |
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DE19927734C2 (en) * | 1999-06-17 | 2002-04-11 | Astrium Gmbh | Thrust chamber of a propulsion engine for satellites and transport devices for space applications |
US7256815B2 (en) * | 2001-12-20 | 2007-08-14 | Ricoh Company, Ltd. | Image forming method, image forming apparatus, optical scan device, and image forming apparatus using the same |
US6918243B2 (en) * | 2003-05-19 | 2005-07-19 | The Boeing Company | Bi-propellant injector with flame-holding zone igniter |
US7400697B1 (en) * | 2003-12-08 | 2008-07-15 | Bwx Technologies, Inc. | Clad tube for nuclear fuel |
US7565795B1 (en) * | 2006-01-17 | 2009-07-28 | Pratt & Whitney Rocketdyne, Inc. | Piezo-resonance igniter and ignition method for propellant liquid rocket engine |
US8122703B2 (en) | 2006-04-28 | 2012-02-28 | United Technologies Corporation | Coaxial ignition assembly |
WO2007139452A1 (en) * | 2006-05-29 | 2007-12-06 | Volvo Aero Corporation | Method for fastening of a heater |
US20080299504A1 (en) * | 2007-06-01 | 2008-12-04 | Mark David Horn | Resonance driven glow plug torch igniter and ignition method |
US20090266870A1 (en) * | 2008-04-23 | 2009-10-29 | The Boeing Company | Joined composite structures with a graded coefficient of thermal expansion for extreme environment applications |
US8512808B2 (en) | 2008-04-28 | 2013-08-20 | The Boeing Company | Built-up composite structures with a graded coefficient of thermal expansion for extreme environment applications |
US8814562B2 (en) * | 2008-06-02 | 2014-08-26 | Aerojet Rocketdyne Of De, Inc. | Igniter/thruster with catalytic decomposition chamber |
US9404441B2 (en) * | 2008-08-18 | 2016-08-02 | Aerojet Rocketdyne Of De, Inc. | Low velocity injector manifold for hypergolic rocket engine |
US8161725B2 (en) | 2008-09-22 | 2012-04-24 | Pratt & Whitney Rocketdyne, Inc. | Compact cyclone combustion torch igniter |
TWI422741B (en) * | 2010-02-24 | 2014-01-11 | Nat Applied Res Laboratories | Motor |
RU2532640C2 (en) * | 2010-11-17 | 2014-11-10 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт машиностроения" (ФГУП "НИИМаш") | Low-thrust liquid propellant rocket engine chamber |
US8505577B2 (en) * | 2010-11-30 | 2013-08-13 | The United States Of America As Represented By The Secretary Of The Army | Pnumatically actuated bi-propellant valve (PABV) system for a throttling vortex engine |
TWI504538B (en) * | 2013-05-31 | 2015-10-21 | Nat Applied Res Laboratories | Dual-vortical-flow hybrid rocket engine |
RU2605267C2 (en) * | 2015-04-29 | 2016-12-20 | Федеральное государственное унитарное предприятие "Научно-исследовательский институт машиностроения" (ФГУП "НИИМаш") | Low-thrust rocket engines unit |
EP3717768B1 (en) * | 2017-12-02 | 2022-12-14 | Aerojet Rocketdyne, Inc. | Copper alloy combustion chamber attached to injector by non-copper weld transition ring |
US11779985B1 (en) * | 2020-11-15 | 2023-10-10 | Herbert U. Fluhler | Fabricating method for low cost liquid fueled rocket engines |
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US3903693A (en) * | 1973-03-26 | 1975-09-09 | Anthony Fox | Rocket motor housing |
US4785748A (en) | 1987-08-24 | 1988-11-22 | The Marquardt Company | Method sudden expansion (SUE) incinerator for destroying hazardous materials & wastes |
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US4936091A (en) * | 1988-03-24 | 1990-06-26 | Aerojet General Corporation | Two stage rocket combustor |
US4882904A (en) * | 1988-03-24 | 1989-11-28 | Aerojet-General Corporation | Two stage rocket combustor |
DE69820173T2 (en) * | 1997-08-29 | 2004-10-14 | Hughes Electronics Corp., El Segundo | Connection ring for a rocket combustion chamber |
EP0899448B1 (en) * | 1997-08-29 | 2003-03-19 | Hughes Electronics Corporation | Fabrication of a rocket engine with transition structure between the combustion chamber and the injector |
US6138451A (en) * | 1998-05-11 | 2000-10-31 | Hughes Electronics Corporation | Rocket engine with combustion chamber step structure insert, and its fabrication |
US6138450A (en) * | 1998-05-11 | 2000-10-31 | Hughes Electronics Corporation | Rocket engine with integral combustion chamber step structure and its fabrication |
-
1998
- 1998-08-27 EP EP98116170A patent/EP0899449B1/en not_active Expired - Lifetime
- 1998-08-27 DE DE69812014T patent/DE69812014T2/en not_active Expired - Lifetime
- 1998-08-31 US US09/144,375 patent/US6381949B1/en not_active Expired - Lifetime
- 1998-08-31 JP JP10244653A patent/JP3051377B2/en not_active Expired - Lifetime
-
2001
- 2001-01-30 US US09/772,196 patent/US20020014070A1/en not_active Abandoned
Cited By (4)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US20080264372A1 (en) * | 2007-03-19 | 2008-10-30 | Sisk David B | Two-stage ignition system |
US20130305687A1 (en) * | 2012-05-17 | 2013-11-21 | Ihi Aerospace Co., Ltd. | Thruster and Spacecraft |
US9376987B2 (en) * | 2012-05-17 | 2016-06-28 | Ihi Aerospace Co., Ltd. | Thruster and spacecraft |
CN115073200A (en) * | 2022-05-19 | 2022-09-20 | 北京控制工程研究所 | Butt-joint sealing structure and method for ceramic reaction chamber and high-temperature alloy injector |
Also Published As
Publication number | Publication date |
---|---|
DE69812014D1 (en) | 2003-04-17 |
JP3051377B2 (en) | 2000-06-12 |
EP0899449B1 (en) | 2003-03-12 |
US6381949B1 (en) | 2002-05-07 |
JPH11132106A (en) | 1999-05-18 |
DE69812014T2 (en) | 2003-12-24 |
EP0899449A2 (en) | 1999-03-03 |
EP0899449A3 (en) | 2000-06-14 |
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STCB | Information on status: application discontinuation |
Free format text: ABANDONED -- FAILURE TO RESPOND TO AN OFFICE ACTION |
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Owner name: AEROJET-GENERAL CORPORATION, WASHINGTON Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:GENERAL DYNAMICS OTS (AEROSPACE), INC.;REEL/FRAME:013552/0854 Effective date: 20021002 |