Metal working – Method of mechanical manufacture – Rocket or jet device making
Reexamination Certificate
1999-04-15
2001-03-27
Cuda, I (Department: 3726)
Metal working
Method of mechanical manufacture
Rocket or jet device making
C029S523000
Reexamination Certificate
active
06205661
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to a rocket thrust chamber for thrusters or bipropellant rocket engines. More particularly, the invention is directed to high temperature thrust chambers for relatively small rocket engines generally of about 5-100 pounds of thrust, which are difficult to cool regeneratively.
BACKGROUND OF THE INVENTION
Rocket engine liquid fueled thrust chambers of the larger thrust type, typically of many hundreds, many thousands or even millions of pounds of thrust, employ regeneratively cooled thrust chambers where pressurized (pumped) propellant is first passed through thrust engine tubing or channels forming the shell or cooling jacket of the chamber, before being injected into the combustion chamber. The cool fuel or oxidizer in a bipropellant system, for example, liquid oxygen at −180° C. or other oxidizer, thus keeps the combustion chamber at a sufficiently low temperature to preserve the structural integrity of the thrust chamber. In the case of smaller thrust thrusters used for intermittent thrust control of a space vehicle or satellite, thrust chambers have employed film cooling. Film cooling employs a protective coating of propellant which is sprayed along the inner surface of the thrust chamber. Evaporation of the film cools the chamber wall. Although film cooling is efficient, it is to be avoided since it lowers the overall specific thrust by using propellant for a purpose other than producing thrust. Regeneratively cooled engines are considered more efficient since coolant is not wasted but, in fact, augments the initial energy at injection by its increased heat content.
Conventional thrusters currently in use have a minimal upper temperature limit of about 2400° F. (1315° C.) and a limited life span of about ten hours. These conventional thrusters, using a hydrazine propellant for example, and a thrust chamber constructed of niobium alloys, necessarily will use about 40% of the fuel for film cooling in order to keep the thrust chamber walls below this temperature. Since the propellant is the major mass item for satellites being put in space, a considerable incentive exists to decrease or obviate the need for film cooling and hence the amount of on-board fuel.
U.S. Pat. No. 3,354,652 discusses the difficulty of regeneratively cooling small liquid propellant engines resulting inter alia in boiling or decompositon of the coolant within the coolant jacket. While it has been suggested to apply high temperature insulation, e.g., metal oxides, to the combustion side of the chamber to reduce the coolant bulk temperature during steady state firing, this can result, upon engine shut down, in additional stored heat in the insulation causing localized heating and decomposition of remaining stagnant propellant. The patent somewhat solves the problem by suggesting a tantalum alloy liner coupled with a stagnant gas or vacuum enclosed space and helical two-way flow coolant channels.
U.S. Pat. No. 3,780,533 discloses the use in regeneratively cooled chambers utilizing cooling channels, of a composite wall including a deposit of electroformed nickel, or a sheet of nickel or of refractory alloys, such as copper-silver or molybdenum-rhenium alloys, brazed to lands in a middle wall component U.S. Pat. No. 3,315,471 shows with respect to thrusters utilizing radioisotope fuel, structural elements of the thruster, namely spaced shells, preferably constructed of tungsten. U.S. Pat. No. 3,723,742 shows the use of noble metals and refractory metals surrounding a radioisotope fuel casing.
U.S. Pat. No. 4,917,968 decribes a thrust chamber structure where a ductile layer of a platinum group metal including iridium is deposited by chemical vapor deposition on a mandrel and a layer of refractory metal deposited thereover also by chemical vapor deposition, with a solid solution of the two metals present between and metallurgically bonded to the two metal layers.
U.S Pat. No. 5,613,299 describes a thrust chamber structure where a layer of a platinum group metal, including iridium, is bonded to the interior of a refractory alloy thrust chamber by pressurizing the exterior of the chamber forcing the chamber to collapse onto the liner, itself supported on a solid mandrel. This present invention differs from the above by having a hollow mandrel so that the interior of the chamber is pressurized and the liner expanded outwards.
SUMMARY OF THE INVENTION
The present invention is directed to a fractory metal one-piece thrust chamber for use in a bipropellant rocket engine employing, for example, hydrazine and nitrogen tetroxide propellants. The chamber typically constructed of refractory alloys such as niobium, tantalum, tantalum (10%)-tungsten, rhenium or rhenium-tungsten alloys, is formed by spin forming, swaging or is machined from bar stock, into which discrete liner sections of platinum group metals (e.g., rhodium, iridium or their alloys) are bonded by diffusion bonding, explosive boding, hot isostatic pressing (HIP), isostatic forging, rapid omnidirectional compaction or by the Ceracon process which utilizes fused silica to apply bonding pressure. A cylindrical first liner section of oxidation resistant material such as rhodium, rhodium alloy, e.g., platinum (80%)-rhodium (20%) alloy or other material selected to provide oxidation protection to the chamber walls, is sized to be bonded to a cylindrical barrel portion of the refractory metal chamber. An essentially conical second liner section of iridium, rhodium, iridium-rhodium alloy or other material selected to provide oxidation protection to the chamber walls, is constructed to be bonded to the downstream end of the barrel portion and to a converging conical portion of the chamber. By scarfed end the second liner section is bonded to a scarfed end of the first liner section. [Alternatively the first and second sections can be fabricated as a single piece.] The second liner section extends from the barrel portion along the inner surface of a first conical portion of the chamber. An essentially conical third liner section of iridium or rhodium or an iridium-rhodium alloy or other material selected to provide oxidation protection to the chamber walls, is constructed to be bonded to a second diverging conical portion of the chamber and also by a scarfed end joined to a scarfed end on an overlapping end of the second liner section. The third liner section forms an engine expansion nozzle with the second conical portion of the outside chamber. High bonding pressure is applied usually at elevated temperature to pressure bond the overlapping liner sections together and to pressure bond the lining sections to the respective thrust chamber portions.
In a preferred fabrication process a refractory alloy thrust chamber is constructed to final size in all aspects except that the chamber is extended in length at each end to allow seal welding of an internal hollow mandrel to the ends of the chamber. The protective liner sections are assembled onto the split, hollow mandrel and placed within the hollow confines of the refractory alloy thrust chamber. The hollow mandrel is preferably constructed of niobium although other materials for example steel can be used. The protective liner sections are assembled ‘on the mandrel to’ confirm fit, particularly at the overlapping scarfed joints. The mandrel is then separated at the split and reassembled within the refractory alloy thrust chamber. The total liner length is designed and constructed to meet the thrust chamber design longitudinal dimension, i.e., without the oversize ends the mandrel however is sized to correspond to the extended chamber length. The ends of the mandrel are electron beam welded to each end of the thrust chamber and the welds helium leak checked. A final weld is then made to join both halves of the hollow mandrel together at the throat region of the mandrel. The entire assembly is again helium leak checked. The completed assembly is then preferably hot isostatically pressured using a high pressure and temperature in an argon atmosphere. The mandr
Cuda I
MacPherson Alan H.
Skjerven Morrill & MacPherson
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