The development of solid propellant thruster technology, both of propellant formulations and system configurations, largely concerns methods of burn rate control, and thus thrust control to satisfy mission design requirements. Key to successful thrust control is not only the ability to adjust the magnitude of thrust, but also the ability to completely extinguish the combustion process upon command and the facility to have subsequent reignition/extinguishment cycles, all on command. Ideally, such thrust control is achievable in real-time and not merely by a thrust profile engineered into the system at the time of manufacture. Although thrust control is readily achievable for liquid propulsion systems and hybrid propulsion systems, these systems are more expensive, more complex and less reliable than comparable solid propellant systems. In addition, the majority of thruster systems that utilize liquid propellants are highly toxic.
Consider, for example a dual thrust mode of operation. The boost/sustain thrust profile is designed to initially provide high thrust to accelerate the propelled object, followed by a reduced thrust mode of operation that merely sustains the propelled object's velocity, thereby enabling extended propellant burn times. The difficulties inherent to controlling conventional propellant burn rates and the state-of-the-art solutions are reviewed in U.S. Pat. No. 6,086,692, issued to Hawkins, et al. and is hereby incorporated by reference. One problem identified by Hawkins, et al. associated with conventional propellant formulations, is the inherent risk of catastrophic failure due to the potential for burn rate runaway. Burn rate runaway can occur because of the relationship between propellant burn rate and pressure. Typically, as pressure is increased, the burn rate increases causing an exponential increase in burn rate and again an increase in pressure, ultimately terminating in critical failure of the motor. Accordingly, controlling the pressure is critical and difficult for such thrusters. Hawkins, et al. propose using propellants whose burn rate is relatively insensitive to pressure offering greater reliability through consistent performance and ease of predictability. However, two significant barriers to total thrust control of solid propellant systems remain with this system: the inability to alter the thrust profile in real-time, and the inability to completely extinguish the propellant combustion followed by reignition upon demand.
Another means of controlling propellant burn rate, is the use of propellants whose burn rate may be tailored by the addition of varying amounts of a given chemical ingredient. This class of propellants can be used to tailor the thrust profile by varying, in accordance with the desired thrust profile, the density of the critical burn rate controlling ingredient throughout the propellant grain. Numerous specific examples of this design philosophy exist, as represented by the following two patents: U.S. Pat. No. 5,372,070 to Neidert, et al., and in U.S. Pat. No. 4,411,717 to Anderson. However, the method of burn rate control by variations in chemical composition throughout the grain also suffer the problems of extinguishability and thrust profile alteration in real-time.
Alternatively, systems have been designed to control propellant burn rates through mechanical means. For example, U.S. Pat. Nos. 4,952,341 to Sayles and 4,756,251 to Hightower, et al. Sayles discloses the method of embedding shrink tubes or spheres within the propellant for the purpose of enhancing propellant burn rate, and Hightower, et al. disclose a method of embedding within the propellant grain a series of reticulated structures such that the propellant burn rate is variable, and similarly the thrust profile may be shaped as desired. Again, though, the thrust profile of the thruster using such a system is fixed at the time of manufacture and there is no provision for a start/stop/restart capability.
Yet another common mechanical means of controlling the burn rate is to construct the grain of the propellant with deliberate variations in exposed propellant surface area and the intentional inclusion of void spaces of designed dimensions. For well characterized propellants, these methods are tolerable, but in addition to the aforementioned problems of extinguishment and the lack of real-time thrust profile adjustment, the efficiency of these motors is decreased due to the increased volume that follows from the requirement for void spaces to control pressure and thus burn rate.
To achieve total control of the thrust profile, not only must propellant burn rate be actively controlled, thrusters must also have the capability of many start/stop/restart cycles. The ability to stop and then restart liquid propellant and hybrid propellant thruster systems is well established; however, problems of high cost, propellant toxicity, system complexity, and reliability encourage the search for solid propellant thruster systems that are capable of multiple start/stop/restart cycles.
Two common methods of achieving multiple start/stop/restart cycles for solid propellant thruster systems are the inclusion of discrete segments of propellant charges separated by some form of frangible material with a separate ignition means for each segment, or a pressure release system designed to vent the chamber pressure rapidly enough that extinguishment occurs. For example, consider U.S. Pat. Nos. 4,972,673 to Carrier, et al., 5,675,966 to Dombrowski, et al., and 4,078,953 to Sayles in which dual stage solid propellant rocket motors that rely upon a physical barrier between two distinct propellant grain regions to be broken upon command are disclosed. Again, although each of these concepts addresses multi cycle operation, none offer complete control, as the segments are always of finite capacity thus preventing smooth variation in the thrust profile, and, as always, the potential for unintentional segment ignition is present.
A particularly innovative means of thrust control that also achieves the zero thrust, start/stop multicycle operation is delineated in McDonald, et al.'s disclosure, U.S. Pat. No. 4,840,024 in which a complex system of valves and ducting is employed to control thrust. However, a detrimental aspect of McDonald, et al.'s solution to the problem of thrust control and restartability is the extreme complexity of the valving apparatus and the significant mass penalty. In addition, the exhaust from a gas generator charge is routed to pass over the main propellant charge for ignition and the exhaust is routed out of the combustion chamber during the extinguishment procedure. Finally, another problem is that the gas generator propellant charge is never extinguished and has a limited lifetime.
Finally, the concept of utilizing electric power to ignite and control propellant burn rate has been previously discussed. Electric power provides one standard means of igniting typical rocket propellants, in that an element is ohmically heated to temperatures above that required for propellant ignition. A comprehensive review of the application of electric power to the problem of enhancing efficiency in gun ammunition is provided in U.S. Pat. No. 5,515,765 to Wilkinson. This patent discloses the concept of controlled ignition of propellants through the use of electrothermal chemical cartridges. However, this application of electrical power to energetic materials is limited to the control of burn rate only, combustion continues despite the removal of electric power.
A related patent, U.S. Pat. No. 5,854,439 to Almstrom, et al. also discloses a method for electrically initiating and controlling the burning of a propellant charge. The innovation in this case relates to the inclusion of multiple and separate electrically conductive surfaces within the propellant for the purpose of passing electric current through the conductive segment, causing electrothermal heating of the segment followed by ignition of the propellant within that segment. The concept capitalizes solely upon the ohmic heating of a conductor to initiate propellant combustion and is unsuited, as disclosed, to the application of thrusters because of the extreme requirement for electric power. More specifically, the patent states the electric power unit is estimated to weigh 100-300 kg per kg charge weight of the propellant, a ratio clearly unfavorable to rocketry.
The state-of-the-art in controllable energetic materials that are influenced by the application of electric power is disclosed in a provisional patent to Katzakian, et al. Prov. Application No. 60/287,799, filed on Apr. 30, 2001, the disclosure of which is incorporated herein by reference. The disclosed invention describes a class of solid propellants whose combustion mechanism depends on electric power such that application of electric power above some critical value results in propellant ignition with subsequent sustained combustion and application of electric power below some critical value results in a cessation of combustion; however, no discussion of how the combustion may be controlled by the application of electric power is given, nor is any mention made of the mechanism.
Accordingly, a need exists for a state-of-the-art propellant control that enables economic and simple thrust control.