Patent Application
20220367077
Aspects of the present disclosure generally relate to a rocket thruster system, and more specifically, to a space-based isotope production system and method of use.
Many nuclear propulsion concepts have been proposed over the decades, using the heat from a reactor or radioisotope to either generate electricity for an electric thruster, or directly heat a propellant. In the case of the latter, the specific impulse is severely limited by the temperature of the materials and lifetime is limited by the corrosive nature of the low molecular weight propellants (such as hydrogen). In the case of the former, the reactor requires a heavy power conversion system such as Brayton or Rankine conversion with complex and expensive components that are difficult to make reliable enough for a spacecraft. Additionally, the nuclear-electric system requires a massive heat rejection system to dump waste heat, and heavy shielding to protect the payload from gamma-rays and neutrons.
Radioisotope thrusters have also been proposed, but to date require isotopes produced on Earth and integrated with a thruster during manufacturing. This conventional approach restricts the options to isotopes that are longer lived that may be handled and launched before expiration (such as Pu-238, or Po-210). However, longer lived isotopes have very low activity (low power density), and due to practical and safety limitations, the thrusters are quite small and low thrust. Only one known thruster concept has a relatively high-power density due its use of a shorter-lived alpha emitter such as Po-210 to generate a high-voltage potential between electrodes in an electrostatic thruster (See U.S. Pat. No. 3,184,915 to Low et al, incorporated by reference in entirety for all purposes).
Further, various reactor configurations and isotope production systems are deployed for terrestrial applications. For example, the High Flux Isotope Reactor at Oak Ridge National Laboratory is described as being based on the flux trap principle, in which the reactor core consists of annular regions of fuel surrounding an unfueled moderating region. The fast neutrons leaking from the fuel are moderated in the moderating region, thus producing a region of very high thermal-neutron flux that is “trapped” within the reactor (See “High Flux Isotope Reactor Technical Parameters,” Oak Ridge National Laboratory, https://neutrons.ornl.gov/hfir/pararneters (accessed May 1, 2022). Again, such reactors for isotope production are currently limited to earth-bound constructions.
The following presents a simplified summary of one or more aspects to provide a basic understanding of such aspects. This summary is not an extensive overview of all contemplated aspects, and is intended to neither identify key or critical elements of all aspects nor delineate the scope of any or all aspects. Its purpose is to present some concepts of one or more aspects in a simplified form as a prelude to the more detailed description that is presented later.
The disclosure overcomes the significant limitations of conventional nuclear propulsion concepts to include conventional radioisotope thruster concepts operated in space.
The disclosure describes various aspects of space-based radioisotope production system for use in a variety of use scenarios, such as a radioisotope rocket thruster and in-space charging station.
In an aspect, a propellant is accelerated by decay energy to yield thrust. The decay energy is provided by activating a target material. In one aspect, the thruster may be recharged or “reactivated” in a space-borne charging station. The activated isotopes may also be used generate electricity. The space-borne charging station may also be used for irradiating other items in space for any number of purposes.
In one aspect of a thruster application, a propellant is injected into a stream of ionizing radiation to yield thrust. The stream of ionizing radiation is provided by activating a target material.
In one embodiment, the stream of particles is controlled by a magnetic field to travel in a circular route within a thrust chamber encapsulating the propellant.
In some embodiments, a radiation absorbing substance is added to the propellant to facilitate heating of the propellant.
In one aspect, the thruster may be recharged or “reactivated” in a space-borne charging station.
In some embodiments, a space-borne radioisotope power system includes a chamber containing a target material, a neutron source in space for producing neutrons. The neutrons are used for activating the target material to produce a radioisotope material. The radioisotope material produces decay energy, and a propellant receives the decay energy to produce at least one of heat and electricity.
In other embodiments, a method for operating a space-borne radioisotope power system includes providing a space-borne radioisotope power system. The space-borne radioisotope power system includes a neutron source and a chamber containing a target material. The method further includes interfacing the space-borne radioisotope power system with a client spacecraft in space. The method also includes producing neutrons using the neutron source, activating the target material using the neutrons to produce a radioisotope material, producing decay energy from the radioisotope material to produce at least one of heat and electricity, and using the at least one of heat and electricity to provide power to the client spacecraft interfaced with the space-borne neutron source.
In another aspect, the client spacecraft may have been launched without radioactive material, then interfaced in space with a space-borne charging station equipped with a reactor or neutron source to activate an inert material on the spacecraft to enable the use of radioactive decay on the spacecraft while in space.
In still another aspect, the method further includes providing an orbital transfer vehicle for navigating the client spacecraft to the space-borne radioisotope power system. In an embodiment, the orbital transfer vehicle may also be charged at the space-borne neutron source.
The appended drawings illustrate only some implementation and are therefore not to be considered limiting of scope.
The detailed description set forth below in connection with the appended drawings is intended as a description of various configurations and is not intended to represent the only configurations in which the concepts described herein may be practiced. The detailed description includes specific details for the purpose of providing a thorough understanding of various concepts. However, it will be apparent to those skilled in the art that these concepts may be practiced without these specific details. In some instances, well known components are shown in block diagram form to avoid obscuring such concepts.
The present disclosure describes various embodiments of a space-borne radioisotope power system for providing power, heat, electricity, and/or thrust for objects in space. For instance, a space-borne reactor or other neutron source (such as a fusion reactor or particle accelerator) may be used to make an inert material radioactive on a client spacecraft for the purpose of using the radioactive decay for some purpose on that client spacecraft (such as generating electric power or thrust). The neutron source may be connected to its own spacecraft (bus) for the purpose of operating it remotely in space, and remains separate from the client spacecraft. The neutron source and its bus may be referred to as a “charging station.” In an example, this station may nominally stay parked in an orbit (e.g., around earth or another object) for use as stationary infrastructure to service client spacecraft. This configuration allows client spacecraft to launch without radioactive material, alleviating launch safety issues and reducing costs.
In an embodiment, the radioisotope power system is used for producing thrust by directing a stream of ionizing radiation to heat and react a propellant, wherein the ionizing radiation is generated by activating a target material. For instance, while beta decay has been utilized in the past to directly produce electricity for terrestrial applications and beta batteries are even sold commercially, the limited supply of isotopes produced by conventional means and safety requirements limits the applications for this type of power source. Furthermore, no known thruster concept uses a flux trap-based reactor to generate a radioisotope in space.
As an example, a space-borne radioisotope power system may be used as a radioisotope rocket thruster system to produce thrust by directing a stream of beta particles to heat a propellant, the beta particles generated by decay of an activated target material. Stated another way, the rocket thruster system is a spaceborne radioisotope power system. Activation of the target material occurs in situ, in the space environment. In an embodiment, a propellant is injected into a beta particle stream to produce thrust. The thrust is directed through a nozzle. In one embodiment, the stream of beta particles is controlled by a magnetic field within a thrust chamber encapsulating the propellant.
As disclosed herein, the spaceborne reactor system for producing the radioisotope relies on relatively short-lived isotopes (with half-lives of hours to months or a few years) made by activating readily available stable elements (such as phosphorus or scandium) while in space. Such stable elements are referred to in the disclosure as “target material.” A beta emitter may be used partly because most naturally occurring elements activate into beta emitters, and the particles are easily deflected with magnetic fields and/or are easily absorbed to generate heat. The beta particles may be steered into long cyclic paths (i.e., longer than the mean free path of the beta particles in the gas) to allow the beta particles to collide with propellant or other thruster components to deposit electric charge or heat. Due to the short half-life and to alleviate launch safety issues, the system activation may occur in space.
Within the present disclosure, the term “isotope” means any of two or more forms of a chemical element having the same number of protons in the nucleus (i.e., the same atomic number) but having different numbers of neutrons in the nucleus, or different atomic weights. The phrase “beta particle” means a high-speed electron or positron emitted by radioactive decay or reaction of an atomic nucleus.
In one embodiment, a reactor with a neutron flux trap large enough to contain the thruster with target material(s) and an opening for a spacecraft to insert the thruster resides in orbit as a kind of “charging station” that may be used by multiple spacecraft. Such reactors may be placed at various orbits or destinations (such as Mars orbit, Lunar orbit, geo-stationary orbit, etc.) so that spacecraft may be re-activated for return trips or other maneuvers. Such an embodiment may be incorporated as part of a multi-orbital transfer vehicle constellation system, as described in U.S. patent application Ser. No. 16/986,517 to Clark, incorporated by reference in entirety for all purposes.
Preliminary analysis shows that, for example, with natural phosphorus as a target material with resulting phosphorus-32 (14-day half-life, 1.7 MeV beta), a power density of 1 kW per kg of phosphorus is feasible. If the propellant mass flow rate is adjusted such that the Isp is 2000 sec., then the resulting initial thrust is about 0.1 N per kg of phosphorus. The resulting thruster has similar specific power to electric thrusters, but without the need for a power-supply, power-conversion, radiation shielding, or power transmission, resulting in a spacecraft specific mass of at least an order of magnitude lower than conventional electric spacecraft. Other shorter lived activated target elements such as manganese may produce higher thrust and specific power, albeit for much shorter durations. It is noted that the term “Isp” or “Isp” refers to specific impulse, which is a measure of impulse delivered to a vehicle per unit of propellant consumed, and thus measured how efficiently propellant is consumed.
Although the disclosed devices, systems, and methods of use will be described relative to space-borne radioisotope rocket thruster, the devices, systems, and methods of use have other applications. For example, the method and/or devices disclosed herein may be used in terrestrial applications.
The disclosed devices, systems, and methods of use will be described with reference to
Target material 11 is, for example, a stable element that forms a radioisotope material when impacted by a stream of neutrons. Once activated, target material 11 emits ionizing radiation, and radioactive decay is ongoing once activated. Candidate target materials may also include yttrium, which activates into Y-90, and other target materials known to those skilled in the art.
Neutron source 12 emits neutrons 20. Neutrons 20 are directed at target material 11 to activate target material 11 and produce or create radioisotope material and ionizing radiation 22.
Neutron source 12 may be a nuclear reactor with a neutron flux trap in one embodiment. Alternatively or additionally, neutron source 12 may be any type of fusion reactor, to include any device that produces or emits neutrons through a fusion reaction, such as for example solar-powered neutron generators. In one embodiment, neutron source 12 is a particle accelerator. Note that once target material 11 is activated, neutron source 12 may be removed from or distanced from the radioisotope engine for radioisotope engine operation.
Ionizing radiation 22 emitted from the radioisotope material formed by target material 11 is received by power conversion system 13, which controls the movement or kinematics of ionizing radiation 22. Power conversion system 13 receives ionizing radiation 22 and outputs or creates heat or electricity 24. A broad objective of power conversion system 13 as shown in
System processor/controller 16 may control one or more aspects or components of radioisotope rocket thruster system 10. For example, system processor/controller 16 may control the rate of emission of neutrons 20 from neutron source 12, and/or control or set parameters for power conversion system 13.
Power conversion system 13 may employ any one or more mechanisms to control ionizing radiation 22. In one embodiment, power conversion system 130 creates and employs a magnetic field to control or direct ionizing radiation 22 into a controlled beta particle stream forming a curvilinear route (see
Electricity or heat 24 from power conversion system 13, is directed to propellant 15. In one embodiment, propellant 15 is contained within and occupies the entirety of chamber 17. In an example, electricity or heat 24 is provided as a beta particle stream moving in a spiraled pattern within chamber 17 so as to remain within chamber 17 (and thus also within propellant 15) until the beta particles are absorbed or stopped by the propellant (see
Any of the steps, functions, and operations discussed herein with respect to
As shown in
Method 200 then proceeds to a step 212, at which the target material is positioned to receive neutrons from a neutron source, such as neutrons 20 from neutron source 12 of
Method 200 then proceeds to a step 216, at which the neutrons emitted or provided by the neutron source are directed to the target material, resulting in creation of an activated target material. As discussed above, once the target material is so activated, the target material produces ionizing radiation.
Continuing to refer to
Then, at a step 224, ionizing radiation is produced as a result of the creation of activated target material from the previously inactive target material.
At a step 228, the radioactive emission is controlled by a power conversion system, such as power conversion system 13 of
At a step 232, the heat or electricity is provided to a propellant (e.g., propellant 15 housed in chamber 17) thereby producing thrust. As described above, in one embodiment, the power conversion system directs or controls the radioactive emission so as for a circular or spiral route so as to remain within the enclosed chamber that encapsulates or contains the propellant. In one embodiment, the propellant flows through a radioisotope rocket thruster engine, the propellant heated by beta particles via collisions with gas molecules while in the chamber (e.g., chamber 17 of
At a step 236, a query is made as to whether the target material should be re-charged (i.e., the target material is no longer radioactive and must again receive neutrons from a neutron source to be reactivated). If the reply is YES, method 200 returns to step 212. If the reply is NO, method 200 proceeds to step 240 and ends.
The radioisotope power systems the present disclosure employ short-lived (and thus high-power density) isotopes to thermally heat a low-pressure gas (i.e., the temperature is high enough to ionize this gas, and is a plasma upon exit). The radioisotope is produced by capture of thermal neutrons in a target material resulting in activation. A low-pressure propellant reduces the heat flux to the walls and engine components. Therefore, unlike a nuclear thermal rocket, the melting point of the materials is not a limiting factor in the deployment of the radioisotope power systems described above. In some cases, a magnetic nozzle may still be required.
The power density and thrust duration of the disclosed radioisotope rocket thruster systems may be adjusted based on the application. For example, phosphorus or yttrium may be used for high thrust, shorter burn durations (14-30 days), and scandium may be used for longer duration, lower thrust applications. Furthermore, target materials may be mixed in one engine where high thrust is needed early, and low thrust later (for initially climbing out of a gravity well, and then slowly accelerating afterwards). The thrust and Isp may also be varied by using multi-mode injection schemes, or by using hypergolic propellants such as hydrazine or ASCENT (Advanced Spacecraft Energetic Non-Toxic) propellant in combination with a catalyst bed (Note: ASCENT, formerly known as AF-M315E, is an advanced monopropellant formulation developed by the Air Force Research Laboratory (AFRL) Rocket Propulsion Division (RQR)). For instance, an element like manganese may be advantageous for laboratory-based, proof of concept experiments. Further, manganese may also be used in very short high-thrust missions. As an example, 55Mn activates into 56Mn, which has a 2.57 hour half-life with a 2.8 MeV beta, which means a more modest neutron source (such as a particle accelerator at a university) may activate a small amount and produce a significant thrust. In this case, with 1 hour of irradiation of 10 g of Mn at 1015 n/cm2/s produces about 120 watts initially, equates to 12 mN of thrust (for Isp=2000 seconds). This short operating life is advantageous for, for example, laboratory settings because the radioisotope does not remain highly radioactive for very long, thus improving laboratory safety considerations.
A great deal of neutron activation is possible with a relatively simple reactor design based on the concepts described herein, and calculations show that achievable neutron fluxes (<1015 n/cm2/s) in existing reactors may produce an initial decay power in phosphorus greater than 1 kW/kg. However, even at lower neutron fluxes, good specific power may be achieved using the concepts described herein. The quantity of activated isotopes reaches an equilibrium after some length of time, which is a function of half-life and neutron flux. For phosphorus, the equilibrium is reached in about 60 days, and the asymptotic limit is higher with higher neutron fluxes. If a lower specific power is acceptable, then irradiation time may be reduced, and this helps to extend reactor life.
Shown below is a list of candidate propellants, such as propellant 140. Other propellants 140 are possible, as known to those skilled in the art.
Shown below is a table of candidate target materials that may serve as radiation sources, such as target material 11 of
It is noted that the decay energy from the activated target material may be used to increase the enthalpy of a propellant to produce heat to be converted to thrust. Alternatively, the energy from the radioactive decay may also be used to accelerate propellant to a high exhaust velocity by some other means than heat. For instance, electricity matched to the requirements of the thruster (e.g., derived from the decay heat or directly from the radiation) may be used to power an electric thruster. This heat-to-electricity or radiation-to-electricity conversion may be performed by any system that produce voltage and current output that is matched to the desired electric thruster inputs. This conversion to electricity eliminates the need for power conditioning equipment to change or regulate the voltage and current, resulting in significant reduction in the overall system mass.
Some simplified reactor models have been evaluated using the OpenMC continuous-energy particle transport reactor physics code (although other neutron transport codes such as MCNP or Serpent could also be used). The reactor uses high-assay low-enriched (HALEU) fuel for lower cost and regulatory restrictions.
In particular, a simplified model engine was included in the OpenMC reactor model to determine the neutron capture rate with phosphorus-31 target material. This OpenMC model is simulating the fission in the reactor as well as neutron scattering and thermalization; thus, there are no assumptions about neutron energy or flux being made to estimate the activation. The target material included a reduced density to approximate a void in the engine. The volume-averaged neutron flux was also measured in the reactor fuel and in the engine. To determine the specific power of the engine, the number of transmuted target atoms must be determined. The average neutron flux may be used to independently verify the neutron capture rate in the target material by solving the system of ordinary differential equations for capture and decay:
NP31 and NP32 are the numbers of the respective phosphorus isotopes, φ is the neutron flux (n/cm2/s) (from OpenMC tally result), σP31 is the thermal neutron capture cross section for 31P, and λP32 is the decay constant of 32P. Additionally, 32P does not have an appreciable capture cross section even at thermal energies, so only 31P and 32P are considered in this simple calculation.
After solving for NP31 and NP32, the decay power may be calculated as p=Ap32 N32 Ep, where E is the energy of the emitted beta particle (1.7 MeV for 32P). For a neutron flux of 1015 n/cm2/s, and an irradiation time of 60 days (when the 32P population reaches equilibrium) and 5 kg of 31P the population of 32P atoms is 2*1022, which is an initial decay power of 4.3 kW, and a specific power of 0.863 kW/kg. This same calculation may be done using the NIST Activation Calculator (https://www.ncnr.nist.gov/resources/activation/.)
Other reactor geometries, such as containing 500 kg or more or phosphorus as target material, are contemplated. In such larger reactor configurations, the solid fuel may not be easily replaced after being consumed. Therefore, in some embodiments, a molten fuel may be used in a manner similar to a terrestrial molten salt reactor. This configuration allows for a larger fuel load that may be stored outside the reactor and circulates through the core, thus extending the life of the reactor and makes it easier to potentially replace the fuel. Such a configuration would also allow many more engines to be charged during the life of the reactor, making it more economical to implement. As a variation, a pebble-bed core would also allow potential online refueling as well as better heat transfer. The flux trap may be further improved using an ultra-cold moderator surrounding the cavity, such as frozen deuterium (at just a few kelvin). This will greatly increase the neutron capture in the engine target by lowering the neutron thermal energy.
An alternative configuration of a reactor geometry is shown in
As shown in
Further possibilities for materials suitable for use as a target material in the aforedescribed space-borne radioisotope power system are listed in the table shown in
The disclosed radioisotope power system may serve a whole class of thrusters and power sources because there are so many possible engine configurations that benefit from space-based in-situ isotope production. For instance, space-based in-situ isotope production could activate a thin bismuth coating on a collapsed isotope sail to produce Po-210, which emits alpha particles from a thin plastic sheet to produce thrust with extremely high Isp. See, for example, R. L. Forward, “Radioisotope sails for deep space propulsion and electrical power,” 31st Joint Propulsion Conference and Exhibit, San Diego, Calif., 1995, which publication is incorporated by reference in entirety for all purposes.
An activation rocket may have a mix of target material that activates into both short and long-lived isotopes to have high initial thrust and then lower, longer lived thrust. In addition to being ideal for orbital transfer vehicles, the disclosed radioisotope power system may be ideal for interplanetary probes because, in addition to activating thrusters, radioisotope batteries may also be charged in the same reactor. The charging station reactor (such as shown in
There are myriad different propulsion and power systems that could benefit from charging in space. Such an arrangement could eliminate many of the logistical and safety issues involved with launching radioisotope systems from the ground, while also dramatically increasing performance of those systems due to the use of shorter-lived isotopes. Having the orbital infrastructure for servicing spacecraft with these high-performance thrusters and power systems enables new capabilities heretofore uncontemplated.
Furthermore, the disclosed radioisotope power system may be placed in multiple orbits for orbital transfer vehicles to make round trips between high and low orbits. For example, the reactors/charging stations may be placed in multiple orbits to service orbital transfer vehicles (OTVs) that contain the thrusters. There would also need to be reactors (or other neutrons sources) placed to re-charge the thrusters in those various orbits (See, e.g., U.S. patent application Ser. No. 16/986,517 to Clark, as cited above). They could also be placed at more distant destinations such as Mars or the asteroid belt, where they could be used for charging engines for return trips to Earth, or trips further out in the solar system. This system allows large and small spacecraft to move nearly anywhere in the solar system. It is conceivable to have large thrusters that produce hundreds of newtons of thrust and are capable of moving bulk cargo and people.
Any number of radioisotope power systems may be placed in any number of activation station orbits to service visiting spacecraft. For example, there could be several at different inclinations and altitudes to re-charge orbital transfer vehicles. Such a scenario is illustrated in
The orbital transfer vehicle may also include isotope propulsion systems that can be activated and/or recharged by the radioisotope power system at the charging station. Additionally, the orbital transfer vehicle may be used for moving the charging station itself to new locations. For instance, the orbital transfer vehicle may be used to change the activation station orbit of the radioisotope power system or to move the charging station to a desired location, such as to a planet or a space station, as shown in
A neutron-flux trap type reactor in orbit may be used to irradiate other types of items (such as other types of thrusters or radioisotope electrical power sources) by neutron capture or other secondary reactions to produce radioisotopes within the item. For example, U.S. Provisional Patent Application No. 63,184,138, filed Apr. 5, 2021 and entitled “RADIOISOTOPE ROCKET THRUSTER SYSTEM AND METHOD OF USE,” which application is incorporated herein in its entirety by reference, describes a neutron-flux trap type reactor suitable for integration with the radioisotope rocket thruster system described herein.
Listed below are some of the possible propulsion systems that may utilize the space-borne flux-trap reactor for activation:
For all of the thruster concepts listed in above, any mix of isotopes may be used to tailor the thrust, specific impulse, and/or burn time for a given mission or application. For instance, the exponential decay of power may be leveraged as an exponentially decreasing thrust with constant Isp, constant mass flow rate with decreasing thrust and Isp, or some combination thereof. Similarly, for isotope power sources, longer lived (i.e., lower power) isotopes may be mixed with shorter lived (i.e., higher power) isotopes to provide more power for a short duration early, and lower power for longer.
The isotope may also be incorporated into thruster and/or battery components to help ionize, heat, or otherwise condition the propellant before being used in the thruster.
Furthermore, it could be possible to combine an isotope power source (like a radioisotope thermoelectric generator (RTG)) into one device (or separate, but physically adjacent to fit within the reactor flux trap), each having their own isotope(s). This would help enable deep space missions by providing propulsion and power by charging in the same reactor.
There are a variety of isotope power sources that may be activated by an orbiting reactor as described herein. Some types of these isotope batteries are listed below.
Note that because the propellant entering the chamber is kept at low mass-density, the stopping distance for emitted beta particles is quite long, which necessitates deflecting the beta particles into spiral paths with magnetic fields. Either axial or radial magnetic fields may be used to this effect. To help prevent the negatively charged betas from impacting the target material and getting re-captured, electrostatic fields are established between the target (or the cladding) and the walls or other components to deflect the orbits away. The orbits eventually migrate towards the anode(s), but this may be delayed long enough for the betas to transfer their kinetic energy to the propellant. Any number of combinations of electric and magnetic fields may be used to deflect the particles long enough to maximize their capture in the propellant. Additionally, nanoparticles, liquid droplets, or dense gasses (such as Xenon) may be added to the propellant to help stop the beta particles and transfer energy to the propellant.
As an example, for the 1.7 MeV beta particles emitted by phosphorus-32 (the activation product of natural phosphorus-31), the magnetic field strength for a 10 cm gyro radius is only 0.015 T. Analysis may determine an optimal geometric configuration of target material and magnetic and electric fields; particle transport simulations show that engines with void fractions up to 90% have power densities above 1 kW per kg of target phosphorus. It should be noted that although denser propellants (such as Xenon) are better at stopping the betas, almost any gas may be used. Solid propellants may be used by sublimating them at a controlled rate. Any molecular gas may be decomposed into its constituent elements due to the extreme temperatures, so a compound that is easily storable as a liquid or solid could be used (such as ammonia). Additionally, a propellant with the target material itself as part of the mix may allow the propellant to self-heat (although magnetic capture of the betas would still be necessary). Adding a powdered target material to the propellant may also help with beta heating or allowing the electrically charged powder to become trapped in the magnetic field as well.
Although the present disclosure has been provided in accordance with the implementations shown, one of ordinary skill in the art will readily recognize that there could be variations to the embodiments and those variations would be within the scope of the present disclosure. Accordingly, many modifications may be made by one of ordinary skill in the art without departing from the scope of the appended claims.
This application is related to U.S. Provisional Patent Application Ser. No. 63/184,138, filed May 3, 2021, and U.S. Provisional Patent Application Ser. No. 63/246,712, filed Sep. 21, 2021, the content of which applications are incorporated herein by reference in its entirety.
| Number | Date | Country | |
|---|---|---|---|
| 63184138 | May 2021 | US |
