Abstract
This proposed system provides a method to generate electrical power for space-based orbiting satellites, probes, stations, habitations, and interplanetary missions. Electricity is generated by collecting the flow of ionized plasma in the solar system for low earth applications and in the solar wind beyond the earth's magnetosphere, then directing the plasma through a channel using the principle of magneto-hydrodynamics (MHD). The channel has conducting electrodes on two sides and a magnetic field directed orthogonally to the plasma flow direction. This results in an electrical current to power spacecraft functions such as batteries, communications, propulsion, guidance, navigation and control. This MHD generator has the potential of providing higher power generation density (e.g., watts/kg) for spacecraft than photo-voltaic panels. The design includes a control system to maintain voltage quality, regulate electromagnet power and control ion inlet scoop RF frequency and voltage in response to changing space ionized plasma conditions.
Claims
1. An electro-mechanical inlet scoop comprises a funnel shape that directs a flow of space-based ionized plasma into an opening of a MHD channel.
2. An inlet scoop of claim 1 further comprises a set of sequentially-spaced electrode wire rings around an inside surface of which gradually decrease in diameter in accordance with an angular slope change of the inlet scoop.
3. Wire rings in claim 2 further comprise a confinement and guidance of ions by applying out-of-phase RF potentials to these rings and a DC voltage gradient along a longitudinal axis of the inlet scoop.
4. An inlet scoop of claim 1 further comprises a mechanical ring located to mechanically affix it to an entrance of the MHD channel.
5. A control system comprises three control loops with regulators and a computer with software logic to maintain a voltage produced from a pair of collector electrodes, adjust a magnetic field current so spacecraft power is within an operational range, and control RF potentials and a DC voltage gradient to rings on a scoop inside surface.
6. This system of controls in claim 5 further receives feedback from a spacecraft voltage level to determine adjustments that will match a spacecraft electrical load.
7. The control system of claim 5 further has a loop within it to adjust current flowing to a pair of electromagnets to control a magnetic field strength which results in regulating a produced power.
8. A second control loop within the control system of claim 5 comprises adjustments to the voltage produced by a pair of MHD electrodes to maintain transient-free output voltage to a battery that stores energy to ensure that it is within an operational tolerance for a spacecraft.
9. A third control loop within the control system of claim 5 further adjusts voltages to wire electrodes that surround the inside surfaces of an inlet scoop to maintain an electromagnetic RF field and voltage gradient to guide plasma particles inwardly toward a MHD generator channel inlet.
10. An MHD channel that comprises a wedge shape with a rectangular cross-section that varies in dimension passes high-velocity ionized solar plasma through it to concentrate a plasma flow and expand it through exhaust ports.
11. The MHD channel of claim 10 further comprises a pair of conductive electrodes mounted on opposite sides of that are orthogonal to a magnetic field.
12. The MHD channel of claim 10 further comprises a pair of electromagnets constructed of a toroid of conductive copper wire wound around a circular ferro-magnetic core that are positioned halfway along a length to provide a magnetic field that is projected across at a right angle.
13. The MHD channel of claim 10 is further surrounded by a metallic box enclosure that limits magnetic field lines from projecting outwardly thereby preventing interference with exterior spacecraft RF signals and other magnetic sources.
Description
DESCRIPTION OF THE DRAWINGS
[0109] An overview of the main components of a space-based MHD generator is depicted in FIGS. 1 through 9.
[0110] FIG. 1 shows a perspective pictorial view of the main components with an inlet scoop 2 that receives the ionized particle flux (plasma) 1 from the sun and funnels it through the MHD channel inside the enclosure box 4 to convert the ionized solar plasma into DC electrical voltage and then exits the flow through exhaust port 6.
[0111] FIG. 2 shows a sectional view cut through the horizontal center plane which depicts the interior of the tapered MHD channel for the flow of ionized plasma. Circular-shaped, wire-wound electromagnets 7a and 7b are mounted on opposite sides of MHD channel to provide the magnetic field that creates DC voltage flowing over electrodes 9.
[0112] FIG. 3 shows a sectional view cut through the vertical center plane depicting the interior of the MHD channel. Wires 10 connect between the electrodes on the top and bottom surfaces. Anti-magnetic enclosure 4 surrounds the channel.
[0113] FIG. 4 shows a schematic diagram of the elements of the MHD generating system starting with the entry of particle flux (plasma) 1 into scoop 2 and flowing through channel 8 and then exiting via the exhaust port 6. Channel 8 has electrodes 9a and 9b on the top and bottom surfaces with magnets 7a and 7b rotationally spaced 90 degrees from the electrodes.
[0114] FIG. 5 shows a diagram of the central computer function that will control and monitor MHD generation functions.
[0115] FIG. 6 is a diagram of the bridge current control circuit for the electromagnets.
[0116] FIG. 7 shows a schematic of the inlet scoop with interior, circumferential metallic bands that are connected to a series of resistor and capacitor components spaced at 90 degree angles apart.
[0117] FIG. 8 is a diagram of the voltage regulator circuit to ensure the output matches the spacecraft load demand.
[0118] FIG. 9 depicts an overview of the software architecture that is used to receive and interpret functional data to logically regulate and control the MHD channel magnetic field and scoop RF frequency, and to manage network connectivity of the DC power to the spacecraft.
DESCRIPTION OF THE EMBODIMENT OF THE INVENTION
[0119] The form and composition of this MHD electrical power generation system for spacecraft applications is illustrated in the accompanying FIGS. 1 through 9.
[0120] FIG. 1 shows a perspective pictorial view of the main exterior components of the MHD generator which has an inlet scoop 2 that receives the ionized particle flux 1 from the sun's plasma and directs it through a channel inside enclosure 4 where the conversion from ionized solar plasma flow into DC electrical voltage occurs and then exits (5a and 5b) through exhaust ports 6. The MU-metal enclosure 4 surrounds the MHD channel to limit the projection of magnetic field lines not to exceed the confines of 4 and not interfere with spacecraft electronics, exterior RF signals or other peripheral magnetic sources. The adapter piece 3 serves to connect the shape of the scoop 2 to the channel inside enclosure 4. The scoop 2 is circumferentially wrapped with metallic bands on the inside surface to create an oscillating RF electromagnetic field and voltage gradient within which the flowing ions will be guided and accelerated into the scoop to the MHD channel, see also FIG. 7, and the voltage to the metallic bands is controlled as described in FIG. 7.
[0121] FIG. 2 depicts a cross-sectional view cut through the horizontal center plane of the MHD generator across the interior of enclosure 4 and channel 8 wherein ionized plasma flows and then exits through exhaust port 6. Circular-shaped, wire-wound electromagnets 7a and 7b are mounted on opposite sides of channel 8 to provide the magnetic field across the particle velocity that creates DC voltage flowing between electrodes 9 along the top and bottom surfaces of channel 8. The metal enclosure 4 is shown enveloping the magnets 7a and 7b next to channel 8. It has six pieces with top and bottom surfaces, left and right sides, and forward and aft plates with openings for the inlet and exhaust ports. The adapter 3 serves to connect the small opening of the scoop 2 to the mounting flange on the entrance of channel 8.
[0122] FIG. 3 depicts a cross-sectional view through the vertical center plane of the MHD generator across the interior of enclosure 4 and channel 8 wherein ionized plasma flows between electrodes 9, on the top and bottom surfaces, which are mounted orthogonally to electromagnets 7a and 7b. The metal enclosure 4 is shown enveloping the channel 8. The electrodes are wired in series with wires 10 between one another and then to the linear voltage regulator circuit shown in FIG. 8.
[0123] FIG. 4 depicts the major interconnected elements of the generation system starting with the entry of particle flux (plasma) 1 into scoop 2 and flowing through channel 8 and then exiting, 5a and 5b, via exhaust port 6. The MHD channel 8 has one electrode 9a mounted on the top and another electrode 9b mounted 180 degrees apart on the bottom, and electromagnets 7a and 7b mounted on the sides that are spaced 90 degrees rotationally from the electrodes. The electrically charged particle flux (plasma) 1 that creates a charge flow across the electrodes 9a and 9b from positive toward negative. This DC voltage that is created flows to a linear voltage regulator circuit 15 within power module 13 to control the input voltage to the battery and other spacecraft functions 16 (27 vdc is commonly used as the principle bus voltage on most spacecraft) via a connectable separation between the MHD generator system and the spacecraft bus electrical system which is used for guidance, navigation and control, instrumentations, and communication. A computer 14, with software, for the MHD system controls and monitors MHD generation functions, including three control loops for voltage regulation, power regulation, ion scoop voltage and RF signal control, and receives data from a Faraday cup (mounted separately aboard the spacecraft) for plasma measurement.
[0124] FIG. 5 is a diagram of the central computer control 14 function that receives data inputs from MHD electrode voltage measurements 17 and plasma-state conditions 18 (velocity and density) to determine outputs to the ion scoop voltage and RF signal 19 through the circumferentially-wrapped wires and electromagnet power 20.
[0125] FIG. 6 is the power regulation control circuit that is designed to control the current flow through the magnet 26 for control of the magnetic field and thus the power produced by the MHD generator. The basic circuit design of an H-bridge power stage is configured with four power Insulated Gate Bi-polar Transistors (IGBT's). The central computer software will send signals to the IGBT's 22, 23, 24 and 25 to control the current magnitude using Pulse Width Modulation (PWM) to the input of the electromagnets 26. The generator power output value is input to the computer which then, based on the difference between the output and load demand values, controls the duty-cycle of the PWM pulses which corresponds the current amplitude in the desired magnetic field strength in electromagnet coils 26 and the resulting MHD generator power output. Input power to the H-bridge circuit from the spacecraft is depicted in 21. The H-bridge circuit design with PWM gate drives allows control of the electromagnet current magnitude (and thus MHD generator power output) and provides the capability to reverse current direction and magnetic field polarity, which is applicable to changes in polarity of the space ionized plasma charged particle mix.
[0126] FIG. 7 is a diagram of the circumferentially-wrapped, electrically conductive strips or wires 27 around the inside surface of the inlet cone-shaped scoop 2. This is thus a stacked ring, radio frequency ion guide with a series of cylindrical ring electrodes. These electrode strips are interconnected by a series of small resistors 28 between adjacent wires running from the front opening to the rear exit. Placed 180 degrees away from the resistors are a series of small capacitors 29 to allow an RF signal to be impressed on adjacent wire rings. This will create a de potential voltage gradient to drive ions along the axis from the front opening to the rear neck of the funnel. Radio Frequency potentials of opposite polarity are applied on adjacent electrodes. The arrangement creates an effective potential (also called pseudo-potential) that radially confines ions inside the ion guide. The effective potential, V*, expressed in Volts, is proportional to the squared amplitude of the local RF electric field E.sub.rf This feature takes advantage of the fact that the electrode ring ion guide geometry is naturally “segmented” in the axial direction. The RF signal impressed in the ion scoop electrode rings will be in the range of 600 to 700 kHz depending on ion plasma makeup. A Faraday cup sensor will be mounted outside of the spacecraft to monitor plasma density and space charge. And the control computer will adjust scoop voltage gradient and RF frequency and amplitude to maximize ion scoop performance.
[0127] The scoop is constructed of a polymeric material (e.g., Kapton, polyurethane, or other fabric or laminated composite material) that is capable of withstanding the space environment. The scoop membrane 2 could be stowed into a smaller package volume in such a manner that the sequential wire ring(s) 27 concentrically surround nearby adjacent rings to form a flat pancake-like stack. This technique minimizes spatial volume when stowed on the spacecraft bus in preparation for launch. Alternatively, a stowed implementation of foldable polymeric rods could form a more compact arrangement that may deploy outwardly into a larger size scoop opening to collect more charged ions and generate more power. These could be deployed by the stowed strain energy in the folded rods or by mechanical methods (e.g. springs and hinges) that connect via incremental lengths to the scoop membrane 2. A motorized system of driven hinges could also deploy a system of separate rods that support the membrane and wires. These alternative deployment methods could be selected from depending on spacecraft interface needs that affect installation methods, electrical power that may necessitate larger or smaller scoop sizes, and the definition of individual spacecraft missions which could necessitate a specialized system installation or tailored mounting arrangement.
[0128] FIG. 8 depicts a block diagram of the linear voltage regulator circuit used to maintain a constant voltage for the spacecraft onboard power for spacecraft operations which is typically tightly controlled within less than 2 to 3%. The proposed voltage regulation control system is designed to adjust voltage 36 to the spacecraft proper level and maintain a constant DC voltage. The MHD electrode voltage in LEO conditions is expected to be 390-492 V, for geosynchronous earth orbit (GEO) and deep space MHD electrode voltages will range from 24-50 kV. Usually, spacecraft load voltage needs to be about 27 vdc. So a voltage divider circuit (e.g. a buck regulator circuit) will be used to step-down the voltage to 15-27 volts. This voltage will be used as an input to the linear voltage regulator circuit which employs negative feedback and will provide a smooth output voltage for spacecraft operations. An energy storage battery 35 will be used to store energy and to smooth out any voltage spikes, and ensure good power quality 36 to the spacecraft. The voltage input from the MHD electrodes comes in from 30. The circuit has the three resistors 31, 32 and 33 and a capacitor 33. Voltage input to 15 is 37, voltage output is 38 and connection to ground/common is 39.
[0129] FIG. 9 depicts the logic of the software architecture that is used to receive and interpret functional data, logically decide and regulate the MHD channel magnetic field and the scoop electromagnetic field, manage network connectivity and control the DC power level. Decisions are made based on the voltage and current generated in the MHD channel and received from the electrodes and flowing to the spacecraft battery to adjust up or down the inlet cone voltage and electromagnet voltages. In this diagram, software is divided into three functions: a) MHD Electromagnet Power Regulation Control 40, b) Ion Scoop RF power Control 41, and c) Voltage Output Regulator Protection for power output to the spacecraft 42.
[0130] a) Block 40 in FIG. 9 shows within it the software logic steps to control the MHD power regulation. The control of power output created by the MHD generator ensures that it matches spacecraft load demand 43. The amount of power generated by the MHD generator is dependent on space ion plasma conditions, and the spacecraft load demand that periodically changes. This power regulation control loop 44 regulates and controls the amount of power that is generated and match spacecraft demand. The MHD generator power output (see equation 1) is proportional to the square of plasma velocity, and the magnetic field strength produced by the electromagnet, and directly proportional to the plasma conductivity and distance the electrodes are separated. The magnetic field produced by the electromagnet is the variable that can be controlled and adjusted to change the amount of power that is being generated. The magnetic field strength is directly proportional to the amount of current that is circulated through the electromagnet, as shown in equation 4. Our electromagnet will be a solenoid coil with a ferromagnetic core which will have a high magnetic permeability.
B=μNI equation 4 (reference 11) [0131] where: B=Magnetic field strength (Tesla) [0132] μ=magnetic permeability (T amp/m) [0133] N=number of turns of coil [0134] I=Amps
The power regulation control is designed to control the current flow through the magnet and thus control the magnetic field and power produced by the MHD generator. The basic circuit design of an H-bridge power stage is configured with four power IGBT's as shown in FIG. 6.
[0135] b) Block 41 in FIG. 9 shows the software logic steps to adjust the RF frequency and voltage of the system of metallic strips, resistors and capacitors that surround the cone of the inlet scoop. Depending on the value of the ion density as measured by a Faraday Cup 45, the ion scoop voltage gradient 46 is adjusted up or down. If the resultant power output of the MHD channel increases, then, if the maximum frequency is reached, the RF frequency adjustment is stopped 47. If it did not increase, then the RF frequency is reduced 48. The power output is again checked and the RF frequency is adjusted accordingly 49.
[0136] c) Block 42 in FIG. 9 shows within it the software steps to logically decide and regulate power output to the spacecraft as a protection measure to prevent an over-voltage condition to the spacecraft energy storage battery. Since the DC voltage produced at the MHD channel electrodes is directly proportional to the ion particle velocity in the MHD channel, the distance between the anode and cathode, and the magnetic field strength created by the electromagnet, as shown previously in equation 1, then the voltage produced at the electrodes can vary if any of these variables change significantly. The equation for the calculation of open circuit voltage is shown in equation 5 below and is used in 50.
Voc=B×ν×δ equation 5 (reference 7) [0137] where: Voc=open circuit voltage [0138] B=Magnetic field strength of electromagnet (Tesla) [0139] ν=ion particle velocity (meters/second) [0140] δ=electrode separation distance (meters)
[0141] A constant voltage for the spacecraft onboard power maintains spacecraft operations. The onboard electronics on spacecraft operate within a fairly tight voltage regulation (<2-3%). In LEO the ion particle velocity is primarily determined by the orbital speed of the spacecraft and is not expected to vary significantly after insertion into orbit. In GEO and deep space the solar wind ion particle velocity can vary significantly. The distance between electrode plates is constant, and therefore will not cause any change in voltage. The magnetic field strength from the electromagnet will be changing due to the power regulation control circuit, which will be automatically adjusting power output to match changing spacecraft load and variations in the plasma characteristics. Because of these variations in the magnetic field strength and plasma conditions, voltages at the electrodes will vary significantly. This hardware controlled voltage regulation control system with software protection, adjusts the voltage to the proper level and maintains a constant supply.
[0142] It is to be understood from the foregoing that, while particular implementations have been illustrated and described, various modifications can be made thereto and are contemplated herein. It is also not intended that this MHD generator system be limited by the specific examples provided within the specification. While the MHD generator system has been described with reference to the aforementioned specification, the descriptions and illustrations of the preferable embodiments herein are not meant to be construed in a limiting sense. Furthermore, it shall be understood that the aspects of this MHD generator system is not limited to the specific depictions, configurations or relative proportions set forth herein which depend upon a variety of conditions and variables. Various modifications in form and detail of the space-based MHD generator system will be apparent to a person skilled in the art. It is therefore contemplated that the system shall also cover any such modifications, variations and equivalents
