Power plants – Reaction motor – Particular exhaust nozzle feature
Reexamination Certificate
2001-01-19
2004-03-30
Gartenberg, Ehud (Department: 3746)
Power plants
Reaction motor
Particular exhaust nozzle feature
C239S265110
Reexamination Certificate
active
06711901
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to a rocket motor nozzle assembly having a throat insert that includes a vacuum plasma-sprayed refractory metallic component for preventing the throat from receding, and an ablative component constructed and arranged to reduce erosion under the refractory metallic component. This invention also relates to a rocket motor engine assembly comprising the nozzle, and to methods of making the rocket motor nozzle assembly and engine assembly.
2. Description of the Related Art
Conventionally, solid rocket motor propellants contain, among other things, oxidizer and fuel components immobilized within a polymeric binder. The solid propellant is cast as a grain within a rigid outer casing or shell of the rocket motor combustion chamber. A heat insulating layer and a liner are usually interposed between the grain and the outer casing to protect the outer casing from the high operating temperatures associated with rocket motor operation and to provide enhanced grain-to-case bonding. The solid-propellant grain is typically configured as either a center perforated grain or an end burner grain.
During firing, the oxidizing agents immobilized within the solid-propellant grain drive combustion reactions to form large quantities of combustion products, which are expelled from the rocket motor through a nozzle in fluid communication with the combustion chamber. The amount of thrust produced by a rocket motor is proportional to the exit gas velocity squared. Nozzles are designed to accelerate the combustion product gases from the propellant grain to the maximum velocity at the exit.
To achieve this end, nozzles usually have forward walls converging to a throat region and have aft walls diverging from the throat region to a larger exit area to form a converging/-diverging nozzle configuration. The nature of compressible gases is such that a converging/-diverging nozzle increases the exit gas velocity and thereby thrust. The proportions of the mass flow pathway, particularly the ratio of area at the exit plane to area at the throat, establish how efficiently the nozzle converts pressure in the mass flow stream to thrust produced by the motor. It is within the purview of those skilled in the art to design a nozzle throat to optimize the ratio of exit area to throat area.
As described above, an insulating layer and liner are typically placed between the propellant and the outer casing of the combustion chamber to protect the outer casing from the extremely high temperatures at which the rocket motor operates. Likewise, the nozzle throat must also be designed to withstand the elevated temperatures and pressures, reactions with the combustion products, and the high velocities at which the combustion products pass over the nozzle inner surface.
Carbon-based and silica-based materials are highly advantageous for use as nozzle insulation due to their excellent ablative properties, inexpensive cost, and relatively low weight. Intuitively, a lesser weight nozzle assembly is desirable because a lesser weight nozzle assembly imparts a lesser weight penalty to the rocket motor than a heavier nozzle assembly, thereby increasing the distance the rocket motor assembly can travel. As referred to herein, carbon-based and silica-based materials include, but are not limited to, carbon, silica, or graphite bulk and composite materials with constituents previously subject to carbonization or graphitization, known as carbon/carbon, graphite/carbon, and cloth, fiber, or powder-filled phenolic composites, and also a large array of metal or silicon carbides.
It is widely acknowledged in the industry, however, that carbon- and silica-based nozzle throats tend to recede, especially at high operating temperatures and pressures. The material loss in the nozzle throat is generally attributed to one or more mechanisms. For example, in the case of silica-cloth phenolics and the like, material loss is attributed to thermal decomposition and melting of the liner material, which can cause the decomposed layer to be sheared off by the high-velocity gas stream. In the case of carbon cloth phenolics and the like, material loss can result from the thermal decomposition of the liner material, which forms pyrolysis gas and char due to chemical reactions of decomposed material (char) with the combustion gases. Typical carbon-carbon materials and the like may undergo chemical reaction with the reactive species in the combustion products.
The recession of the nozzle throat inner surface during motor operation is a source of several problems in rocket operation. As the nozzle throat material recedes, the exit area to throat area ratio (or expansion ratio) diminishes, thereby decreasing the efficiency of the nozzle and causing loss in motor performance. Additionally, rough nozzle surfaces, which tend to form during nozzle recession, have been shown to undergo recession at faster rates than smooth surfaces. Thus, the nozzle throat recession process can be characterized as a self-perpetuating phenomenon. Another problem is having higher than predicted nozzle recession. Calculations for determining payload weight and corresponding motor designs must be accurate to ensure that the payload will reach its intended target. The calculations necessary for ascertaining rocket dimensions and payloads are dependent upon many variables, including nozzle throat diameter. In-flight variations of nozzle throat diameter due to recession can reduce motor performance.
To address the shortcomings of carbon- and silica-based nozzle throats, refractory metal and metal alloys are occasionally used in rocket motors, especially as the throat insert. Examples of such refractory materials are tungsten and its alloys.
However, the weight penalty and expense associated with the presence of the tungsten and other refractory metals often make these refractory materials impractical and uneconomical for applications involving bulky throat insert cross sections. Additionally, refractory metals exposed to hot combustion exhaust gases and particles are subject to tensile stresses due to thermal shock early in motor burn when thermal expansions near the rapidly heated exposed surfaces are restrained by cooler regions of the cross section farther from the exposed surfaces. Indeed, surface heating can be so intense that temperature gradients of thousands of degrees per inch are possible. Such thermal stresses in both the axial and tangential (or hoop) directions can produce thermal fractures in the nozzle component, and potential ejection of the throat insert from the motor.
It would, therefore, be a significant advancement in the art to provide a nozzle assembly that takes advantage of the low weight of carbon- and/or silica-based materials and the erosion resistance of metals and alloys, yet avoids significant nozzle recession experienced by carbon- and silica-based materials and reduces the risks associated with thermal stresses encountered by thick refractory materials, even when the operating conditions are characterized by high temperatures and pressures.
SUMMARY OF THE INVENTION
This invention realizes the above-discussed advancement in the art. In accordance with the principles of this invention, this and other advantages are attained by the provision of a rocket motor nozzle assembly defining a converging/diverging pathway with a throat of restricted cross-sectional area and operatively engageable with a rocket motor to receive the combustion products of the rocket motor, pass the combustion products through the converging/diverging pathway, and discharge the combustion products to propel and/or divert the rocket motor assembly.
In accordance with an embodiment of this invention, the rocket motor nozzle assembly comprises a throat insert comprising a carbon-based throat support and a vacuum plasma-sprayed metal shell comprising at least one refractory metal and/or refractory metal alloy having a melting temperature above 2700° C., preferably above 2760° C. The assembly further comprises a carbon- o
Canfield Alan R.
Shigley John K.
Alliant Techsystems Inc.
Gartenberg Ehud
TraskBritt
LandOfFree
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