Explosive and thermic compositions or charges – Structure or arrangement of component or product – Solid particles dispersed in solid solution or matrix
Reexamination Certificate
2003-02-03
2004-01-27
Hardee, John (Department: 1751)
Explosive and thermic compositions or charges
Structure or arrangement of component or product
Solid particles dispersed in solid solution or matrix
Reexamination Certificate
active
06682615
ABSTRACT:
FIELD OF THE INVENTION
This invention relates to solid propellant formulations, which exhibit good processing properties, good safety characteristics, desirable combustion properties, good safety characteristics and excellent mechanical attributes. More specifically, this invention relates to solid propellant formulations, which have a burning rate of 0.5 in/s at 1000 psi and show no pressure slope break up to 8000 psi.
BACKGROUND OF THE INVENTION
Solid propellants are used extensively in the aerospace industry. Solid propellants have developed as the preferred method of powering most missiles and rockets for military, commercial, and space applications. Solid rocket motor propellants have become widely accepted because of the fact that they are relatively simple to formulate and use, and they have excellent performance characteristics. Furthermore, solid propellant rocket motors are generally very simple when compared to liquid fuel rocket motors. For all of these reasons, it is found that solid rocket propellants are often preferred over other alternatives, such as liquid propellant rocket motors.
Conventional solid propellant rocket motors operate by generating large amounts of hot gases from the combustion of a solid propellant formulation stored in the motor casing. Typical solid rocket motor propellants are generally formulated having an oxidizing agent, a fuel, and a binder. At times, the binder and the fuel may be the same. In addition to the basic components set forth above, it is conventional to add various plasticizers, curing agents, cure catalysts, and other similar materials which aid in the processing and curing of the propellant. A significant body of technology has developed related solely to the processing and curing of solid propellants, and this technology is well known to those skilled in the art.
During operation, the gases generated from the combustion of the solid propellant accumulate within the combustion chamber until enough pressure is amassed within the casing to force the gases out of the casing and through an exhaust port. The expulsion of the gases from the rocket motor and into the environment produces thrust. Thrust is measured as the product of the total mass flow rate of the combustion products exiting the rocket multiplied by the velocity of the exiting combustion products plus the product of the change in pressure at the exit plane multiplied by the exit area. Increasing the pressure at which the gases are expelled from the combustion chamber raises the thrust level, which in turn increases the propulsion rate of the vehicle containing the rocket motor to thereby permit the vehicle to achieve higher speeds.
One type of propellant that is widely used incorporates ammonium perchlorate (AP) as the oxidizer. The AP oxidizer may then, for example, be incorporated into a propellant that is bound together by a hydroxy-terminated polybutadiene (HTPB) binder. Such binders are widely used and commercially available. It has been found that such propellant compositions provide ease of manufacture, relative ease of handling, good performance characteristics; and are at the same time economical and reliable. In essence it can be said that AP composite propellants have been the backbone of the solid propulsion industry for approximately the past 40 years.
High-performance booster propellants are being developed not only because they can deliver superior results but also because they can be used in numerous propulsion systems. One of the problems associated with conventional propellant formulations having an exponentially increasing propellant burn rate is that an increase consumption of propellant generally increases the operating pressure, which in turn increases the risk of catastrophic failure of the rocket motor casing. This insufficiency springs from the inherent combustion characteristics of AP, which often impart the propellant a high-pressure slope break between 2500 to 3500 psi. Therefore, to avoid this slope break region, propulsion systems designers generally devise motors that operate below 2000 psi, even though it is known that much higher performance can be achieved if the motor can operate at pressures higher than 2000 psi. The change in burning rate (r
b
) as a function of the pressure change, P
c
, is defined as the burning rate slope, n:
n
=
(
∂
ln
r
b
∂
ln
P
c
)
T
Data for determining burning rates at different pressures are typically gathered either by standard strand testing or by test motor analysis. The determination of burning rates by such testing procedures is well known in the art. Generally, conventional solid rocket propellant formulations have burn rate slopes of 0.15 or greater, but well below 1.0. Conventional propellants usually exhibit a dramatic positive increase in burning rate slope at pressures above about 2500 psi to 3500 psi, as illustrated in FIG.
1
.
The conventional solution to avoiding catastrophic failure of the rocket motor casing is to strengthen the rocket motor casing by constructing the casings with thick walls from strong, dense materials, such as steel. This approach, however, deleteriously imparts a severe weight penalty to the vehicle. Consequently, a greater amount of thrust and an increased propellant burn rate is required to propel the vehicle at a comparable rate. Because of the recent successes in manufacturing new composite motor cases, which can sustain much higher chamber pressures, it is possible to develop a motor that can deliver increased performance as measured by specific impulse (Isp).
Regardless of advances in rocket motor casings, it is highly desirable to search for a high-energy propellant material that does not exhibit a pressure slope break. While the exact cause is unknown, one theory is that it is due to a change in the mechanism from diffusion flame control to solid AP flame control at the higher pressures that exist in burning propellants, as described in “A Review of Models and Mechanisms for Pressure Exponent Breaks in Composite Solid Propellants”, Cohen, N.,
Proceedings of
23
rd
JANNAF Combustion Meeting
, CPIA Publication 457, Vol. 1 (October 1986), incorporated herein by reference. For example, burning solid AP is much more sensitive to pressure fluctuation (i.e., high slope) than materials experiencing diffusion-control-type burning.
Because conventional booster propellants (HTPB/AP/aluminum) are often composed of a mixture of different AP particle sizes, optimum packing efficiency (i.e., good processing properties) is important in attaining high performance and, at times, in providing burn rate tailorability. The various AP powders often contain large pieces (i.e., 200 &mgr;m to 400 &mgr;m) and it is extremely likely that the large-size AP particles contribute to the pressure slope break problem. Previous work has demonstrated that the slope break is delayed from 2500 psi to much higher pressures (as high as 10,000 psi) when the propellant formulations do not contain large AP particles. Therefore, many formulations instead incorporate high levels of ultrafine AP and exotic burn rate catalysts, such as superfine iron oxide, chromic oxide, catocene, or carboranes to achieve the higher pressure slope break.
Attempts have been made to solve the problems associated with the high-pressure slope break that occurs between 2500 psi and 3500 psi with ammonium perchlorate as an oxidizer. The problem is discussed in detail in the following publications: Atwood, A. I. et al., “High-pressure Burning Rate Studies of Ammonium Perchlorate (AP) Based Propellants,”
Proceedings for Research and Technology Agency of North Atlantic Treaty Organization
(
NATO
) 1999
Meeting on Small Rocket Motors and Gas Generators for Land, Sea, and Air Launched Weapon Systems
, Apr. 19-23, 1999, Corfu, Greece and Boggs, T. L., et al., “Ammonium Perchlorate Combustion: Effects of Sample Preparation; Ingredient Type; and Pressure, Temperature and Acceleration Environments,”
J. Combustion Science and Technology
, Vol. 7, pp. 177-183 (1973), incorpor
Chan May L.
Turner Alan D.
Haley Charlene A.
Hardee John
The United States of America as represented by the Secretary of
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