Power plants – Reaction motor – Solid propellant
Reexamination Certificate
2001-03-29
2002-12-31
Gartenberg, Ehud (Department: 2859)
Power plants
Reaction motor
Solid propellant
C060S257000, C060S803000, C374S144000, C374S148000
Reexamination Certificate
active
06499289
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates to a gauge for measuring heat fluxes, especially in high temperature, highly corrosive environments. In particular, the present invention relates to a gauge for measuring heat fluxes in rocket motor nozzles, including sub-scale and full-scale rocket motor nozzles, heat shields of re-entry vehicles, rocket motor combustion chambers, and metal processing. This invention also relates to a method of measuring heat fluxes with the gauge in various high temperature, corrosive environments.
2. Description of the Related Art
Solid rocket motors typically include an outer case or shell housing a solid propellant grain which, in the case of a hybrid motor, is a solid fuel or oxidizer grain. The rocket motor case is conventionally manufactured from a rigid, yet durable, material such as steel or filament-wound composite. The propellant is housed within the case. There are several basic propellant grain configurations known in the art (and compatible with the use of this invention) for loading propellant within the case. The two most commonly used configurations are the center-perforated grain configuration and the end burning grain configuration. In the center-perforated grain configuration, the flame front advances radially from the center perforation towards the outer case. On the other hand, in the end-burning grain configuration the flame front advances axially from the nozzle end of the motor towards the forward dome.
During firing, oxidizing agents of the propellant serve to drive combustion reactions in a combustion chamber to form large quantities of combustion products, which are expelled from the rocket motor through a nozzle in fluid communication with the combustion chamber. 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 restricted throat region and aft walls diverging from the throat region to a larger exit area, thus defining a converging/diverging contoured pathway. 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.
During operation, combustion of solid rocket propellant generates extreme conditions within the case and along the contoured nozzle pathway of the rocket motor. For example, temperatures inside the rocket motor case can exceed 2760° C. (5,000° F.), and interior pressures can exceed 1,500 psi. These factors combine to create a high degree of turbulence for particles entrained in the combustion gases.
A heat insulating layer (insulation) protects the rocket motor case from the hot gas and highly erosive particle streams generated by the combusting propellant. Typically, the propellant grain is bonded to the insulation and/or non-insulated portions of the case by use of a lining layer (liner). In addition to its adhesive function of bonding the propellant to the insulation and any non-insulated portions of the case, the liner also supplements the insulator by functioning to inhibit the burning surface of the propellant grain when the propellant/liner interface is exposed to an approaching flame front. Additionally, the liner isolates propellant within the case from the environment and prevents leakage of combustion gases or liquid into or through the case.
Likewise, the contoured nozzle pathway, including the restricted nozzle region, must also be insulated to withstand the elevated temperatures and pressures of the combustion products, as well as the erosive effects of turbulent particles entrained within the combustion gas. Carbon-based and silica-based ablative materials are highly advantageous for use as nozzle insulation due to their excellent ablative properties, low cost, and relatively light weight.
It is widely acknowledged in the industry, however, that carbon-and silica-based insulation and ablative materials, such as those present at case insulation-to-propellant interfaces and those defining nozzle contour pathways, are highly susceptible to recession at high operating temperatures. Convective and radiative heating of the ablative materials by the combustion products increases the vulnerability of the ablative materials to recession. As incident heat is conducted to the nozzle throat ablative material, the ablative material tends to decompose into pyrolysis gases and residual carbon, which are carried away with the propellant combustion gases. If not accounted for, the recession of the nozzle throat inner surface during motor operation may become a source of several problems in rocket operation, including decreased efficiency and loss of predictability.
Thus, the design of rocket insulation, including the selection of an appropriate ablative materials and insulation thickness, is dependent upon the convective and radiative heat fluxes incident on ablative surfaces at various locals in rocket motors and nozzles. Because no reliable measuring technique has heretofore been known for determining actual heat fluxes, heat fluxes are usually estimated by modeling based on data obtained from such techniques as computational fluid dynamics.
The criticality of predication accuracy is demonstrated by the severity and magnitude of the risk of failure due to erosion. Most insulation is, of necessity, “man-rated” in the sense that a catastrophic failure can result in the loss of human life—whether the rocket motor is used as a booster for launch of a manned rocket or is carried tactically underneath the wing of an attack aircraft. The monetary cost of failure in satellite launches is well-publicized and can run into the hundreds of millions of dollars. Additionally, as mentioned above, unforeseen amounts of recession in the nozzle throat insulation can significantly affect the flow expansion contour in the nozzle, affecting motor performance.
For these reasons, there is a strong desire in the art to validate heat flux predictions through actual measurement of the convective and radiative heat fluxes incident on the ablative surfaces of rocket motors and nozzles. As described in Wool et al., “Measurement of Convective and Radiative Heat Fluxes at the Surface of an Ablative Material” (1970), thermocouples have been proposed for validating heat flux predictions. Wool et al. state that thermocouples used for this purpose should be capable of accurately measuring the high temperatures encountered in rocket motor materials and, because ablative material recedes and decomposes, should be capable of obtaining sufficient data for a continuous evaluation of the heat flux. Additionally, the gauge must be capable of withstanding the high operating temperatures to which it is exposed for a sufficient time to obtain heat flux data. Although the gauges disclosed in Wool et al. reportedly attain these objects, the structure of the Wool et al. thermocouple devices requires that data reduction be done by a complex analytical technique in which assumptions are made concerning phenomena such as surface thermochemical reactions, transient heat conduction, and in-depth pyrolysis gas generation.
SUMMARY OF THE INVENTION
It is, therefore, an object of this invention to provide a gauge capable of measuring either heat fluxes (convective and radiative fluxes) or radiative fluxes at a rocket motor ablative surface with sufficiently high accuracy and for a sufficiently long period of time to permit validation of known prediction models, without the need for complex analytical data reduction.
In accordance with the principles of this invention, the above and other objects are attained by a novel gauge for measuring heat flux, especially
Bunker Robert C.
Ewing Mark E.
Shipley John L.
Alliant Technologies Inc.
Gartenberg Ehud
Sullivan Law Group
LandOfFree
Pyrolytic graphite gauge for measuring heat flux does not yet have a rating. At this time, there are no reviews or comments for this patent.
If you have personal experience with Pyrolytic graphite gauge for measuring heat flux, we encourage you to share that experience with our LandOfFree.com community. Your opinion is very important and Pyrolytic graphite gauge for measuring heat flux will most certainly appreciate the feedback.
Profile ID: LFUS-PAI-O-2931393