WO2020033009A9

WO2020033009A9 - Systems and methods for microwave electrothermal propulsion

Info

Publication number: WO2020033009A9
Application number: PCT/US2019/022621
Authority: WO
Inventors: Mikhail KOKORICH
Original assignee: Space Apprentices Enterprises, Inc.
Priority date: 2018-03-15
Filing date: 2019-03-15
Publication date: 2020-05-07
Also published as: WO2020033009A3; WO2020033009A2

Abstract

Certain improvements in microwave electrothermal (MET) propulsion systems comprise modifications to the internal structure of the resonant chamber and optimization of microwave power coupling into the chamber. Various embodiments disclosed herein improve upon the thermal efficiency of the chamber by the use of certain dielectric and metal inserts strategically placed within the chamber. Other improvements relate to the use of waste heat from the microwave generator or cavity to raise the enthalpy of a propellant, the configuration of the propellant tank for radiative shielding, the optimization of propellant injection, as well as methods of injecting energy into the diverging portion of the nozzle.

Description

SYSTEMS AND METHODS FOR MICROWAVE

ELECTROTHERMAL PROPULSION

CROSS-REFERENCE TO RELATED APPLICATION

[0001] This application is based upon and claims the benefit of priority from United States Provisional Patent Application No. 62/643,709 filed on March 15, 2018, the entire content of which is hereby incorporated herein by reference in its entirety and made a part of this specification.

TECHNICAL FIELD

[0002] The present disclosure relates to methods, apparatus, and techniques for improved microwave electrothermal propulsion systems for spacecraft

BACKGROUND

[0003] Microwave electrothermal thrusters (METs) are a method of electrical propulsion which use a microwave power source to generate and heat a plasma contained within a resonant chamber or resonant cavity. The microwave energy is provided to the chamber typically by means of a probe (sometimes referred to as an“antenna”), or, in some embodiments, a waveguide. This heating is used to raise the temperature of a gaseous propellant and provide thrust in the form of hot gases exiting the chamber through a nozzle.

[0004] Existing forms of microwave electrothermal thrusters have been tested in research universities. These prototypes reveal inefficiencies in the transfer of microwave power to the resonant cavity, the loss of thermal energy to the nozzle of the system, as well as issues with low-power efficiency. Optimization of the resonant cavity of the thruster may yield large efficiency gains at low power. Similarly, the recovery of waste energy from the microwave generator can improve efficiency. The operating temperature of the thruster has been limited by the use of low temperature materials.

SUMMARY

[0005] For purposes of this summary, certain aspects, advantages, and novel features of the invention are described herein. It is to be understood that not necessarily all such advantages may be achieved in accordance with any particular embodiment of the invention. Thus, for example, those skilled in the art will recognize that the invention may be embodied or carried out in a manner that achieves one advantage or group of advantages as taught herein without necessarily achieving other advantages as may be taught or suggested herein.

[0006] The following embodiments are various methods and configurations of a microwave electrothermal thruster system. Various embodiments disclosed herein improve upon the thermal efficiency of the chamber. This is accomplished at least in part by the use of certain dielectric and metal inserts strategically placed within the chamber. In one embodiment, the chamber is divided into two separate cavities, top and bottom, by a horizontally placed dielectric chamber divider.

[0007] The dielectric material used would necessarily have a high relative permittivity and low dielectric loss, as well as high melting points: materials such as beryllium oxide, alumina, sapphire, diamond, zirconia, quartz, hexagonal boron nitride, or yttrium are good examples. Metal inserts would similarly be moderate-to-high electrical conductivity metals with high melting points (stainless steel, nickel, nickel-based superalloys, tungsten or other refractory metals). The materials used therein must be selected carefully to be compatible with gaseous propellants. The use of water vapor necessitates oxidation resistant materials such as sapphire and nickel-based superalloys.

[0008] Other embodiments provide for efficiencies with respect to other thruster or propulsion system parameters, such as, coupling efficiency (the efficiency by which microwave energy is“coupled” into the resonant chamber), nozzle efficiency, electron loss, heating efficiency, amongst others. The combination of these configurations improves the usefulness of a microwave electrothermal thruster as a space propulsion device.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] Figure 1 is a diagram of a microwave electrothermal thruster that utilizes different geometries of a dielectric to confine a generated plasma or to shape the electric field.

[0010] Figure 2 is a diagram of the usage of metallic inserts in a microwave electrothermal thruster to confine a plasma or shape the electric field (in conjunction with dielectric material).

[0011] Figure 3 is a diagram of the usage of cermet inserts in a microwave

electrothermal thruster to confine a plasma or shape the electric field.

[0012] Figure 4 is a diagram of the usage of multiple probe couplers to excite the electric field within a microwave electrothermal thruster cavity. [0013] Figure 5 is a diagram of the usage of multiple loop couplers to excite the magnetic field within a microwave electrothermal thruster cavity.

[0014] Figure 6 is a diagram of the usage of multiple probe couplers with dual frequencies.

[0015] Figure 7 is a diagram of the usage of multiple loop couplers with dual frequencies.

[0016] Figure 8 is a diagram of the usage of an aperture coupler with a waveguide step transition.

[0017] Figure 9 is a diagram of the usage of a pin probe coupler to excite the electric field within a microwave electrothermal thruster cavity.

[0018] Figure 10 is a diagram of the usage of a metal insert with a DC bias to influence the motion of electrons within a microwave electrothermal thruster cavity.

[0019] Figure 11 is a diagram of the usage of a nozzle with a DC bias to influence the motion of electrons within a microwave electrothermal thruster cavity.

[0020] Figure 12 is a diagram of the usage of a permanent magnet to influence the motion of electrons within a microwave electrothermal thruster cavity.

[0020.1] Figure 13 is a diagram of the usage of two cylindrical resonant cavities configured to insert power into the diverging portion of a converging-diverging nozzle.

[0021] Figure 14 is a diagram of the usage of a heat exchanger and a Peltier cooler to impart heat to a propellant.

[0022] Figure 15 is a diagram of the usage of a propellant tank to shield spacecraft components from in- space radiation.

[0023] Figure 16 is a perspective view of a nozzle plate of a resonant chamber or cavity manufactured in accordance with one embodiment.

[0024] Figure 17 is a perspective view of a bottom plate of a resonant chamber or cavity manufactured in accordance with one embodiment.

[0025] Figure 18 is perspective view of a top plate of a resonant chamber or cavity manufactured in accordance with one embodiment.

[0026] Figure 19 is a side view of an MET thruster manufactured in accordance with one embodiment and shown ready to be mounted on a spacecraft

[0027] Figure 20 is a cross-sectional view of the MET thruster of Figure 19 illustrating a few of the details of a resonant chamber or cavity manufactured in accordance with the principles of the present disclosure, at least with respect to one embodiment. [0028] Figure 21 is a cross-sectional view of a resonant chamber manufactured in accordance with one embodiment of the present application

[0029] Figure 22 is an exploded view of a resonant chamber or cavity manufactured in accordance with the embodiment illustrated in Figure 21.

[0030] Figures 23-25 are graphs illustrating various simulations of the intensity of an electromagnetic field generated within a resonant chamber in accordance with various embodiments of the present application, and Figures 25A-25C show an example metal ring insert for a resonant chamber.

[0031] Figure 26 is a graph illustrating the how the efficiency of a thruster increases with specific power.

DETAILED DESCRIPTION

[0032] Systems, methods, and processes, which represent various embodiments will now be described with reference to the drawings. Variations to the systems, methods, and processes which represent other embodiments will also be described. Thus, the present invention is not limited by the type of environment in which the systems, methods, and processes are used, meaning that they can be utilized in connection with various types of spacecraft operating at various orbits or altitudes, as well as terrestrial systems.

Overview

[0033] The microwave electrothermal thruster system improvements comprise modifications to the internal structure of the resonant chamber and optimization of microwave power coupling into the chamber. Although the terms chamber and cavity are used interchangeably throughout this disclosure, persons of ordinary skill will readily understand that the principles disclosed herein would apply to either type of resonant receptacle, if any differences between a chamber and cavity are perceived. Furthermore, in some embodiments, the term“cavity” may refer to a portion of a“chamber” which is formed therein. Thus, various embodiments of resonant chambers are disclosed, wherein several have two or more cavities formed within the chamber.

[0034] Other improvements relate to the use of waste heat from the microwave generator or cavity to raise the enthalpy of a propellant, the configuration of the propellant tank for radiative shielding, the optimization of propellant injection, as well as methods of injecting energy into the diverging portion of the nozzle.

[0035] In the various figures it will be noted that the resonant chamber (as well as the MET thruster itself) are oriented in a vertical position, meaning that the vertical axis is oriented with respect to the top and bottom of the page. However, those or ordinary skill will recognize that the principles described herein will also apply to chambers and cavities of other orientations. In fact, certain advantages can be achieved by orienting certain components transversely to the vertical axis of the chamber, and certain embodiments illustrate such transverse orientation. Therefore, references to“top” or “bottom” or“upper” or“lower,” or other terms of orientation used herein, should not be considered as limiting. Furthermore, although the preferred embodiments illustrate a resonant chamber which is cylindrical, other configurations are possible.

Section 1 - Thermal Efficiency Improvements

[0036] In order to achieve the present improvements in thermal efficiency in METs, certain embodiments will now be described in detail.

[0037] Embodiment 1 : The use of a large dielectric insert in the body of the resonant cavity. Increasing the volume of the insert reduces the size of the region in which the plasma can form. By decreasing the plasma region, more power can be absorbed in a smaller region and raises the temperature of the propellant in this region. This improves the thermal efficiency of the thruster. Figure 1 shows a possible configuration where the top half of the cavity is almost completely filled with a dielectric, reducing the volume in which the propellant can flow. This thereby reduces the volume of the plasma.

[0038] In this embodiment, as shown in Figure 1, the thruster body 100 is comprised of a resonant chamber 102 which is divided into a lower cavity 104 and a reduced volume upper cavity 106 which confines the plasma discharge 108. Thrust is generated by the discharge of hot gasses through the nozzle 109. Microwave energy is coupled to the chamber by the probe 110 (probe coupler). In this embodiment, a horizontal dielectric material 112 divides the chamber at about mid-plane into the lower cavity 104 and the upper cavity 106. The large dielectric material 114 mentioned above is shown in Figure 1 just above and adjacent to the horizontal dielectric material 112. This material 112 and additional, optional dielectric blocks 116, 118 are described below in connection with Embodiment 2.

[0039] Embodiment 2: The system of Figure 1 may be optimized by the use of a precisely shaped dielectric inserts 116, 118 in the body of the resonant cavity. By placing a high permittivity dielectric 114, 116, 118 near regions of high field intensity, the electric field intensity in the region around the dielectric is increased. The plasma temperature will rise in this region, and thus the propellant temperature will also rise. This improves the thermal efficiency of the thruster. Figure 1 shows a possible configuration, where the dielectric materials 112, 114, 116, and 118 are located both on the midplane and near the top of the central axis. It should be noted that, given a preferred cylindrical chamber 102, and given the cross-sectional view of Figure 1, the inserts 116, 118 may be circular in shape to confine the plasma discharge region 108.

[0040] Embodiment 3: A metal insert in the body of the resonant cavity. Increasing the volume of the insert reduces the size of the region in which the plasma can form. By a similar mechanism to Embodiment 2, this will increase the propellant temperature. This improves the thermal efficiency of the thruster. Figure 2 demonstrates two metal inserts 120, 122 protruding into the cylindrical cavity in which around a region of higher field intensity will form. As shown in Figure 2, the metal inserts 120, 122 may be embedded in the dielectric blocks 116, 118. The metallic inserts 116, 118 may comprise a ring-like structure. In fact, the inserts may comprise a metal ring having a circular shape. The ring may have a blunt profile, as shown in Figure 2, or may have a sharp edge protruding into the chamber 102, as shown in Figure 25. The ring may also be comprised of a series of sharp metal points arranged closely adjacent to one another so as to essentially form a ring.

[0041] Embodiment 4: the use of a metal insert in the body of the resonant cavity. By precisely shaping and locating the metal insert, the electric field intensity around the insert is increased. By a similar mechanism to Embodiment 2, this will increase the plasma temperature. This improves the thermal efficiency of the thruster. Figure 2 demonstrate two metal inserts 120, 122 near a region of high field strength (the top of the central axis of the cavity).

[0042] Embodiment 5 : the use of a metal and dielectric insert in the body of the resonant cavity. By placing a metal insert near regions of high field intensity, and covering the metal insert with a dielectric, the electric field intensity in this region is increased. By a similar mechanism to Embodiment 2, the plasma temperature will rise. This improves the thermal efficiency of the thruster. Figure 2 shows a combination of dielectric and metal inserts in these regions of high field strength.

[0043] Embodiment 6: the use of a cermet insert in the body of the resonant cavity. By placing a cermet insert near regions of high field intensity, the electric field in this region is increased. By a similar mechanism to Embodiment 2, the plasma temperature will rise. This improves the thermal efficiency of the thruster. Figure 3 demonstrates a cermet insert being used in these regions of high field strength. [0044] Embodiment 7 : the use of a ceramic insert near the nozzle throat of the thruster. By placing a ceramic near the throat of the nozzle, the electric field intensity in this region will be increased. Simultaneously, the temperature handling properties of ceramics are superior to metals. This will improve the thermal efficiency and power transmission efficiency of the system.

Section 2 - Microwave Efficiency Improvements

[0045] Embodiment 8: the use of multiple coaxial probes to excite the resonant modes of the resonant cavity. By using multiple coaxial probes, the resonant cavity can receive power without the loss associated with a power combiner. This increases the power transmission efficiency of the system. Figure 4 shows one possible configuration of multiple coaxial probes.

[0046] Embodiment 9: the use of multiple loop couplers to excite the resonant modes of the resonant cavity. By using multiple loop couplers, the resonant cavity can receive power without the loss associated with a power combiner. This increases the power transmission efficiency of the system. Figure 5 shows one possible configuration of multiple loop couplers.

[0047] Embodiment 10: the use of a combination of loop and coaxial coupling mechanisms. By using the combination of couplers, the resonant cavity can receive power without the loss associated with a power combiner. This increases the power transmission efficiency of the system.

[0048] Embodiment 11 : the use of multiple coaxial probes with two separate frequencies. One frequency excites the resonant mode of the cavity. The other frequency couples a large amount of power into the cavity at a lower frequency. Lower frequency microwave generators are more efficient. This increases the microwave efficiency of the system. Figure 6 shows one possible configuration with the low frequency signal applied to the on-axis probe and the high frequency signal applied to the probe at the midplane. The low frequency signal could be at a different resonant mode of the cavity, or only applied once the discharge is ignited and the quality factor of the resonant cavity is low. This embodiment is not limited to a set number of coaxial probes, or only two frequencies, or to a particular configuration of low and high frequency assigned to a particular coupler.

[0049] Embodiment 12: the use of multiple loop couplers with two separate

frequencies. One frequency excites the resonant mode of the cavity. The other frequency couples a large amount of power into the cavity at a lower frequency. Lower frequency microwave generators are more efficient. This increases the microwave efficiency of the system. Figure 7 shows one possible configuration with the low frequency signal applied to the on- axis loop coupler and the high frequency signal applied to the probe at the midplane.

[0050] Embodiment 13: the use of a waveguide step transformer and aperture coupling. The waveguide step transformer takes a rectangular waveguide and gradually converts the mode into a useful one for exciting the resonant frequency of a cylindrical cavity. The aperture coupling provides a method for coupling power to the cavity which does not require a dielectric material or a probe intruding into the cavity. This improves the microwave efficiency of the system and reduces the weight. Figure 8 shows one possible configuration where the coupling aperture is located on the axis at the base of the thruster. The coupling aperture must be located at a region of field intensity in both the resonant cavity and in the rectangular waveguide. Traditionally, this location can be precisely defined in a waveguide as one-quarter of a guide wavelength from the shorted end wall of the waveguide, but the periodic nature of field intensity inside the waveguide allows for many possible locations for this coupling aperture. To increase the field strength at the aperture, the waveguide would be half height or quarter height waveguide, or some other reduction of height. This allows for more power to be coupled into the cavity.

Additionally, the waveguide height reduction could be used in conjunction with a quarter- wave transformer to account for any mismatch in impedance between the resonant cavity impedance and the waveguide impedance.

[0051] Embodiment 14: the use of a pin coaxial probe to excite the resonant mode of the cavity. The coaxial probe will produce a stronger electric field on the axis of the thruster if the probe is sharpened to a point. This stronger field at the tip of the“pin” is mirrored by a stronger electric field near the nozzle region of the cavity. This improves the microwave efficiency of the system, as well as the thermal efficiency of the system. Figure 9 shows one possible configuration where the pin coaxial probe is located on the axis at the base of the thruster.

Section 3 - Electron Confinement Improvements

[0052] Embodiment 15: the use of a direct current voltage applied to an electrically isolated insert in the thruster. This influences free electron motion within the cavity. By precisely applying voltages in certain regions, this increases the thermal efficiency of the system. Figure 10 shows one possible configuration where a metal insert is placed near the midplane of the thruster. [0053] Embodiment 16: the use of a direct current voltage applied to an electrically isolated nozzle in the thruster. This influences free electron motion to prevent electron loss to the nozzle. This increases the thermal efficiency of the system. Figure 11 shows one possible configuration where a DC bias is applied to the entirety of the isolated nozzle.

[0054] Embodiment 17: the use of a cylindrical permanent magnet placed around the body of the thruster. This influences free electron motion to prevent electron loss, and thus increases the thermal efficiency of the system. Figure 12 shows one possible configuration where the magnet is placed around the nozzle of the thruster.

Section 4 - Power Addition in a Nozzle

[0055] Embodiment 18: the use of multiple cylindrical resonant cavities oriented such that they insert power in a diverging portion of a converging-diverging nozzle. This mechanism allows for the creation of a plasma discharge that does not suffer nozzle throat losses (either boundary layer losses or heat transfer losses) or causes nozzle erosion. Figure 13 shows one possible configuration of cylindrical cavities inserting power into a diverging portion of a nozzle.

Section 5 - Using Microwave Generator Heat

[0056] Embodiment 19: the use of a heat exchanger with the microwave generator to pre-heat a propellant before it is injected into the resonant cavity. This increases the thermal efficiency of the system.

[0057] Embodiment 20: the use of a heat exchanger, combined with a Peltier cooler, in order to use the waste heat from the microwave generator to pre-heat the propellant before it is injected into the resonant cavity. This increases the thermal efficiency of the system. Figure 14 demonstrates one possible configuration with the use of a Peltier cooler.

[0058] Embodiment 21: the use of a heat exchanger (with or without a Peltier cooler) in order to use the waste heat from a microwave generator to heat, vaporize, or sublimate a propellant. This increases the thermal efficiency of the system. Figure 14 demonstrates one possible configuration with the use of a Peltier cooler.

[0059] Embodiment 22: the use of a heat exchanger (with or without a Peltier cooler) in order to use the waste heat from both the microwave generator and the nozzle of the thruster to heat, vaporize, or sublimate a propellant. This increases the thermal efficiency of the system. [0060] Embodiment 23: the use of the resonant cavity to vaporize a propellant. In order for a solid or liquid propellant to be used in a space environment, the propellant must first be heated until it reaches a gaseous state. When the microwave electrothermal thruster system is started from a cold state, the system can provide the heat necessary to vaporize or sublimate the propellant by transmitting power to the resonant cavity without any propellant. The power will be dissipated as heat into the walls of the cavity. This decreases the need for secondary components, and thus reduces the weight of the system.

Section 6 - Radiation Shielding

[0061] Embodiment 24: the use of a dense propellant tank surrounding the direct current components of a space propulsion system, which are prone to failure from space radiation energy. The dense propellant helps shield these components from space radiation, helping increase the longevity of the system. Figure 15 shows one possible configuration of a propellant tank that helps shield direct current components.

[0062] Embodiment 25: the use of a dense propellant tank surrounding sensitive instruments - which are prone to failure and interference from radiation energy. The dense propellant helps shield these components from radiation, which will allow increased longevity and more sensitive, accurate measurements. Figure 15 shows one possible configuration of a propellant tank that helps shield sensors or sensitive instruments.

Resonant Chamber Configurations and Thruster Embodiments

[0063] One embodiment of a resonant chamber manufactured in accordance with the present principles is illustrated in Figures 16 - 18. In addition, one embodiment of a MET propulsion system configured to be integrated into a spacecraft is illustrated in Figures 19 - 20.

[0064] With respect to Figure 16, there is shown a perspective view of a nozzle plate 200 mountable to a resonant chamber (not shown) and having various apertures 202 for mounting the nozzle plate 200 to a top chamber plate or component (shown in Figure 18). The nozzle aperture 204 is where hot gasses are emitted to provide thrust. Pressure transducer port 206 can regulate the pressure/flow rate of the propellant which is injected into the muzzle region through a separate port (not shown). Propellant gas could be various well-known gases such as helium, nitrogen, hydrogen, ammonia, argon, xenon, and water vapor. It should be noted that no propellant is provided to the lower cavity such that the plasma discharge is only located in the upper cavity. [0065] Figure 17 is a perspective view of one embodiment of a bottom chamber component or plate 208 illustrating a cylindrical lower cavity 210, which comprises a portion of the resonant chamber of this embodiment.

[0066] Figure 18 is a perspective view of an upper chamber component or plate 212 illustrating a cavity 214 which comprises the upper cavity of the resonant chamber of this embodiment.

[0067] Figure 19 is a side view of an MET thruster 215 of one embodiment shown configured for mounting on a spacecraft. The thruster utilizes a resonant chamber 216 in accordance with the embodiments illustrated herein. Flanges 218 serve to mount the resonant chamber/thruster 215/216 to the spacecraft.

[0068] Figure 20 is a cross-sectional view of the thruster 215 of Figure 19 illustrating the inner components of the thruster, including the resonant chamber 216, the upper and lower cavities, 220, 222, the probe coupler 224, the horizontal dividing dielectric material 226, and the nozzle 228. In addition, the nozzle plate 200, and the upper and lower chamber plates 212, 208, described above in connection with Figures 16 - 18, are shown in this view.

[0069] Similarly, Figure 21 illustrates in greater detail a cross-sectional view of the resonant chamber 216 of this embodiment, including the same components illustrated in the cross-sectional view of Figure 20. Furthermore, Figure 22 illustrates an exploded view of the resonant chamber of Figure 21, including the same components, but also illustrating certain fasteners 230 for assembly of the resonant chamber 216 and O-rings 232 for sealing purposes. In addition, upper resonant chamber plate 212 illustrates a port 213 for propellant injection into the upper cavity 220.

[0070] Figures 23 - 25 are simulation graphs which illustrate the intensity of the electromagnetic field generated by the resonant chamber 216 of various embodiments disclosed herein. These graphs illustrate that the field can vary in intensity by vertical and circumferential location within the upper and lower cavities 220, 222. More importantly, these graphs illustrate how the strategic placement of certain dielectric and metallic inserts in the resonant chamber can significantly increase such intensity, by as much as 50% or even 100% over baseline.

[0071] Figure 23 illustrates the baseline intensity of a resonant chamber 216 for purposes of these illustrations, including an upper cavity 220 and a lower cavity 222 divided by a horizontal dividing dielectric material 226. The probe 224 is shown in the lower cavity region 222. The darker, black regions of this image represent the colder regions within the chamber (temperature being a function of intensity), whereas the grey scale regions 233, 234 represent the regions of greater electromagnetic intensity and higher temperatures. It should be noted that a field of high intensity 234 is shown in the upper cavity near the nozzle of the thruster.

[0072] Figure 24 is a simulation of a similar resonant chamber 216; however, in this embodiment, a ring 235 of dielectric material is placed in the top of the upper cavity 214 of the resonant chamber 218, whereas in the simulation of Figure 23 has no such insert.

In this case, the field intensity 236 increases by a significant amount (50% or so) next to that region compared to the baseline of Figure 23. In Figure 25, the same sized resonant chamber 216 is illustrated, except in this simulation, a metal ring 237, such as a sharp metal insert, is embedded into the dielectric material 235. In this embodiment, the intensity in the upper region 238 is a 100% improvement in electric field maximum intensity as compared to baseline. The metal insert has a significantly higher maximum field at a point right below the metal insert (vs the dielectric block alone). The metal ring insert 237 is shown in greater detail in Figures 25A - 25C, including the sharp edge 239 formed on the lower surface thereof. In another embodiment, illustrated for example in Figure 9, rather than a blunt end probe, a sharp pin probe conductor is used. This embodiment could be combined with these other embodiments to achieve additional improvements in field intensity.

[0073] In the simulation graphs of Figures 23 - 25, color would be the ideal indicator of field intensity. In these graphs, an intensity of 0-200,000 V/m is shown and an intensity greater than this magnitude shows a dark zone for any region that is over 200,000 V/m (the units of electric field magnitude).

[0074] Resonant chambers in accordance with these principles can be sized for various power versus propellant mass flow rate ratios. Thrusters of various powers and mass flow rates can be envisioned based on the embodiments described herein. Mass flow rate can be adjusted in the resonant chamber by adjustable aperture orifices or adjustable pressure valves. In addition, the following equation is used to size resonant cavities. In this equation, fr is the frequency, TMz_mnp is the mode of the cavity (TM011 in this case), mu is the permeability of free space, epsilon is the permittivity of free space, chi is the root of the Bessel function, and , a is the radius and h is the height of the cavity.

[0075] Figure 26 is a graph illustrating how the efficiency of a thruster increases with specific power. Each trendline represents a fixed power input with varying mass. As the power input into the thruster increases, the thruster efficiency increases. As power is directly related to electric field intensity, increasing the efficiency by which power is transmitted to the cavity will increase overall thruster efficiency. In this instance, a low power trendline could be about 25% of max operating power of the thruster, while a medium power trendline would represent about 50% of max power, and a high power would be 100%

Methods of Operation and Commercial Applications

[0076] The present MET base propulsion system can be utilized with 0.5-1 kW, 50-200 kg small satellites. MET propulsion performance parameters are at an optimum between the high specific impulse and low thrust of Hall effect and gridded ion engines and the high thrust but low specific impulse of resistojets. METs offer high thrust at small footprints. Because they do not require fully ionized gases, they can be more efficient at optimal mission specific impulse than electrostatic or electromagnetic thrusters.

[0077] The present thruster comprises a propulsion system that may rely on water vapor as a propellant. The prospect of using water vapor as propellant in a system with a specific impulse significantly higher than a resistojet has two very important mission and commercial implications. First, because water is inexpensive, dense, non-toxic, and easily stored, the use of water as a propellant will greatly reduce system integration costs and propulsion system footprint on high AV spacecraft. Second, because water is readily available on many planetary bodies including the lunar poles, Mars and many asteroids, it raises the possibility of In Situ Resource Utilization (ISRU) based high performance electric propulsion for both NASA missions of exploration and commercial resupply services . The present propulsion system will address Strategic Thrust (ST) and enable efficient and safe transportation into and through space.

[0078] The rapid expansion of the small satellite market has enabled many new applications for a fraction of the cost of traditional satellites: remote sensing, communications, and scientific experiments. Off-the-shelf technology, iterative development, and ride share launches have laid the groundwork for a new commercial space market. However, low-cost space missions beyond LEO remain unattainable with the current state of propulsion systems: small satellite electric propulsion is dominated by high ISP, low thrust systems. Helicon thrusters, Hall-effect thrusters, ion thrusters, ionic liquid thrusters, and FEEP thrusters all possess poor thrust characteristics at the 1 kW level. These characteristics do not permit reasonable flight times for small satellite missions.

[0079] More traditional propulsion systems fare no better: resistojets and

monopropellant (hydrazine, HAN, among others) thrusters have fundamental physical limitations with respect to their efficiency. With thermal limitations to the heating elements or chemical energy limitations, these thrusters can only achieve a moderate specific impulse. The proportion of a small spacecraft dedicated to propellant then becomes unreasonable to achieve large AV. The present propulsion system offers a low- cost propulsion solution for high AV small satellite missions. The system offers certain advantages:

1. Attain a high enough ISP such that, for 2-5 km/s deltaV, the spacecraft propellant mass fraction is less than 50%;

2. Achieve higher thrust available than with traditional electric propulsion concepts at the same power level - roughly 1 mN / kg of spacecraft (such that a 1 km/s AV

maneuver could be completed in roughly 10 days);

3. Employ dense, non-cryogenic propellants stored at roughly atmospheric pressure.

[0080] The present water-based MET propulsion system strikes a balance between efficiency and thrust. A MET uses a resonant cavity to generate a plasma from a gaseous propellant. Joule heating of the plasma then transfers energy to the bulk gas, which is transformed into kinetic energy by a converging-diverging nozzle. METs allow small spacecraft to use a variety of propellants, the most desirable being dense, chemically nonreactive ones, such as water. They do not require costly materials or rare-earth magnets and have electronically neutral exhaust. They emit little interference besides the microwave source band. The overall system efficiencies of METs are primarily dictated by the efficiency of microwave sources, which continue to improve year after year. They provide an ideal propulsion mechanism for non-traditional fuels.

[0081] The established trendline of previous MET thruster performance indicates that they tend to perform similarly to Hall-effect thrusters when their resonant frequency is in the X-band (8-12 GHz) and that higher frequency thrusters offer better specific impulses at lower powers. Gallium nitride solid state power amplifiers similarly are most efficient in this band. With similar molecular weights and specific heats, ammonia and water vapor are expected to have similar characteristics at the proposed thruster size.

[0082] The present propellant system may utilize waste heat from the thruster and microwave source to pre-heat (potentially to vaporize) water, that is kept in liquid state during flight, as well as protect the DC electronics from space radiation as the system travels through the Van Allen belts. Regenerative cooling of the thruster will allow the system to ran at higher temperatures. To thermally stress the system, verify repeatability, and accommodate weight and volume considerations for solar panels, the present propulsion system will ran at a reduced duty cycle. A reduced duty cycle tests worst-case behavior, as repeated thermal cycling will result in accelerated wear from a typical use scenario.

[0083] For small satellites, there are limited options in terms of efficient and low-cost propulsion systems. Because small satellites are launched piggyback with main payloads, launch vehicle providers have strict requirements about the categories of propulsion systems that can be placed on small satellites. The present propulsion system surmounts the propellant and pressurized tank limitations usually imposed on secondary payloads.

[0084] In the simplest sense, the present propulsion system allows orbit-raising or mission extension for small satellites to be much more achievable. Because the density of liquid propellant is high, the footprint of the satellite can be reduced, along with the launch cost. For a viable space ecosystem (including manufacturing, space resource utilization, and refueling) propulsion systems are a primary concern. The broader goal of the present propulsion system is to create an economical and efficient propulsion system that can function with commonly available space resources. Water is a recoverable space resource that does not require complex storage mechanisms in a space environment. The availability of water in NEAs (Near-Earth asteroids) allows a more sustainable fuel delivery and refueling cycle than would be available with traditional propellants. This

45 identifiable near-term necessity of water as a space resource makes a strong business case for water-based propellant systems.

[0085] The further development of water-based propulsion systems also creates a broader synergy with asteroid mining and prospecting. If fuel is recoverable from a NEA during a prospecting or mining mission, the spacecraft can be re-used. Without the lifetime limitations of hollow cathode erosion, the reusability prospects of the present propulsion system are high.

[0086] The present MET thruster is a unique plasma thruster technology in that the manufacturing complexity and cost of the system is low. Once a design is finalized, the thruster itself is manufactured using traditional machining techniques with low-cost materials. Fluid systems for water vapor are so common in commercial usage that their adaptation to the present system does not add significant cost.

[0087] Microwave sources, depending on frequency range and power level, can be a trivial portion of the product cost. Magnetrons especially offer a cost-effective, high efficiency method of generating large amounts of microwave power. We expect to offer thruster systems that exceed the thrust characteristics of Hall-effect thrusters by several times at a fraction of the cost. The low manufacturing complexity of the system would permit large production volumes and fast lead times. The absence of hollow cathodes and other high-risk, high-complexity aspects of traditional EP designs gives the present system a competitive edge in the marketplace.

[0088] The present high-thrust, small satellite propulsion system can reduce orbit raising times, improve the“reach” of small satellite missions, and lower the cost and propulsion mass. These advantages make the present propulsion system appealing to a wide variety of commercial satellite manufacturers. Missions enabled by the present propulsion system include small satellite GEO communications, time-efficient satellite de-orbiting, small satellite interplanetary or cis-lunar missions, as well as NEA prospecting and ISRU missions.

[0089] As the technological capabilities of small satellites increase, they become a more cost-effective option for government usage. The critical barrier to the usability of small satellites as an effective platform for sensing, science, and communication is the ability to have reliable, effective propulsion that is low-risk. Part of the justification for the small satellite as a spacecraft is the cost effectiveness of launching such vehicles en- masse to low orbits. This“splitting of risk” allows government organizations to reduce cost for specialized missions and to reduce lead times. Without effective propulsion systems that still allow for these small satellites as secondary payloads, they have limited lifetimes and are unable to participate in near-space missions. The usage of small, deep- space capable satellites for scientific missions, both in NASA and other governmental agencies, allows greater flexibility: more potential providers, and more opportunities to launch.

Additional Embodiments

[0090] Some disclosed embodiments include a microwave electrothermal thruster, comprising: a resonant chamber having an axis; a dielectric dividing material dividing the resonant chamber completely in a transverse direction to the axis to form a first cavity and a second cavity within the chamber; a microwave probe inserted into the first cavity to couple microwave energy into the chamber; an inlet in the second cavity to allow propellant to enter the second cavity, the propellant comprising a gaseous propellant which under microwave excitation forms a plasma in the second cavity; and a nozzle formed in the second cavity to allow the exit of hot gasses whereby a thrust is generated. The second cavity of the thruster can include a dielectric material surrounding the nozzle. The thruster can have a metallic ring embedded in the dielectric material. The metallic ring can have a sharp edge formed on one surface and adjacent the plasma discharge. The probe of the thruster can have a blunt end. The probe can have a sharp or pointed end.

The thruster’ s second cavity can have at least one additional dielectric material to reduce the volume in which the plasma forms. The additional dielectric material can

substantially occupy the second cavity. The second probe can be inserted into the resonant chamber for coupling a second source of microwave energy into the chamber. The frequency of the second probe can be greater than that of the first probe. The probe can be replaced by a microwave wave guide having an aperture in the resonant chamber. The second cavity can comprise a metal insert having a DC bias. The second cavity can comprise a nozzle having a DC bias.

[0091] A method of optimizing a resonant chamber of a microwave electrothermal rocket thruster yielding large efficiency gains at low power, can comprise the steps of: recovering waste energy from the microwave generator; reducing the temperature of the thruster allowing the use of low temperature materials; and improving one or more system efficiency parameters such as: thermal efficiency, coupling efficiency, nozzle efficiency, electron loss, and plasma volume. The microwave chamber can comprise a microwave cavity resonator, such as a TM011 resonator, wherein: a rocket nozzle

47 penetrates a wall of the resonator; and various dielectric or metal or cermet structures are incorporated within the resonator which increase the electric field intensity and plasma heating ability of the resonator in the vicinity of a rocket nozzle. The resonant chamber can include multiple input ports for microwave energy. The microwave energy at a relatively high resonant frequency can be input through one or multiple input ports to initiate a plasma formation in a gaseous propellant; and additional microwave energy at non-resonant relatively lower frequencies can be input through additional input ports to further heat the plasma in a gaseous propellant. The gaseous propellant can be injected into the resonator through injectors which are fabricated integral to and in close proximity to the rocket nozzle. The continuous bias voltage can be applied to electrically conducting structures for the purpose of improving electron confinement in the plasma to improve thermal conversion efficiency.

[0092] A resonator for a microwave electrothermal thruster can comprise: a resonator body; a rocket nozzle electrically insulated from the resonator body, wherein a continuous bias voltage is applied to the nozzle; a permanent magnet surrounding the exit cone of the rocket nozzle; and one or more microwave resonators positioned in close proximity to the expanding portion or exit cone of the rocket nozzle, wherein concentrated microwave energy from the resonators is used to add additional thermal energy to the rocket after propellant has passed through the throat of the rocket nozzle.

[0093] While certain embodiments of the invention have been described, these embodiments have been presented by way of example only, and are not intended to limit the scope of the present invention. Accordingly, the breadth and scope of the present invention should be defined in accordance with the following claims and their equivalents.

Claims

WHAT IS CLAIMED IS:

1. A microwave electrothermal thruster, comprising: a resonant chamber having an axis;

a dielectric dividing material dividing the resonant chamber completely in a transverse direction to the axis to form a first cavity and a second cavity within the chamber;

a microwave probe inserted into the first cavity to couple microwave energy into the chamber;

an inlet in the second cavity to allow propellant to enter the second cavity, the propellant comprising a gaseous propellant which under microwave excitation forms a plasma in the second cavity;

an nozzle formed in the second cavity to allow the exit of hot gasses whereby a thrust is generated.

2. The microwave electrothermal thruster of Claim 1 wherein the second cavity

comprises a dielectric material surrounding the nozzle.

3. The microwave electrothermal thruster of Claim 2 wherein a metallic ring is

embedded in the dielectric material.

4. The microwave electrothermal thruster of Claim 3 wherein the metallic ring has a sharp edge formed on one surface and adjacent the plasma discharge.

5. The microwave electrothermal thruster of Claim 1 wherein the probe has a blunt end.

6. The microwave electrothermal thruster of Claim 1 wherein the probe has a sharp or pointed end.

7. The microwave electrothermal thruster of Claim 1 wherein the second cavity

comprises at least one additional dielectric material to reduce the volume in which the plasma forms.

8. The microwave electrothermal thruster of Claim 7 wherein the additional

dielectric material substantially occupies the second cavity.

9. The microwave electrothermal thruster of Claim 1 wherein a second probe is inserted into the resonant chamber for coupling a second source of microwave energy into the chamber.

10. The microwave electrothermal thruster of Claim 9 wherein the frequency of the second probe is greater than that of the first probe.

11. The microwave electrothermal thruster of Claim 1 wherein the probe is replaced by a microwave wave guide having an aperture into the resonant chamber.

12. The microwave electrothermal thruster of Claim 1 wherein the second cavity comprises a metal insert having a DC bias.

13. The microwave electrothermal thruster of Claim 1 wherein the second cavity comprises a nozzle having a DC bias.

14. A method of optimizing a resonant chamber of a microwave electrothermal rocket thruster yielding large efficiency gains at low power, comprising the steps of: recovering waste energy from the microwave generator;

reducing the temperature of the thruster allowing the use of low temperature materials; and

improving one or more system efficiency parameters such as: thermal efficiency, coupling efficiency, nozzle efficiency, electron loss, and plasma volume.

15. The method of Claim 14 wherein the microwave chamber comprises a microwave cavity resonator, such as a TM011 resonator, wherein:

a rocket nozzle penetrates a wall of the resonator; and

various dielectric or metal or cermet structures are incorporated within the resonator which increase the electric field intensity and plasma heating ability of the resonator in the vicinity of a rocket nozzle.

16. The method of Claim 14 wherein the resonant chamber includes multiple input ports for microwave energy.

17. The method of Claim 16 wherein microwave energy at a relatively high resonant frequency is input through one or multiple input ports to initiate a plasma formation in a gaseous propellant; and additional microwave energy at non-resonant relatively lower frequencies is input through additional input ports to further heat the plasma in a gaseous propellant.

18. The method of Claim 15 wherein a gaseous propellant is injected into the resonator through injectors which are fabricated integral to and in close proximity to the rocket nozzle.

19. The method of Claim 14 wherein a continuous bias voltage is applied to electrically conducting structures for the purpose of improving electron confinement in the plasma to improve thermal conversion efficiency.

20. A resonator for a microwave electrothermal thruster, comprising:

a resonator body;

a rocket nozzle electrically insulated from the resonator body, wherein a continuous bias voltage is applied to the nozzle;

a permanent magnet surrounding the exit cone of the rocket nozzle; and one or more microwave resonators positioned in close proximity to the expanding portion or exit cone of the rocket nozzle,

wherein concentrated microwave energy from the resonators is used to add additional thermal energy to the rocket after propellant has passed through the throat of the rocket nozzle.