Metal working – Method of mechanical manufacture – Rocket or jet device making
Reexamination Certificate
2000-10-05
2001-10-30
Freay, Charles G. (Department: 3746)
Metal working
Method of mechanical manufacture
Rocket or jet device making
C029S458000, C029S527200, C427S452000, C427S455000, C156S175000, C156S153000, C060S253000
Reexamination Certificate
active
06308408
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to rocket engines. It relates in particular to a combustion chamber
ozzle assembly for a rocket engine and to a process for its fabrication.
2. Description of Related Art
Conventional combustion/chamber nozzle assemblies for rocket engines are usually actively cooled. That is, they generally contain integral cooling passages for cooling fluid within the combustion chamber and nozzle walls, which tubular cooling passages are fed by manifolds. A complex, weighty structure is presented, the fabrication of which requires the construction and assemblage of multiple piece parts through numerous procedural steps, including machining, plating, welding, and brazing. Such a complex, weighty structure, as well as its complicated method of fabrication, are both disadvantageous and in need of improvement, as is well known in this art.
Summary of the Invention
It is accordingly a primary object of the present invention to provide what is lacking in the prior art, especially a simple, yet highly efficient, lightweight integral combustion chamber
ozzle assembly for a rocket engine. It is also a primary object of the present invention to provide an uncomplicated and highly reliable process for the fabrication of a simple, lightweight integral combustion chamber
ozzle assembly for a rocket engine, which process is effective and highly efficient, especially in respect of the utilization of time and materials.
These objects and other related benefits are achieved by the present invention, which in one aspect thereof is an integral, lightweight combustion chamber
ozzle assembly which has a shell of a refractory material, such as an alloy of niobium, having a configuration defining a chamber of generally frusto-conical contour. The chamber communicates at its smaller end with a rocket body, and terminates at its larger end in a cone open at its terminus, which serves as a nozzle for the rocket engine. The inner surface of the chamber has applied thereto a thermal and oxidation barrier layer, especially of a silicide or aluminum oxide. An ablative silica phenolic insert, which is bonded to the thermal and oxidation barrier layer, is configured to provide a chosen inner contour for the combustion chamber.
The ablative silica phenolic insert additionally has a taper or reduction in thickness toward the open terminus of the nozzle.
In another aspect, the present invention is a process for fabricating the integral, lightweight combustion chamber
ozzle assembly, which process is set forth in detail hereinafter.
The integral, combustion chamber
ozzle assembly according to the present invention is simple in design and is much lighter than conventional combustion chamber
ozzle assemblies, making it a highly desirable replacement assembly for these reasons alone. Moreover, the integral combustion chamber
ozzle assembly according to the present invention is not actively cooled, as are the assemblies of the prior art, so that there is no need for cooling passages therein. Fabrication is therefore greatly simplified, and accordingly accelerated, resulting in a highly desirable economy in respect of both resources and time.
During the firing operation of a rocket engine employing the integral, lightweight combustion chamber
ozzle assembly of the present invention, resins boil off from the ablative silica phenolic insert, thereby cooling the inner surface of the insert and leaving behind a layer of char. This layer of char, along with the remaining silica phenolic layer, acts as an insulator and protects the refractory metal shell against overheating. The thickness of the ablative insert is chosen so that the layer of char does not penetrate too deeply during the design life of the combustion chamber
ozzle assembly.
During the firing operation of a rocket engine employing the integral, lightweight combustion chamber
ozzle assembly of the present invention, the temperature inside the combustion chamber
ozzle assembly decreases toward the open end of the nozzle. Therefore the thickness of the ablative silica phenolic insert is fashioned to taper down toward the open terminus of the nozzle, at which terminus the refractory metal can, along with the thermal and oxidation layer applied thereto, survive without any ablative protection.
REFERENCES:
patent: 3597821 (1971-08-01), Emerson
patent: 3871173 (1975-03-01), McKenna
patent: 6025661 (2001-03-01), Ring
Sutton, G.P. and Ross, D.M. “Rocket Propulsion Elements”, John Wiley & Sons, 4th Ed, New-York 1976, p. 276.
Cornelius Charles S.
Myers W. Neill
Freay Charles G.
Gartenberg Ehud
Seemann Jerry L.
The United States of America as represented by the Administrator
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