Patent Application
20220333582
The present invention relates to a plasma thruster device.
Classical pulsed plasma thrusters are known as a propulsion technology e.g. for use in satellites. Although the thrust to power ratio of a pulse plasma thruster (PPT) is limited, their simplicity, reliability and often solid state (e.g. PTFE) propellant makes them attractive as a thruster for small satellites. These pulsed plasma thrusters can operate at high frequencies so an almost continuous operation of the thruster may be obtained. Such known systems typically have two main electrical circuits. The first main electrical circuit is an ignition circuit, which may for example have a capacitive circuit for storing electrical power, and a switching circuit, for releasing the electrical power and generating an electrical arc. This electrical arc ablates and ionizes a small fraction of the propellant into a low energy plasma. The second main electrical circuit generates an electrical discharge through the formed low energy plasma, thereby generating a Lorentz force due to the interaction of a magnetic field and the electric discharge current through the plasma. This Lorentz force accelerates the plasma out of the thruster. An advantage of the PPT is its simplicity in design and operations. This means it is very robust and can effectively be made very small (which is advantageous for modern miniaturized spacecraft). A disadvantage is that the thruster efficiency of a classical PPT is quite low, resulting in a relatively low thrust to power ratio. Furthermore the anode and cathode plates of the accelerator stage, as well as the anode and cathode of the igniter (e.g. spark plug) of the igniter discharge stage, suffer from erosion.
A background of an advanced pulsed plasma thruster concept is given in T. E. Markusic, Y. C. F. Thio, and J. T Cassibry, “Design of a High-Energy, Two-Stage Pulsed Plasma Thruster,” presented at 38th AIAA Joint Propulsion Conference, Indianapolis, Ind., Jul. 7-10, 2002. In the proposed structure liquefied lithium is pumped between electrodes arranged at one end of an acceleration channel, where a lithium droplet grows in size until the distance between the electrodes is completely bridged—thereby closing an electrical circuit resulting in a discharge of a high power capacitor—thereby ionizing the lithium droplet. The resulting plasma is accelerated and is jetted out of the acceleration channel. The lithium has a low density (0.53 g/cm3) and is liquefied by heating it above its melting point of 454 K. Because of the size of the droplet the generated plasma has a low velocity and low thrust power. A second electric acceleration stage is used wherein an electric circuit deposits most energy in the plasma for the accelerator stage, to increase the plasma velocity for generating thrust. This multi stage process typically lasts several microseconds to deliver a complete cycle that can be repeated. Also, in the known thruster, the electrodes for ionizing the propellant are eroded after continued use, so that lifetime is limited and structure, materials and geometry of the known plasma thrusters are quite constrained in their operational use. The development of a small and reliable thruster device is therefore desirable but needs an improvement of the system before it can be miniaturized.
In one aspect of the invention there is provided the features listed in claims 1. In particular, a plasma thruster device comprises: an electrically insulating substrate, said substrate comprising one or more feed channels for feeding an electrically conductive liquid to a bridge structure; said substrate further provided with electrical terminals;
said bridge structure configured to form, when provided with the electrically conductive liquid, an electrical conducting bridge;
said bridge structure configured to form contact areas in electrical contact with said electrical terminals, said bridge structure thereby connecting the contact areas, said bridge structure arranged for forming a plasma of said electrically conductive liquid, when the electrically conductive liquid is ionized by a current peak flow circuit that contacts the contact areas via said electrical terminals. By using a liquid as a source material for ionization the conductive path can be renewed after a current pulse. The electrical connection to the part that will be turned into plasma can be made through the lines through which the liquid is fed. This avoids any erosion to the electrodes, as the electrical contacts in direct contact with the plasma are a liquid and these will be replenished after each electrical discharge. This feature provides renewable electrodes and bridge between those electrodes to which a high voltage of several kV can be applied. The application of this voltage causes a high, sharp-peaked current of several kA through the bridge structure that rapidly heats, melts, evaporates, and turns into plasma. The plasma expands at several km/s, without the need for a dedicated accelerator stage. This single stage electrically conductive liquid propellant pulsed plasma thruster has an excellent thruster to power ratio, and can be used regeneratively in a way that will prevent wear out or malfunction due to the ionisation of its propellant. It may have a high volumetric density, so that it can be miniaturized and used e.g. in nano satellites and in some embodiments may generate a high velocity (several km/s) plasma without the need of a Lorenz force accelerator.
Embodiments of the invention will now be described, by way of example only, with reference to the accompanying schematic drawings in which corresponding reference symbols indicate corresponding parts, and in which:
Unless otherwise defined, all terms (including technical and scientific terms) used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this disclosure belongs as read in the context of the description and drawings. It will be further understood that terms, such as those defined in commonly used dictionaries, should be interpreted as having a meaning that is consistent with their meaning in the context of the relevant art and will not be interpreted in an idealized or overly formal sense unless expressly so defined herein. In some instances, detailed descriptions of well-known devices and methods may be omitted so as not to obscure the description of the present systems and methods. Terminology used for describing particular embodiments is not intended to be limiting of the invention. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. The term “and/or” includes any and all combinations of one or more of the associated listed items. It will be further understood that the terms “comprises” and/or “comprising” specify the presence of stated features but do not preclude the presence or addition of one or more other features. All publications, patent applications, patents, and other references mentioned herein are incorporated by reference in their entirety. In case of conflict, the present specification, including definitions, will control.
The “substrate” may be a ceramic substrate or any other suitable non-conductive substrate, such as silicon or silicon like substrate (e.g. pyrex). This substrate may be part of the satellite that is facing the bridge. This substrate is non-conductive, non-reactive with the conductive liquid, hard, tough, strong, and erosion resistant. A “current peak flow circuit” may be a conventional circuit suitable for activating the plasma thruster device; i.e. by ionization, i.e. plasmafication of the bridge structure. Examples are presented in
The invention pertains, in some embodiments, to the field of nano satellites, in particular CubeSats. Satellites typically have plasma thrusters in order to maintain or alter course while orbiting earth. Space propulsion systems operate on the principle of accelerating a working mass (the propellant) to a high velocity, thereby producing thrust and changing the velocity of the spacecraft. The maneuverability of a satellite is usually expressed in terms of the velocity increase (or ΔV) of the satellite that can be imparted by its propulsion system. Each type of manoeuvre requires a certain ΔV. If a satellite must perform a series of specific manoeuvres, its propulsion system must be capable of producing a certain total ΔV, which is the sum of the ΔV of each individual manoeuvre. The total ΔV that can be delivered by a propulsion system depends on the amount of propellant on board and on the efficiency with which this propellant is used to generate thrust. The ‘propellant efficiency’ is usually expressed in terms of ‘specific impulse’ (Isp), which is the total impulse that the propulsion system can deliver per unit of propellant weight (gravimetric specific impulse) or by unit of propellant volume (volumetric specific impulse). Due to the small size of CubeSats, the available propellant storage volume is limited and consequently the thruster's total impulse (or the total Delta-V capability) is also limited. The invention lies in providing direct plasmafication of the propellant of a pulse plasma thruster, omitting the need for a separate igniter and accelerator—thereby enabling the use of a high density electrically conductive liquid as a propellant. The direct plasmafication, electrically conductive liquid propellant pulse plasma thruster is capable of providing nano satellites with improved thruster efficiency, while at the same time not taking up too large a volume of the satellite.
The bridge material 12, i.e. an electrically conductive liquid, from which the high velocity plasma is formed at bridge 13 has a relatively low electrical resistance for which the total dynamics of the electrical current peak flow circuit 30 is optimized so that most of the energy of the capacitor will be put in the bridge 13 of the thruster device. For example and without limiting, in some applications a resistance around 2Ω appears to be a maximum value for the bridge resistance.
A bridge structure may be of a size as small as about 200×300×5 micrometer, but other dimensions are suitable depending on the application and the propellant used. With the value of the density and the volume of the electric conductive liquid at the bridge, the mass of the propellant can be calculated that is turned in to a plasma during each pulsing cycle. For forming a plasma, first the materials have to be heated up to the boiling point, evaporate and turn into plasma. Using the proper values for the specific heat, the enthalpy of vaporization etc. the amount of energy needed to vaporise the bridge may be calculated. Additional energy is needed to heat-up this vapour further to turn it into a high temperature plasma. The resistance of bridge 13 strongly depends on the form, thickness and length-width ratio and should be rather low, e.g. in the order of 0.1-5 Ohm.
Bridge structure 13 provides an electrical connection (bridge) between anode and cathode, and is arranged for forming a plasma when the bridge structure 13 is ionized by a current peak flow circuit e.g. provided by current peak flow circuit 30 of
In
De current of such a system can be described as:
With Uo the voltage over the capacitor
ω=√(1/LC) the circular frequency
L=the induction of the circuit and
τ=(2L/R) the time constant of the circuit.
An example of such a discharge is found in
The conductive liquid can be an ionic liquid, molten salt, liquid metal, or any other substance that can be used in liquid form and that has sufficient electrical conductivity. The liquid is either a pure substance, a mixture, or possible a fluid with suspended solid particles. Ideally the liquid has low or negligible vapour pressure, so that it does not evaporate by itself when exposed to the vacuum of space. Furthermore a melting point around room temperature is preferred, as spacecraft are in general maintained around this temperature and having the liquid at this temperature means a low amount of energy is needed to make it liquid and to keep it liquid. Finally a high density of the liquid is desirable for the space application because in small satellites the constraining parameter is usually the volume, not the mass. A high density propellant allows for a high volumetric specific impulse.
A liquid metal is preferred for the intended space application because metals in general have sufficient conductivity and high density. Examples of pure metals that are possible propellants include gallium, indium, tin, cadmium, lead, bismuth, lithium, sodium, potassium, and mercury. Alloys of these and other metals are also interesting. All these examples vary in suitability for the application due to their specific properties such as density, conductivity, reactivity, toxicity, vapour pressure, melting point, molecular mass, specific heat, surface tension, surface wetting properties, chemical compatibility with other materials, and possible other properties. Gallium and its low melting point alloys such as gallium-indium eutectic and gallium-indium-tin (“GalInStan”) are suitable. Supplying of conductive liquid towards the bridge after each discharge may be done by a number of ways. These include: 1) ‘normal’ mechanical pumping systems such as rotating pumps or positive displacement pumps, 2) pressure fed pumping by pressurizing the propellant tank of the system by some gas, 3) electromagnetic pumping, 4) magnetic forces applied via electromagnets or moving permanent magnets (if the liquid has a sufficient magnetic susceptibility, or sufficient ferromagnetic properties (e.g. because of suspended iron particles in the liquid)), or 5) capillary action (either by intrinsic affinity of the liquid to the surfaces of the system, or electrowetting on a dielectric (EWOD).
In an embodiment, bridge material may be liquid Gallium. This material is relatively non-toxic, and has a low melting point (30° C.), high density (5900 kg m−3) and favorable electrical properties. Furthermore, Gallium has a negligible vapor pressure, which prevents it from boiling off when subjected to the vacuum of space.
Thrust Generation System (TGS)
This subsystem comprises the plasma thruster device disclosed hereabove arranged to generate a small thrust, using the principle of the regenerating bridge structure. The device may comprise one or more regenerating bridge structures (e.g. in an array), and an electrical circuit containing the switch(es) and capacitor(s) (Switch Capacitor Array, or SCA). Furthermore, a heater may be needed to keep the liquid metal propellant in the liquid phase.
Propellant Feed System (PFS)
This subsystem stores the conductive liquid propellant, e.g. as shown in
Power Control System (PCS)
This subsystem contains power electronics for distribution of electrical power over the different subsystems and subassemblies and for generating a high voltage for charging the capacitor in the current peak flow circuit. The PCS may comprise of a High Voltage Power Supply (HVPS), a Low Voltage Power Control System (LVPC) and a Digital Control Unit (DCU).
The electrically conductive liquid propellant pulsed plasma thruster as disclosed herein uses a conductive liquid propellant (such as liquid Gallium), instead of an insulating solid propellant.
Since the propellant is already conductive, the electrically conductive liquid propellant pulsed plasma thruster device does not require an igniter. Therefore, the electrically conductive liquid propellant pulsed plasma thruster device generates a single discharge per pulse (instead of an ‘ignition’ discharge and a ‘main’ discharge).
The electrically conductive liquid propellant pulsed plasma thruster device uses a switch to close the electrical circuit and trigger the discharge.
The discharge in a electrically conductive liquid propellant pulsed plasma thruster device can be an order of magnitude shorter than conventional pulsed plasma thruster devices (i.e. ˜0.5 μs instead of ˜10 μs), resulting in higher discharge currents which will result in better energy coupling with the propellant.
An electrically conductive liquid propellant pulsed plasma thruster device does not have physical electrodes between which the discharge is generated. The propellant basin acts as the electrodes and regenerates after the discharge. Hence, an electrically conductive liquid propellant pulsed plasma thruster device is not susceptible to electrode erosion.
The gravimetric specific impulse is directly related to the exhaust velocity of the propulsion system:
Isp_grav=Ueff/g0 [1]
In this equation, Isp_gray is the gravimetric specific impulse [s], Ueff is the effective exhaust velocity [m s−1] and g0 is the gravitational acceleration at sea level [m s−2]. This equation shows that in order to obtain a high gravimetric specific impulse (high mass efficiency), the propulsion system should be able to accelerate its propellant to a high velocity. For electric or thermo-electric plasma propulsion system, the relationship between electric power consumption, specific impulse and thrust level is given by the following equation:
P=Isp·F·g0/2·ηt [2]
In this equation, P is the power consumption [W], Isp is the gravimetric specific impulse [s], g0 is the gravitational acceleration at sea level [m s−2] and ηt is the thruster efficiency [-], which is the ratio between the kinetic jet power of the exhaust plume and the electrical input power to the propulsion system. The thruster efficiency is the product of several sub-efficiencies that take into account the losses of the different energy conversion steps in the propulsion system. Based on the experimental data, a value of at least ηt=0.25 can be assumed, which is a conservative estimate for the thruster efficiency. For nano-satellites, it is important that the propulsion system occupies as little volume as possible, so a nanosatellite propulsion system may be optimized for maximum volumetric specific impulse, which is simply the product of the gravimetric specific impulse and the propellant density:
I
vol
=I
sp·ρp [3]
In this equation, Ivol is the volumetric specific impulse [kg s m−3] and ρp is the density of the propellant [kg m−3]. Hence, in order to obtain a high volumetric specific impulse, the propulsion system preferably operates at a high gravimetric specific impulse and/or use a propellant with a high density. The disclosed plasma thruster uses electrically conductive liquid, for example a liquid metal such as Gallium or Galinstan as propellant, which has a density that is 2.7 times higher than that of a solid propellant used in conventional plasma thruster using solid PTFE as propellant (i.e. 5900 kg m−3 compared to 2200 kg m−3). The gravimetric specific impulse of the propulsion system can be calculated with equation 1 and could be equal to 408 s for a plasma velocity of 4000 m/s. This is a conservative estimate, and may be much higher. The volumetric specific impulse of the propulsion system can be calculated with equation 2 and is the product of the gravimetric specific impulse and the propellant density. With a gravimetric specific impulse of 408 s and a propellant density of 5907 kg m−3 (density of Gallium at 1 atm. and 298.15K), the volumetric specific impulse may be about 2.4×10{circumflex over ( )}6 kg s m−3 or higher. This means that the electrically conductive liquid propellant pulsed plasma thruster could operate at a 2.7 times lower gravimetric specific impulse than a conventional plasma thruster, while having the same volumetric specific impulse and a significantly increased thrust to power ratio. As the thrust to power ratio is inversely proportional to the gravimetric specific impulse, this would result in a 2.7 times higher thrust to power ratio. The electrically conductive liquid propellant pulsed plasma thruster concept has the potential of reaching a substantially higher thrust to power ratio at the same volumetric specific impulse, or a substantially higher volumetric specific impulse at the same thrust to power ratio, than a conventional plasma thruster.
ΔV=Isp·g0·ln[m0/(m0-mp)] [4]
In this equation, m0 is the initial satellite mass including propellant [kg] and mp is the propellant mass [kg]. The total propellant mass (mp) depends on the volume that is allocated to the propulsion system and on the volumetric loading fraction of the propulsion system (i.e. the fraction of the propulsion system volume that is occupied by the propellant). Consider a hypothetical nano-satellite with the following mass and volume distribution:
In
While example embodiments were shown for systems and methods, also alternative ways may be envisaged by those skilled in the art having the benefit of the present disclosure for achieving a similar function and result. E.g. some components may be combined or split up into one or more alternative components.
For example, the above-discussion is intended to be merely illustrative of the present system and should not be construed as limiting the appended claims to any particular embodiment or group of embodiments. Thus, while the present system has been described in particular detail with reference to specific exemplary embodiments thereof, it should also be appreciated that numerous modifications and alternative embodiments may be devised by those having ordinary skill in the art without departing from the scope of the present systems and methods as set forth in the claims that follow. The specification and drawings are accordingly to be regarded in an illustrative manner and are not intended to limit the scope of the appended claims.
In interpreting the appended claims, it should be understood that the word “comprising” does not exclude the presence of other elements or acts than those listed in a given claim; the word “a” or “an” preceding an element does not exclude the presence of a plurality of such elements; any reference signs in the claims do not limit their scope; several “means” may be represented by the same or different item(s) or implemented structure or function; any of the disclosed devices or portions thereof may be combined together or separated into further portions unless specifically stated otherwise. The mere fact that certain measures are recited in mutually different claims does not indicate that a combination of these measures cannot be used to advantage.
| Number | Date | Country | Kind |
|---|---|---|---|
| 19196003.8 | Sep 2019 | EP | regional |
| Filing Document | Filing Date | Country | Kind |
|---|---|---|---|
| PCT/NL2020/050548 | 9/4/2020 | WO |
