Method and system for wireless transmission of energy in space

Abstract

A method for wireless transmission of energy in space is described herein, in which a constellation of transmitting satellites (10) orbiting a celestial body, such as the planet (P), Earth, or the moon (L), is equipped with a laser device (12) for transmitting a beam of coherent electromagnetic waves (25) to another receiving target (20), such as an orbiting satellite, a mobile vehicle, or a fixed base on the general celestial body, for transmitting energy thereto. As a result of the total transmitted power being divided among the orbiting satellites (10) in the constellation, there are fewer contraindications than in the case of a single transmitting system with equal total power. The advantage of such division among the satellite constellation is primarily due to shorter periods of solar eclipse, since sunlight is required for the operation of the transmission system. The receiving target or satellite (20) converts the beam of electromagnetic waves (25) into electrical energy, which can then be used for the desired function.

Description

[Technical Field]

[0001] The present invention relates generally to the field of wireless transmission of energy in space. [Background technology]

[0002] As is known, in space missions and activities involving orbiting satellites or stations, energy generation, storage and (internal) distribution are fundamental aspects for ensuring the proper operation of all on-board equipment and instruments, as well as for sustaining human life whenever manned vehicles are utilized.

[0003] The main source of energy is the Sun, whose radiation in space, i.e. outside the Earth's atmosphere, covers the entire electromagnetic spectrum, since it is not altered by reflection, refraction and filtering phenomena, which in turn affect solar radiation as it reaches the Earth.

[0004] Solar energy is converted into electrical energy by photovoltaic panels and is therefore available for on-board use, such as to run devices and telecommunications systems, charge batteries, etc.

[0005] For this reason, orbiting stations as well as satellites and generally all spacecraft are equipped with photovoltaic panels and associated devices for converting solar energy into electrical energy; the latter is produced as direct current but can be converted to alternating current if necessary.

[0006] This energy management for space applications has been known and established for many years and is now used reliably in most space applications, however, it suffers from a few drawbacks that reduce its effectiveness.

[0007] The first drawback is that the conversion of solar energy into electrical energy by photovoltaic panels occurs with an efficiency of 25-35%; therefore, even if the solar energy illuminating the photovoltaic panel is virtually infinite, the end result (i.e. the electrical energy obtained from the panel) is nevertheless limited by the efficiency and dimensions of the panel itself, which must be transported into space for this very purpose.

[0008] Furthermore, photovoltaic panels are known to have a limited useful lifespan, ie, a lifespan after which their efficiency decreases and energy production is no longer satisfactory.

[0009] Such lifespans are measured in years (typically 5 to 15 years, depending on the case) and imply that the panels need to be replaced after a certain period of time, which, however, is not practical in space, resulting in the decommissioning of satellites with degraded panels.

[0010] As a corollary, when high power output is required, currently employed technological solutions attempt to solve this problem by using oversized solar panel power generating systems or by utilizing nuclear power generators. The impossibility of implementing such techniques, for example on mobile ground vehicles or small orbiting probes, necessarily limits the amount of energy available and thus the performance of the spacecraft.

[0011] From the above discussion, it can be seen that, in a generally broad sense, there is an ongoing need to improve upon the most prevalent forms of energy management today, particularly when it comes to high power and / or long distances between points in space.

[0012] The technical problem underlying the present invention is therefore to meet this need; in other words, the present invention aims to provide a method for energy management in space that is able to overcome the above-mentioned drawbacks and limitations of what is currently known in the art.

[0013] The concept to solve this problem is to generate electromagnetic energy in multiple orbiting space units such as satellites, stations, probes, etc., and transmit it remotely as electromagnetic radiation to a point in space for interception by a receiver located on the orbiting satellite or on the surface of a celestial body.

[0014] According to a preferred embodiment, the energy is transmitted remotely by a laser beam, i.e., a coherent monochromatic wave in the visible and / or infrared range.

[0015] From a technical point of view, the advantage of this solution is that it makes it possible to transmit energy to regions of space that are not reached by solar radiation, which is the main source of power utilized in all current solutions.

[0016] According to a preferred embodiment, the electromagnetic radiation or laser beam is emitted by a network or constellation of orbiting satellites to provide energy transmission that is not affected by the duration of the solar eclipse.

[0017] The features of the method according to the invention are set out with particularity in the appended claims.The invention further comprises a system for transmitting energy in space, the features of which are also set out in the claims. [Brief explanation of the drawings]

[0018] Such features, as well as the effects resulting therefrom and advantages realized by the present invention, will become more apparent upon consideration of the following description of two preferred but non-limiting examples of embodiments of the invention, where reference is made to the accompanying illustrative but non-limiting drawings. [Figure 1] (a), (b) Schematic representation of respective embodiments of the method according to the invention. [Figure 2] 1(a), 1(b) is a schematic representation of a system for transmitting / receiving energy in the embodiment of FIGS. [Figure 3a]1 shows the orbital configuration of a set of satellites used to implement the method according to the invention; [Figure 3b] 1 shows the orbital configuration of a set of satellites used to implement the method according to the invention; [Figure 4a] 3(a) shows the ground track of the satellite of FIG. 3(a) and of a receiving satellite representing the transmitting target. [Figure 4b] The ground track of the satellite in FIG. 3(b) and the points on the surface representing the transmission targets are shown. [Figure 5] 3(b) is a graph showing the percentage of eclipse time for the satellite of FIG. 3(a) over a one-year period. [Figure 6] 1 is a graph showing the power received by a rectifying antenna (rectenna) used in the present invention divided according to the size of the receiver installed on the receiving satellite used as the transmission target, relative to their distance. [Figure 7] 3(b) is a graph showing the available transmission time, also called contact time, for the satellite of FIG. 3(a) in an ideal case and in a practical case. [Figure 8] 3(b) is a graph showing the available transmission time, also called contact time, for the satellite of FIG. 3(a) in an ideal case and in a practical case. [Figure 9] 4 is a graph showing the profile of the power transmitted by the satellite in an ideal case and in a practical case. [Figure 10] 4 is a graph showing the profile of the power transmitted by the satellite in an ideal case and in a practical case. [Figure 11] Shows the ground tracks of the group of moons orbiting the Moon. [Figure 12a] A possible embodiment of a receiving device 22 placed on the surface of a celestial body is shown. [Figure 12b] A possible embodiment of a receiving device 22 placed on the surface of a celestial body is shown. [Figure 13]1 illustrates the application of the invention to a target on lunar soil. DETAILED DESCRIPTION OF THE INVENTION

[0019] With reference to the figures listed above, the number 1 indicates as a whole a system for energy transmission according to the invention, comprising a number of satellites 10 orbiting a planet or celestial body P, which may be the Earth, as in the case shown in the drawings, or the Moon, Mars, etc.

[0020] For simplicity, the following description will refer to a planet (or planets) P ​​of the solar system, such as Earth, but this should not be understood as a limiting factor, and the following description also applies to other celestial bodies, such as the Moon, asteroids, etc.

[0021] The satellites 10 may or may not be equally spaced and may all be positioned in the same orbit O (thus effectively constituting a constellation or network of transmitting satellites) so that each is capable of periodically making direct contact with the target 20 and transmitting energy thereto in the form of a beam 25 of coherent electromagnetic waves.

[0022] Preferably, the gravitational orbit O of the satellite 10 is low earth orbit, ie at an altitude of several hundred kilometers (approximately 400-1,000 km) in the examples considered in the specification referring to the Earth and the Moon.

[0023] The same criteria are applicable to other planets or celestial bodies, such as the Moon or Mars.

[0024] According to the invention, the distance between the satellite 10 and the target 20 for transmitting energy to the latter is preferably less than 3,000 km.

[0025] Furthermore, the electromagnetic waves used for wireless energy transmission are preferably those of a laser beam 25 or are in any case coherent monochromatic waves having wavelengths in the visible range (i.e., approximately 400 to 700 nm) or in the near-infrared range (approximately 700 nm to 1.4 μm).

[0026] Generation of the beam of electromagnetic radiation 25 can be obtained either by electrical energy supplied by photovoltaic panels associated with each transmitting satellite 10, or preferably directly from sunlight that has been focused and processed through suitable optics and then processed by electronic means to make it coherent and monochromatic.

[0027] FIG. 2 shows diagrammatically an apparatus 12 installed on a satellite 10 for transmitting a laser beam 25 .

[0028] The device 12 preferably comprises a solid-state laser powered by an energy source 13 which, depending on the embodiment chosen, can preferably be supplied with electrical energy by using a photovoltaic system for converting the solar energy incident on the satellite 10, or an optical and / or electronic system for concentrating the solar radiation and obtaining a beam of coherent monochromatic light within the above-mentioned frequency range.

[0029] The beam 25 is remotely transmitted to a target 20 carrying a receiving structure or device 22 configured substantially as a matrix or the like, as discussed further below.

[0030] In this context, it should be pointed out that the device 12 also comprises a system (not shown in the drawings) for controlling the orientation of the beam 25 in order to keep it pointed at the target 20 during the orbital motion of the satellite 10 on which the device 12 is located.

[0031] Indeed, as shown in Figures 1(a) and 1(b), target 20 may either be a mobile vehicle or a fixed base on planet P, for example adapted to receive beam 25, or may be in orbit as a satellite or another object orbiting the Earth, the Moon or another celestial body.

[0032] In both cases, the orbital motion of the satellite 10 transmitting the energy beam 25 requires continuous beam pointing control depending on its velocity relative to the target 20.

[0033] For this purpose, the device 12 comprises electromechanical means (not shown in the drawings), such as an electric motor, a transmission (e.g., gear transmission, belt transmission, etc.), a ball joint, and any other means that may be necessary to orient the beam 25.

[0034] The electromechanical solution employed may vary depending on the type of device 12, the size and power of the laser beam 25, the speed of the satellite 10, the distance from the target 20, and other structural-functional parameters of the system 1 for wireless energy transmission.

[0035] Some of such parameters are listed below with reference to one possible application of a system for wireless energy transmission devised by the present applicant (abbreviated as ORiS).

[0036] A constellation of satellites 10 may be placed in orbit around the Earth as defined by the data contained in the table below. [Table 1]

[0037] As can be seen in Figures 5 and 6, this is an orbit at an approximate distance of 680 km from the Earth's surface that passes through the poles and covers a band on the Earth that extends evenly in both hemispheres; this results in an eclipse period of approximately 2-10% of the orbital period, i.e., a time when the satellite is not illuminated by sunlight, as shown in the graph in Figure 5.

[0038] For simplicity, it has been assumed that the energy source 13 for the laser beam is electrical energy supplied by solar panels associated with the orbiting satellite 10 .

[0039] The solar panels on each satellite 10 are sized to have a degradation factor of 0.77, losses due to incident ray tilt of 0.99, a yearly degradation factor of 3.75%, and a satellite life of 5 years. 2 Average solar intensity in LEO of , 32% solar panel efficiency, 100 W consumption to power the satellite's subsystems, 1,000 to 1,500 W consumption by Laser Apparatus 12 at 60% efficiency, and a solar eclipse duration spanning 2% of the orbital period were also assumed.

[0040] The resulting solar panel area is approximately 10 m 2 It was.

[0041] The received power as a function of distance follows the Gaussian model:

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[0042] In low Earth orbit, the constellation of satellites 10 contacts target satellites 20 in two regions of the orbit. Assuming that transmission can occur over a maximum distance of 3,000 km, it is possible to calculate the time during which power can be transmitted to the receiving satellite 20.

[0043] The ideal case is obtained when it is possible to select the satellite 10 of the constellation that is closest to the target 20 at a given time and therefore able to transmit the highest power. In the graph of Figure 7 it can be noticed that to obtain such an optimal configuration it may be necessary to change the transmitting satellite 10 many times during the face-to-face time between the constellation and the target 20, which in this example is another satellite (see Figure 1(b)).

[0044] In a practical case optimized to transmit as much power as possible, the transmitting satellite 10 is changed at most once during the facing time.

[0045] The actual case situation is shown in the graph of FIG. 8, where the actual transmission time is the sum of the highlighted segments.

[0046] This analysis has general validity as far as techniques for the selection of transmitting satellites are concerned.

[0047] In the ideal transmission case discussed earlier in the specification, the power received by the target satellite 20 was determined assuming an efficiency of the rectenna 22 of 0.8.

[0048] FIG. 9 shows the power profile for the ideal case, where it can be noticed that the average power guaranteed to the target satellite 20 over a 24-hour period is approximately 221 W, and FIG. 10 shows the power profile for the practical case, where the average power guaranteed over a 24-hour period is approximately 167 W.

[0049] Of course, this is the power transmitted by only one energy transmitting satellite 10, and therefore the total power is given by the sum of the power values ​​transmitted by each transmitting satellite 10 to the receiving satellite or target 20.

[0050] Thus, with a constellation of several tens of satellites 10, power values ​​in the range of several kW can be obtained.

[0051] For applications to the celestial body Moon P, the differences are as follows: A face-to-face time between a group in lunar orbit and a general target 20 on the planet's surface; a receiving device 22 of a structure as described below.

[0052] The lunar orbit constellation preferably utilizes one (or more) of four lunar frozen orbits that minimize the need for orbital corrections during the orbiting satellite's operational year.

[0053] In this case, for the second exemplary embodiment, a lunar base for NASA's Artemis program has been considered. Because an actual human base (lunar habitat) needs to be powered, the energy requirements of a typical target 20 vary and are estimated to reach 5 kW at night (the system would need to be significantly scaled up to provide the required energy during the day). To transmit the correct amount of energy to the lunar surface, more powerful lasers, on the order of tens of kW, are utilized. Therefore, each transmitting satellite 10 is larger and heavier. To determine the sizing of the transmitting satellite 10 and the receiver on the lunar surface, a receiver located at the same latitude and longitude as the future lunar landing site of Artemis III, i.e., -89.54° and 0°, respectively, was considered. An efficiency of η = 0.95 for conversion to direct current and η = 0.60 for laser output power conversion were also assumed. The laser beam reaches the surface of the target celestial body, which is equipped with multiple receivers that reconvert it into electrical energy. As shown diagrammatically in Figure 13, transmission is possible when the satellite is not under a solar eclipse and the angle of elevation between the laser beam 25 and the receiving device 22 is α ≥ 30°. Considering a laser output power of 12 kW, the maximum distance covered by the laser beam is 1,400 km if the minimum angle of elevation α = 30°. At such a distance, the minimum area of ​​the receiver on the lunar surface is 9.6 m 2 and the diameter is D rec = 3.5 m. It must be pointed out that such dimensions do not take into account any safety margins nor any other margins due to inaccuracies in the laser beam stabilization technique.

[0054] According to the above-described scenario of transmission to a target located on the lunar surface, fewer satellites are required than in the Earth case. As shown in FIG. 11, it is possible to cover almost the entire lunar surface using only a few satellites. The architecture presented by the present invention consists of a number n of satellites around a celestial body other than the Earth, in this particular example the Moon (L), utilizing a suitable frozen orbit and laser beam transmission and reception system similar to that previously described in the specification for the Earth transmission case, with some differences in the definition of the face time between the transmitting satellite 10 and the receiver device 20 and in the latter's implementation. Based on such data, it can be estimated that a constellation including five satellites located in a selected orbit at an altitude of approximately 600 km with an inclination angle i = 86° can transmit approximately 37,300 kWh to the lunar base located at the south pole of the Moon, where a receiver shaped as a spherical cap with a radius of 3.6 m and a height of 2.72 m is located, as shown in FIG. 12(b).

[0055] From the above description it can be seen how a method for wireless energy transmission and a system 1 implementing such a method can solve the technical problem underlying the present invention.

[0056] In fact, the deployment of a constellation of satellites 10 instead of one large orbiting station offers several improvements in the architecture of the energy transmission system, including maximizing face time with the receiving target 20 and total coverage of low Earth orbit or the lunar surface.

[0057] Furthermore, because the total transmitted power is divided among multiple satellites 10, contraindications due to the high temperatures of the laser devices 12 and the energy sources 13 required to operate them are limited compared to using a single transmission system with equal total power.

[0058] In the present invention, the transmitting satellite 10 includes a laser 12, preferably solid state, sized to transmit a monochromatic beam 25 adapted to transmit energy in the form of coherent electromagnetic radiation. At the receiving satellite 20 in orbit, or at a target (mobile vehicle or fixed base) on planet P, the monochromatic beam 25 is transformed by a suitable device 22, which converts such electromagnetic energy into electrical current for on-board use, e.g., for motors, propulsion devices, etc.

[0059] A first configuration of receiver 22 may include a matrix of optical rectifying antennas (rectennas), which are devices consisting of antennas and rectifying diodes that convert electromagnetic waves at near-infrared frequencies into direct current.

[0060] A second embodiment of the receiver element 22 comprises a matrix of photovoltaic cells designed for highly efficient conversion of a laser-generated monochromatic beam.

[0061] In the case where the invention is implemented in Earth orbit, in order to realize wireless energy transmission between two points in space, the orbit selected for the transmitting satellite 10 is of the sun-synchronous type, more particularly of the Doondusk type, which advantageously ensures the longest exposure time or the shortest eclipse time for the transmitting satellite 10 and thus allows storing as much energy as possible to be utilized for remote transmission.

[0062] In the case where the invention is implemented in lunar orbit (F), in order to realize wireless energy transmission between a point in space (sending satellite 10) and a point on the surface (target 20), the orbit selected is one of the lunar frozen orbits, i.e., by definition, a quasi-circular orbit with an inclination of 28°, 51°, 76° or 86° and an altitude in the range of 500 to 700 km. The selected orbit shall always ensure the shortest eclipse time for the sending satellite.

[0063] In this latter embodiment, the receiving structure shall have a shape suitable for minimizing cosine loss (i.e., loss due to the angle between the electromagnetic beam and the impact surface) and optimizing the minimum angle of attack at which the transmitting satellite 10 can transmit energy. Rather than a simple flat structure, which may have an efficiency of 82.7% due to cosine loss, the present specification proposes using a raised edge and a dome (i.e., spherical cap) shape, as shown schematically in Figures 12(a) and 12(b). In the future, these two configurations may be merged into one, with further optimization being pursued.

[0064] All of these features and variations are intended to fall within the scope of the following claims.

Claims

1. 1. A method for wireless transmission of energy in space, comprising: - providing a plurality of transmitting satellites orbiting a planet or celestial body and each having a device for emitting a beam of coherent electromagnetic waves; providing at least one target with a receiving device substantially constructed as a matrix of components adapted to convert said beam of coherent electromagnetic waves into electrical energy; - transmitting energy to said target by emitting a beam of coherent electromagnetic waves from said sending satellite towards said target; A method comprising:

2. The method of claim 1 , wherein the beam of coherent electromagnetic radiation is a laser beam.

3. 3. The method of claim 1, wherein the radiation of the coherent electromagnetic beam occurs when the distance of each satellite from the target is less than 3,000 km.

4. 3. The method according to claim 1 or 2, wherein the receiving device of the target substantially comprises a matrix of photovoltaic components and / or rectifying antennas (rectennas).

5. The method of claim 1 or 2, wherein the target of energy transmission comprises one of an orbiting satellite, a mobile vehicle on a planet or celestial body, or a fixed receiver on a planet or celestial body.

6. The method of claim 1 or 2, wherein the orbit of the transmitting satellite is sun-synchronous.

7. The method of claim 1 or 2, wherein the orbit of the transmitting satellite is a low lunar orbit.

8. 3. The method according to claim 1 or 2, wherein the at least one target with a receiving device has a raised edge shape or a spherical cap or dome shape.

9. 3. A system for implementing the method of claim 1 or 2, comprising a plurality of transmitting satellites orbiting a planet or celestial body, each having a device for emitting a beam of coherent electromagnetic waves.

10. The system of claim 9 , wherein the beam is a laser beam.

11. 10. The system of claim 9, wherein the transmitting satellites travel in the same orbit.

12. 10. The system of claim 9, wherein the transmitting satellites are equidistant from one another.

13. 10. The system of claim 9, comprising a receiving target carrying at least one matrix of photovoltaic components and / or rectifying antennas (rectennas) adapted to be illuminated by the transmitting satellite.

14. The system of claim 13 , wherein the receiving target has a raised edge shape or a spherical cap or dome shape.

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