Patent Application
20250081321
The present invention relates generally to spacecraft propulsion systems and, more particularly, to spacecraft pulsed plasma thruster systems.
A spacecraft is a machine or vehicle that is designed to operate in space. Such spacecraft include, but are not limited to, rockets, space shuttles, satellites, and space stations. Spacecraft are used for a variety of purposes, including communications, navigation, scientific research and discovery, meteorology, and the like. Spacecraft propulsion is any method used to accelerate a spacecraft. In-space propulsion exclusively deals with propulsion systems used in space which is a near vacuum environment. Space is a near vacuum environment, which presents difficulties in operating, maintaining, and fueling spacecraft. These difficulties are magnified by the large distances involved and the consequential costs and timings to perform those operations.
Spacecraft propulsion systems are typically divided into four basic groups: (1) chemical propulsion, (2) electric propulsion, (3) advanced propulsion, and (4) supporting technologies; based on the physics of the propulsion system and how it derives thrust.
Chemical propulsion obtains energy needed to generate thrust by chemical reactions that create a hot gas that is expanded to produce thrust. Many different propellant combinations have been used to obtain these chemical reactions, including, for example but not limited to, hydrazine, liquid oxygen, liquid hydrogen, nitrous oxide, and hydrogen peroxide. Non-toxic “green” alternatives are now being developed to replace hydrazine. ASCENT propellant (formerly known as AF-M315E) is an example of one such green propellant. Initially developed to provide a safer handling environment than hydrazine, ASCENT offers higher overall performance than a hydrazine equivalent by creating more heat during ignition.
Electric propulsion is commonly used for station keeping of communications satellites and for prime propulsion on some scientific space missions because of their high specific impulse. Rather than relying on high temperature and fluid dynamics to accelerate the reaction mass to high speeds, there are a variety of methods that use electrostatic or electromagnetic forces to accelerate the reaction mass directly, where the reaction mass is usually a stream of ions. Electromagnetic propulsion systems include, but are not limited to: (1) ion thrustors which first accelerate and later neutralize the ion beam with an electron stream emitted from a cathode; (2) electrothermal thrusters which use electromagnetic fields to generate a plasma to increase the heat of a propellant and is then converted into kinetic energy by a nozzle; and (3) electromagnetic thrusters which accelerate ions by either Lorentz Force or electromagnetic fields. Electromagnetic thrusters include but are not limited to: (1) plasma propulsion engines which generate thrust from a quasi-neutral plasma; (2) magnetoplasmadynamic thrusters (MPDT) which use the Lorentz force to generate thrust; (3) electrodeless plasma thrusters uses ponderomotive forces to accelerate a plasma; (4) pulsed inductive thruster (PIT) which uses perpendicular electric and magnetic fields to accelerate a propellant with no electrode; (5) pulsed plasma thruster (PPT) which passes an arc through fuel to create a plasma which completes a circuit between two charged plates and is accelerated out an exhaust to create thrust; and (6) vacuum arc thrusters (VAT) which use a vacuum arc discharge, across an insulator, between two electrodes to produce thrust.
The rapid development small spacecraft including, but not limited to, micro and nano-satellite technology, has sparked interest in robust, low power, and high specific impulse propulsion systems. PPTs have been extensively investigated and employed to fill such needs. PPTs accelerate plasma propellant through the Lorentz-force-preferably with a minimum of thermal and electromagnetic loss. In lieu of an applied external magnetic field, an induced component resulting from current traversal through the electrodes and plasma supplies the B-field required for acceleration. High discharge currents (typically around 100 amps) enable sufficiently high J×B Lorentz-force magnitudes (where J is the current density and B is the magnetic field). Typically, PPTs utilize a capacitor bank for energy storage, which is subsequently converted into kinetic motion, heating, and propellant ionization upon initiation of the discharge. Accordingly, stored energy may be implemented in power-limited nanosatellites for attitude control maneuvers. Another particularly significant benefit to pulsed operation is increased thrust efficiency, enabled by the ability to operate at higher discharge currents without thermal electrode damage. The operation at these higher discharge currents results in an increase in thrust.
Ablative pulsed plasma thrusters (APPTs) typically operate with solid phase propellants such as polytetrafluoroethylene or similar fluorocarbons. Surfaces of these propellants are vaporized with high currents, and the resulting plasma is accelerated to produce thrust. This process places a harsh limit on efficiency (typically <15%)—plagued by late-ablation and the presence of thermally expelled macroparticles. Nonuniform ablation, low mass flow control, and contamination pose further problems to the implementation of APPTs. These disadvantages are often disregarded, however, as high reliability and low tankage fractions often provide an attractive thruster option.
Gas-fed pulsed plasma thrusters (GF-PPTs) pose a stark contrast to APPTs. These variants offer relatively high efficiencies (typically 20-70%) and precision mass flow-control at the cost of complex injection systems. GF-PPTs can strain implementation in satellites which have limited volumetric capabilities, namely CubeSats. Gas injection arrangements also place a mechanical limit on firing frequency due to the limitations imposed by mechanical solenoid valves. Further, difficulties often arise in attempting to establish a desired mass density distribution before ignition-where inductive delay and gas injection must be properly timed.
Liquid-fed pulsed plasma thrusters (LF-PPTs) have been proposed which include two pairs of electrodes. The first pair of electrodes form an igniter assembly that ignites the liquid propellant to form a plasma cloud, and the second pair of electrodes form a Lorentz-force pulsed plasma accelerator that receives and accelerates the plasma cloud generated by the igniter assembly. For example, but not limited to, see U.S. Pat. No. 11,554,883, the disclosure of which is expressly incorporated herein in its entirety by reference.
Although PPTs with solid and gaseous propellants have been employed with limited degrees of success in small spacecraft, simple and robust PPT systems have yet to be developed for such applications. Additionally, proposed PPTs with liquid propellants appear to be overly complex and unreliable. Accordingly, there remains a need for simple and robust PPT systems for use on small, as well as larger, spacecraft applications.
The present invention overcomes the foregoing problems and other shortcomings, drawbacks, and challenges of spacecraft pulsed plasma thruster systems. While the invention will be described in connection with certain embodiments, it will be understood that the invention is not limited to these embodiments. To the contrary, this invention includes all alternatives, modifications, and equivalents as may be included within the spirit and scope of the present invention. According to one embodiment of the present invention, a conductive liquid-fed pulsed plasma thruster comprises a first electrode having a conductive solid portion and a conductive liquid portion, a second electrode separated from the first electrode to define an ignition space therebetween, at least one electric insulator separating the first and second electrodes, and a conductive-liquid passage extending within the conductive solid portion through which the conductive liquid portion flows from an inlet to an outlet located at the ignition space. The first and second electrodes are configured so that a drop of the conductive liquid forms and grows at the outlet when the conductive liquid flows through the conductive liquid passage until the drop of the conductive liquid causes an arc discharge between the drop and the second electrode to ignite the drop and produce a plasma cloud that generates thrust when exhausted.
According to another embodiment of the present invention, a spacecraft propulsion system comprises a conductive liquid-fed pulsed plasma thruster including a first electrode having a conductive solid portion and a conductive liquid portion, a second electrode separated to define an ignition space therebetween, at least one electric insulator separating the first and second electrodes, and a conductive-liquid passage extending within the conductive solid portion through which the conductive solid portion flows from an inlet to an outlet located at the ignition space. The first and second electrodes are configured so that a drop of the conductive liquid portion forms and grows at the outlet when the conductive liquid portion flows through the conductive liquid passage until the drop of the conductive liquid causes an arc discharge between the drop and the second electrode to ignite the drop and produce a plasma cloud that generates thrust when exhausted. The spacecraft propulsion system also comprises a power source operatively connected to the first and second electrodes, a conductive-liquid reservoir for holding the conductive liquid; and a pump connected in fluidic communication with the conductive-liquid reservoir and the inlet of the first electrode. According to yet another embodiment of the present invention, a method for propelling a spacecraft is provided for a spacecraft including a conductive liquid-fed pulsed plasma thruster comprising a first electrode having a conductive solid portion and a conductive liquid portion, and a second electrode separated from the first electrode to define an ignition space therebetween, at least one electric insulator separating the first and second electrodes, and a conductive-liquid passage extending within the conductive solid portion through which the conductive liquid portion flows from an inlet to an outlet located at the ignition space. The method comprises the steps of forming—a drop of the conductive liquid outside the outlet of the first electrode by supplying the conductive liquid into the inlet of the first electrode, growing the drop of the conductive liquid outside the outlet of the first electrode by continuing to supply the conductive liquid into the inlet of the first electrode until the drop of the conductive liquid causes an arc discharge between the drop of the conductive liquid and the second electrode that ignites the drop of the conductive liquid to produce a plasma cloud that generates thrust when exhausted, and after generating thrust, repeating the steps of forming the drop and growing the drop.
Additional objects, advantages, and novel features of the invention will be set forth in part in the description which follows, and in part will become apparent to those skilled in the art upon examination of the following or may be learned by practice of the invention. The objects and advantages of the invention may be realized and attained by means of the instrumentalities and combinations particularly pointed out in the appended claims.
The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with a general description of the invention given above, and the detailed description of the embodiments given below, serve to explain the principles of the present invention.
It should be understood that the appended drawings are not necessarily to scale, presenting a somewhat simplified representation of various features illustrative of the basic principles of the invention. The specific design features of the sequence of operations as disclosed herein, including, for example, specific dimensions, orientations, locations, and shapes of various illustrated components, will be determined in part by the particular intended application and use environment. Certain features of the illustrated embodiments have been enlarged or distorted relative to others to facilitate visualization and clear understanding. In particular, thin features may be thickened, for example, for clarity or illustration.
The following examples illustrate particular properties and advantages of some of the embodiments of the present invention. Furthermore, these are examples of reduction to practice of the present invention and confirmation that the principles described in the present invention are therefore valid but should not be construed as in any way limiting the scope of the invention.
The illustrated CLF-PPT 112A (best shown in
The illustrated first electrode 126 includes an electrically conductive solid portion 126A and an electrically conductive liquid portion 126B. The illustrated solid portion 126A of the first electrode 126 is a substantially straight and elongate circular tube having a centrally located and substantially straight conductive-liquid passage 134 extending entirely therethrough. A first or rear end of the conductive-liquid passage 134 forms an inlet 136 and a second or front end of the conductive-liquid passage forms an outlet 138. The outlet 138 is located at the ignition space 130. The solid portion 126A of the first electrode 126 can comprise any suitable electrically conductive material such as, for example but not limited to, a metal such as, for example but not limited to, copper or the like. The liquid portion 126B of the first electrode 126 is formed by the conductive liquid 120 located within the conductive liquid passage 134 and any conductive liquid 120 forwardly extending therefrom into the ignition space 130 as described in more detail hereinafter. The illustrated conductive-liquid passage 134 is needle like or elongate so that the liquid portion 126B of the first electrode 126 located therein has a high surface area to volume ratio so that the liquid portion 12B of the first electrode 126 receives an adequate electrical charge from the solid portion 126A of the first electrode 126 which is in contact therewith via the conductive-liquid passage 134. The rear end of the illustrated solid portion 126A of the first electrode 126 is formed to receive an outlet end of a first liquid conduit 140 in a manner that the conductive liquid can flow from the first liquid conduit 140 to the inlet 136 of the conductive liquid passage 134 as described in more detail hereinbelow. It is noted that the first electrode 126 can alternatively have any other suitable configuration or form.
The illustrated at least one insulator 132 is a substantially straight and elongate circular tube having a centrally located and substantially straight passage 142 extending entirely therethrough. The passage 142 is sized and shaped for receiving the solid portion 126A of the first electrode 126 therein. The illustrated insulator 132 is mechanically secured to the solid portion 126A of the first electrode 126 but can be additionally or alternately secured thereto in any other suitable manner. The insulator 126 can comprise any suitable electrically insulating material, that is a dielectric material such as, for example but not limited to, a dielectric ceramic and the like. It is noted that alternatively there can be more than one insulator 132, that is, a plurality of the insulators 132 electrically isolating the second electrode 128 from the first electrode 126. It is also noted that the insulator 132 can alternatively have any other suitable configuration or form.
The illustrated second electrode 128 is a substantially straight and elongate circular tube having a centrally located and substantially straight passage 144 extending entirely therethrough. The passage 144 is sized and shaped for receiving a forward portion of the insulator 132 therein so that the second electrode 128 forwardly extends from the forward end of the isolator 132 in a cantilevered manner and encircles the ignition space 130. The passage 144 has an open forward end 145 in communication with the ignition space 130. The illustrated second electrode 128 is mechanically secured to the isolator 132 but can be additionally or alternately secured thereto in any other suitable manner. The second electrode 128 can comprise any suitable electrically conductive material such as, for example but not limited to, a metal such as, for example but not limited to, copper or the like. It is also noted that the second electrode can alternatively have any other suitable configuration or form.
The illustrated first and second electrodes 126, 128 are thus in a coaxial configuration, that is, a configuration where an arc forms in a radial direction within the ignition space 130 between the free outer end of the conductive liquid portion 126A of the first electrode 126 and an inner contact surface 146 of the second electrode 128 within the second electrode passage 144 at the ignition space 130. The first and second electrodes 126 and 128 are also configured so that a drop 148 of the conductive liquid 120 forms and grows at the outlet 138 of the solid portion 126A of the first electrode 126 as the conductive liquid 120 flows through the conductive liquid passage 134. Once the conductive liquid 120 traverses the solid portion 126A, a droplet or drop 148 forms with its shape constrained by the surface tension of the liquid.
The drop 148 continues to grow until the drop 148 of the conductive liquid 120 grows in the radial direction enough to cause an arc discharge between the first and second electrodes 126, 128. That is, when a part of the drop 148 becomes close enough to the second electrode 128 that the voltage between second electrode 128 and drop 148 exceeds the breakdown voltage of any residual gas or outgassing within the tube-shaped second-electrode 128. The arc of electrons from the second electrode 128 to the drop 148 is struck between the conductive inner surface 146 of second electrode 128 and the liquid surface of the drop 148. This arc deposits electron energy within a thin surface of the droplet material which is heated and vaporizes. As the surface of the drop 148 vaporizes, the vacuum gap fills with conductive plasma facilitating additional ablative energy deposition in the neutral droplet material. The goal is sizing the capacitive stored energy 116 and the high voltage converter 117 to match the deposited energy with the ablative consumption of the droplet material back towards the solid portion 126A of the first electrode 126. That is, depleting the stored the stored energy prior to erosion of solid electrode material (either anode or cathode). In other words, the plasma discharge vaporizes all of the drop 148, but the stored energy is low enough to mitigate solid electrode erosion.
Once the superheated high-density plasma cloud is formed, it expands out of the tubular-shaped second electrode 128 and through the nozzle 145. While the plasma temperature exceeds the melting temperature of the tube wall material of the second electrode 128, rapid expansion and cooling makes the duration of this over-temperature condition brief. The material's thermal inertia is tuned such that the surface of the material can withstand this brief high temperature exposure. As a plasma plume expands down the tubular-shaped second electrode 128 and cools, radiation from within the plasma and from the heated walls of the second electrode 128 is recaptured allowing further acceleration of the cooling plasma. With the high densities encountered in the plasma, rapid recombination occurs within the accelerated flow enabling a primarily neutral high ISP plume to be ejected from the thruster 112A to generate thrust. The tube length of the second electrode 128 is sized to optimize this acceleration while allowing for sufficient blow-down to vacuum conditions to allow for a low enough density such that the vacuum resistivity is sufficient to avoid premature breakdown of the subsequent pulse. This blowdown, in combination with the capacitive stored energy 116 and thermal constraints, determine the maximum sustained repetition rate for the pulse cycle.
It is noted that the first and second electrodes 126, 128 can alternatively have any other suitable configuration. For example, but not limited to, in a coaxial configuration the first electrode 126 can be a solid electrode and the second electrode 128 can be a liquid sacrificial electrode so that a drop 148 of conductive liquid 120 grows radially inward from the second electrode 128 to the first electrode 126. Also, for example but not limited to, when in a coaxial configuration both the first electrode 126 and the second electrode 128 can be liquid sacrificial electrodes so that a pair of drops 148 of conductive liquid 120 grow radially inward toward each other from the first electrode 126 and the second electrode 128.
The capacitive stored energy 116 can be of any suitable type to provide a suitable high-energy voltage difference across the first and second electrodes 126, 128 to operate as described herein. The illustrated capacitive stored energy 116 is a high-voltage capacitor bank for storing energy. The high voltage capacitor bank is initially charged using a DC-DC high energy or boost convertor 117 from available low voltage spacecraft bus. Charging time depends on current and bus current limit boost convertor 117, but these only limit repetition rate rather than other performance parameters of the thrusters 112A. The capacitor bank voltage and capacity are sized to match the desired energy deposition within one ablative pulse of propellant for the target dense plasma slug. It is noted that the high voltage power can alternatively be provided by any other suitable type of power supply.
The capacitive stored energy 116 is electrically connected to the first and second electrodes 126, 128. In the illustrated embodiment a negative terminal of the capacitive stored energy 116 is operably connected to the first electrode 126 and a positive terminal of the capacitive stored energy 116 is operably connected to the second electrode 128. Thus, the first electrode 126 operates as the anode and the second electrode 128 operates as the cathode. It is noted, however, that the polarity can be reversed if desired. In this case the positive terminal of the capacitive stored energy 116 is operably connected to the first electrode 126 and the negative terminal of the capacitive stored energy 116 is operably connected to the second electrode 128. Thus, the first electrode 126 operates as the cathode and the second electrode 128 operates as the anode. It is noted that the power source can alternatively have any other suitable configuration.
The illustrated conductive-liquid reservoir or tank 118 forms an enclosed interior space for storing a suitable quantity of the conductive liquid 120. The reservoir 118 can be formed of any suitable material. An outlet of the conductive liquid reservoir 118 is in fluidic communication with an inlet of the pump 122 via a second liquid conduit 150. The second liquid conduit 150 can be of any suitable type. It is noted that the conductive liquid reservoir 118 can alternatively have any other suitable configuration.
The conductive liquid 120 preferably has a low vapor pressure and can be of any suitable type which performs as described herein. The vapor pressure preferably is low enough that the arc does not break down prematurely and too much mass is not lost to make the thruster 112A inefficient. This may not be required if a lossy liquid is utilized. The conductive liquid 120 can be, for example but not limited to, (1) an ionic liquid such as, for example but not limited to 1-ethyl-3-methylimidazolium tetrafluoroborate (EMI-BF4), and the like, (2) an energetic liquid or propellant such as, for example but not limited to, Advanced Spacecraft Energetic Non-Toxic (ASCENT) formerly known as AF-M315D, hydrazine, and the like, or (3) a liquid metal such as, for example but not limited to lithium, mercury, and the like.
The pump 122 can be of any suitable type for pumping the conductive liquid 120 from the conductive liquid reservoir 118 to the CLF-PPT 112A during operation of the CLF-PPT 112A. The pump 122 can be, for example but not limited to, a needle pump, syringe pump, infusion pump, and the like. An inlet of the pump 122 is in fluidic communication with the outlet of the conductive liquid reservoir 118 via the second liquid conduit 150. An outlet of the illustrated pump 122 is in fluidic communication with the inlet 136 of the first electrode 126 via the first liquid conduit 140. It is noted that the pump 122 can alternatively have any other suitable configuration.
The controller 124 comprises suitable processors and memory and is programmed for operating the components of the propulsion system 110A as described herein. The controller 124 can be a stand-alone component or part of a larger controller/computer of the spacecraft 100. The illustrated controller 124 is in operable communication with the capacitive stored energy 116 and the pump 122. The illustrated controller 124 is configured to operate the capacitive stored energy 116 to provide voltage to the first and second electrodes as needed. The illustrated controller 124 is also configured to operate the pump 122 as needed to provide the conductive liquid 120 to the CLF-PPT 112A as needed to fire the CLF-PPT 112A as needed. It is noted that the controller 124 can alternatively or additionally be configured to operate any other components as needed. It is also noted that the illustrated propulsion system 110A is simplified compared to the prior art propulsion systems because an igniter system is not required, and a switching system is not required.
The second electrode 128 is a body of suitable material having a planar contact surface 156. The illustrated second electrode 12SB is generally shaped as a rectangular block but any other suitable shape can alternatively be utilized. It is noted that the planar contact surface 156 can also have any other suitable shape. The second electrode 128 is positioned so that the planar contact surface 156 is spaced-apart from and facing the conductive liquid passage outlet 138 of the first electrode 126. Thus, the planar contact surface 156 is substantially perpendicular to the direction of travel of the conductive liquid 120 exiting the conductive liquid passage outlet 138 of the first electrode 126. As the drop 148 of conductive liquid 120 forms and grows, it ignites once it has sufficiently grown in the linear direction toward the contact surface 156 of the second electrode 128 to cause an arc therebetween.
The second electrode 128 is configured as a liquid sacrificial electrode like the first electrode 126 having a conductive solid portion 128A and a conductive liquid portion 128B. Both the first and second electrodes 126, 128 are provided with the conductive liquid 120 by the pump 122 via the first liquid conduits 140. It is noted that alternatively a second pump could be utilized 122. The second electrode 128 is spaced apart from and facing the first electrode 126 so that the conductive liquid passages 134 are coaxial. Thus, the direction of travel of the conductive liquid 120 exiting the conductive liquid passage outlets 138 of the two first electrodes 126 are directly opposed to one another. That is, aligned on a collision course. As the drops 148 of conductive liquid 120 form and grow toward one another, they each ignite once they have sufficiently grown in the linear direction toward one another to cause an arc therebetween.
The insulator 132 of the fourth embodiment of the present invention is provided with planar wetting walls 160 located adjacent to the outlets 138 of the conductive liquid passages 134 and perpendicular to the outlets 138 of the conductive liquid passages 134. The wetting walls 160 are configured and positioned so that the drops 148 form along the wetting walls 160 and “wet” the walls 160. The drop 148 forms with its shape constrained by the surface tension of the liquid and liquid-solid wetting angle. The wetting wall 160 support the drops 148 unlike the unsupported free drops 148 of the first and second embodiments. The wetting walls 160 can be of any suitable size and can alternatively have any other suitable configuration.
The illustrated insulator 132 includes a pair of laterally spaced apart side walls 162 that enclose the space between the first and second electrodes 126, 128 including the ignition space 130. A top portion of the side walls 162 angle outward in an to form a nozzle 164. It is noted that the side walls and/or nozzle can alternatively have any other suitable configuration.
This multimode concept enables flexibility between low thrust, high efficiency electric propulsion maneuvers and high thrust, low efficiency chemical propulsion maneuvers. This flexibility provides significant advantages for spacecraft resilience. A liquid fed pulsed plasma thruster, especially one compatible with multimode propulsion propellants, has the potential to support this role. Compared to other electric propulsion systems, the pulsed plasma thruster has a number of potential advantages such as small size and minimal complexity. The potential compatibility with existing chemical thruster form factors also makes it a promising candidate for a single thruster/single tank ideal configuration for a multimode system.
It can be appreciated from the above disclosure that a conductive liquid can be used as both propellant and sacrificial anode in a pulsed plasma thruster in order to mitigate performance limitations due to anode erosion in pulsed plasma thrusters. By flowing the conductive liquid through a high temperature dielectric material into a conductive solid material, the liquid itself can serve as both propellant and anode in striking a high voltage arc discharge to deposit high energy density into a small volume of propellant. The propellant is ohmically ablated and ejected from the thruster to provide thrust. As the anode material ablates, the surface recesses breaking the circuit to allow energy storage for the subsequent pulse. It can further be appreciated from the above disclosure that a further advantage of the disclosed low specific mass high power electric propulsion system compatible with existing ionic liquids (such as, for example but not limited to, EMI-BF4), energetic ionic liquids (such as, for example but not limited to, ASCENT), and liquid metal propellants (such as, for example but not limited to) lithium or mercury.
It can also be appreciated from the above disclosure that one of the key advantages of the thrusters according to the present invention is mechanical and electrical simplicity. It can further be appreciated that another advantage comes from avoiding the need for highspeed vacuum compatible pulsed electronics because the erosion of the fuel and subsequent liquid feed replenishment provides the necessary vacuum switching mechanism and avoids inefficiencies due to incomplete ionization and fuel blow-by and other inefficiencies of prior art active pulsed electronics systems.
It is noted that each of the features and components of the various embodiments of the present can be used with each of the other embodiments of the present invention if desired.
While the present invention has been illustrated by a description of one or more embodiments thereof and while these embodiments have been described in considerable detail, they are not intended to restrict or in any way limit the scope of the appended claims to such detail. Additional advantages and modifications will readily appear to those skilled in the art. The invention in its broader aspects is therefore not limited to the specific details, representative apparatus and method, and illustrative examples shown and described. Accordingly, departures may be made from such details without departing from the scope of the general inventive concept.
The invention described herein may be manufactured and used by or for the Government of the United States for all governmental purposes without the payment of any royalty.
