GB2551013A

GB2551013A - Remotely powered propulsion system

Info

Publication number: GB2551013A
Application number: GB1705307.5A
Authority: GB
Inventors: Quaw M'dimoir
Original assignee: Individual
Current assignee: Individual
Priority date: 2016-04-01
Filing date: 2017-04-03
Publication date: 2017-12-06
Also published as: GB201705307D0

Abstract

A remotely powered propulsion system 1 comprises electromagnetic (EM) energy (2q1, 2q2, 2q3, 2q4, figs 9 and 10) transmitting means 2, 2s1, 2s2, for emitting EM energy and a receiver 3 which can receive said energy for propulsion means 3s1, 3s2. The propulsion means is able to convert the transmitted EM energy into kinetic energy for the purpose of materially propellant-free aerospace propulsion applications. The movement of matter external to the craft is not required for said crafts propulsion. The propulsion means has a number of applications such as thermonuclear impact fusion for electrical power, heat generation, radioisotope production, nuclear reaction product generation, astronomical body impacting for earth-crossing threats, rapid missile interception for national defence, terminal ballistic testing of spacecraft fuselage and unprecedented impact physics research.

Description

Remotely Powered Propulsion System

Background

This invention relates to a propulsion system for accelerating macroscopic objects to unprecedented velocities in the laboratory reference frame.

When macroscopic objects are to be accelerated to hypervelocity, substantial quantities of energy must be efficiently transformed into kinetic energy for an appreciable increase in projectile velocity to be observed.

However, the use of conventional macroscopic hypervelocity accelerators can lead to a number of difficulties. If light-gas guns are used, the projectile rapidly accelerates to a velocity in which it "outruns" the expanding gas that had previously driven it; as such, efficiency is limited as the projectile velocity becomes comparable to the gas atoms' thermal speed. Furthermore, electromagnetic rail guns produce friction between the sabot or projectile and the rails, thermal plasma forms the electrically-conductive contact between sabot or projectile and rails and energy is lost as heat, the gun's rails are eroded and the repulsive forces are not localised to the projectile and threaten to damage the rails.

Statement of Invention

To overcome the previously outlined problems, the present invention proposes a remotely powered propulsion system, with energy transmitting means, for emitting electromagnetic or EM energy, a receiver which can receive said energy for the propulsive means, and said propulsive means for converting the transmitting means' received EM energy into the propulsive means' kinetic energy .

Advantages

The transmitting means is preferably provided by an electric dipole antenna, although the transmitting means may also be provided by other means , such as a magnetic dipole antenna, or a length of electrical current-carrying cable.

The transmitting means may be adjustable so that the driving frequency of the electromagnetic or EM radiation can be modified to suit the receiving antenna's relative position, ferromagnetic material selection, EM cavity dimension and the user's requirement.

The Remotely Powered Propulsion System or RPPS may be a self-contained unit with all required components enclosed within the chassis or fuselage of a vehicle or craft; alternatively, the RPPS may have its transmitting means mechanically decoupled from its receiver's propulsive means.

The propulsive means is preferably provided by the Lorentz force, although the propulsive means may also be provided by other means, such as a magnetoplasmadynamic-type thruster, or arc-jet type of thermal plasma electrical propulsion or any electrically-powered rocket motor-type that exploits electromagnetic forces and thermal pressure to derive thrust .

The propulsive means may be adjustable so that the phase relation between magnetic flux and current density vectors can be modified to subsequently enhance the Lorentz force acting on the ferromagnetic current-carrying components to suit the user's requirement.

The RPPS may be comprised of elliptical reflectors so as to efficiently focus most of the transmitting means' EM energy onto the receiver for conversion to kinetic energy by the propulsive means.

Introduction to Drawings

The invention will now be described solely by way of example and with reference to the accompanying drawings in which:

Figure 1 introduces the first embodiment of the labelled Remotely Powered Propulsion System or RPPS mark I, with features such as the transmitting antenna, said antenna's electrical power supply leads, the antenna terminals, thruster assembly, ferromagnetic receiving antenna terminals, tuning capacitor, tuning inductor, antenna support structure or vehicular fuselage or chassis,

Figure 2 follows on from figure 1 and shows the transmitting antenna energised with the two upwardly-pointing arrow headed lines denoting electrical dipole moment, the two diverging upwardly-pointing diagonal undulating arrow-headed curves denote electromagnetic or EM radiation of wavelengths shorter than the separation between the aforementioned transmitting antenna and ferromagnetic receiving antennae terminals, said radiation propagates from the transmitting antenna to at least one receiving antenna assembly,

Figure 3 follows on from figure 2 and shows the Lorentz forces acting parallel to the direction of propagation of the aforementioned EM radiation, the Lorentz force acts upon the ferromagnetic receiver terminals which are mechanically supported by the fuselage or the frame, the Lorentz force is depicted by the two pairs of diagonal arrow-headed lines, the lift vector is depicted by the vertical arrow-headed line,

Figure 4 introduces the labelled second embodiment of the Remotely Powered Propulsion System or RPPS mark II, the RPPS mark II as depicted in figure 4 is in perspective view, the largest cylinder is the transmitting antenna, labelled features are said transmitting antenna and its electrical power input and output terminals, receiving antenna's tube and its receiving antenna inlet or breech as well as its receiving antenna outlet or muzzle, upper or dorsal elliptical reflector, lower or ventral elliptical reflector,

Figure 5 follows on from figure 4 and shows the RPPS mark II from its distal or proximal end, to reiterate, the viewing axis is parallel to the aforementioned cylindrical transmitting antenna's axis of rotational symmetry,

Figure 6 follows on from figure 5 and shows the transmitting antenna energised and emitting EM radiation radially in cylindrical geometry, the electric field wave-vector component, the magnetic field wave-vector component and Poynting vector of said radiation is depicted by the arrow-headed lines,

Figure 7 follows on from figure 6 and shows the EM radiation propagate radially outwards towards the vacant receiving antenna's tube,

Figure 8 introduces the labelled RPPS mark II's moving part - the receiving-antenna rocket or RAR,

Figure 9 is a perspective view of the RPPS mark II and follows on from figure 4, figure 7 and figure 8 by showing two of the aforementioned and labelled RARs entering into the receiving antenna's tubes, figure 9 also depicts the labelled electrical current or electrical field responsible for the EM radiation generation and propagation in figure 6 and in figure 7,

Figure 10 is a proximal or breech end view of the RPPS mark II and follows on from figure 7 and from figure 9 and shows the EM radiation approach the presently RAR-filled receiving antennas' tubes - either directly or by reflection from the elliptical reflector, RAR is labelled, as is the EM radiation Poynting vector, dorsal and ventral elliptical reflectors ,

Figure 11 follows on from figure 8 to figure 10 inclusive and depicts the labelled electrical field component of the EM wave energising the propellant energising assembly or PEA upon the RAR's entry into the receiving antenna's tube, the RAR's labelled features are payload, nozzle media, energising assembly, rifled jacket, complementarily said tube is a barrel which is also labelled and rifled,

Figure 12 follows on from figure 11 and shows the transformation of the PEA into plasma which is expelled from the RAR with a force which is labelled, the aforementioned RAR rifled jacket and receiver antenna's rifled barrel induce RAR rotation - which is also labelled, the barrel is labelled, vaporised nozzle media and the aforesaid plasma - are also labelled, figure 12 shows the RAR begin to emerge out from the barrel's muzzle,

Figure 13 follows on from figure 9 to figure 12 inclusive; figure 13 shows two RAR remnants emerge from the aforementioned barrel's muzzle, the RAR interior continues to eject matter from its nozzle - even though it is clear of the muzzle, and no longer energised by the EM wave, the rotating or spinning RAR is subsequently stabilised in flight and travels by rocket propulsion; accelerating until it has ejected all of its vaporised material or acquires terminal velocity through the medium in which it travels.

Detailed Description

The following section will describe the present invention in detail and elaborate on the accompanying drawings. In general, the drawings are arranged in a sequence so as to illustrate the present invention's operation in chronological order. Prior to describing the invention, the following paragraphs aim to introduce the reader to the formalities and conventions used in the drawings.

Components are geometrically described by lines without arrows. Components may not necessarily be drawn to scale. A blue-print or engineering-type drawing would not concisely illustrate the operating principle of the invention. Rather, a symbolic representation is provided in the drawings in order to readily confer understanding of the invention to the reader. To indicate the required sizes, dimensions, angles, mechanical and thermal radiation quantities and properties, arrows are used in the drawings to highlight important features and requirements to carry out the invention which are to be explained in the present detailed description section. The reader is not to confuse arrows with those seen in engineering-type drawings explicitly showing construction lines, materials or dimensions.

The requirements the components must fulfil to carry out the invention are explained. The arrows in the accompanying drawings of the present invention are to indicate quantities of mechanical and electromagnetic or EM relevance which may be calculated by following the guiding physical principles alluded to in the present document.

Components and quantities are labelled and described qualitatively and are essential or at least preferable to carry out the present invention. For example, component 2 is the transmitting antenna and its subcomponent 1 is one of its electrical power lead. Therefore 2sl is a transmitting antenna's electrical power lead. By extension, the primary function of the transmitting antenna is to radiate EM waves. Therefore, 2q4 is the Poynting vector or EM wave energy-flux emitted from the antenna. Closed arrow-headed lines or closed arrows indicate the movement of matter. Open arrow-headed lines or open arrows indicate EM field or EM flux direction.

As a point of note, the present patent application is to claim priority from patent application numbered GB1605639.2 filed on April 1st 2016. The present application is a development of all priority date linked-patent applications numbered GB1116356.5 filed on 20th September 2011 through to the granted patent application granted on 15th June 2016 with patent serial number GB2494941. As such, the present patent application makes reference to all the above published patent applications as background and supporting prior-art knowledge [1]. Additionally, the Massachusetts Institute of Technology "Climate CoLab" "Energy Supply" contest held in the year 2016 was entered by the author and the financier of the aforesaid granted patent; links to their contest entry will be made available on the author's website [2]. The contest entry was titled; "Small Guided Missile or Dart for Thermonuclear Impact Fusion: Vast, Clean Energy". Said contest entry proposal has a number of references which may be referred to with the present patent application; in particular the "Proceedings of the Impact Fusion Workshop: National Security & Resources Study Centre, Los Alamos Scientific Laboratory, Los Alamos, New Mexico, July 10-12, 1979". The commonality of the technologies outlined in the present application allow for the acceleration of macroscopic objects to extreme hypervelocities by the use of electromagnetic radiation.

The remainder of the detailed description section is dedicated to the chronologically-sequenced descriptive narrative explaining the operation of the invention.

Figure 1 depicts a simplified ellipsoidal craft 1 whose fuselage cross-section is described as an ellipse and rotationally-symmetric about the minor axis. Figure 1 depicts the craft with a flat base for stable landing. At the horizontal centre of the circular cross-section and near the base or vertical centre there is a transmitting antenna 2 with electrical power leads 2sl, 2s2, leading to a source of time-varying electrical field 2ql such as an alternating current generator; although a pulsed EM power source instead of, or as well as, the implied sinusoidal EM power source can be envisaged. The transmitting antenna is coupled to said power source.

The frequency of the EM power source is selected such that the spatial separation between the transmitting antenna terminals 2s3, 2s4 and the radiated EM wave is multiple wavelengths away from the receiving antenna 3 or thruster assemblies' ferromagnetic 3sl,3s2 terminals. This is required so that the EM coupling between the transmitting antenna 2 and thruster assembly is minimised such that momentum transfer is desirable. Said requirement allows the craft to be propelled by an on-board, mechanically coupled and self-contained transmitting antenna 2 in a similar fashion to being propelled by a remote, mechanically isolated antenna. An on-board, mechanically coupled and self-contained transmitting antenna can propel the craft that it is mounted within, on or to, without violating the conservation of momentum.

Said condition holds for as long as the EM radiation is not considered in the "near field" or "near zone" as it impinges upon the thruster assemblies 3 within, on or mechanically attached to the same craft. This nuance will be explained further in the present document.

In the case of the near-zone EM fields from a transmitting antenna coupling with surrounding matter, and moving said matter so as to propel said antenna in an opposite direction; the reader is directed to patent application numbered GB1610677.5 and the relevant priority documents. Said application describes a spacecraft, configured to be, or connected to, a rectifying antenna. Said antenna derives its propulsion by inducing electrical currents in, and exerting pressure on, surrounding astrophysical plasmas. Said plasma includes the solar wind and planetary ionospheres. In this case, the plasma and spacecraft move in opposite directions. As such, the antenna and surrounding plasma must be directly momentum-coupled, which implies "nearzone" operation of the antenna at wavelengths exceeding the dimensions of the craft.

In the present invention, a self-enclosed craft, contains an antenna, whose EM radiation is preferentially absorbed, reflected or transmitted and intensified on different sections of its surfaces so as to derive thrust on the proviso that the radiation transmitted from antenna to said surfaces has wavelengths smaller than the craft's dimensions. To reiterate, exemplify and simplify, an antenna contained within a spacecraft can propel said spacecraft without the requirement of matter flowing out from, or surrounding, the spacecraft.

On the other hand, a remote, mechanically isolated antenna is discussed in the priority patent application numbered GB1605639.2; from which the present patent application claims priority; an alternative embodiment of the present invention. To reiterate, exemplify and simplify, a ground-based antenna can propel a space-borne spacecraft.

Additionally, the applied frequency of the EM power source must be low enough, or the EM wavelength must be long enough so that the magnetic permeability of the ferromagnetic 3sl, 3s2 terminals exceeds that of paramagnetic material.

To reiterate, at the EM power source frequency, the ferromagnetic or non-paramagnetic 3sl, 3s2 terminals must be able to substantially enhance the magnitude of the magnetic flux density 2q3 of the EM wave that impinges upon it. These design specifications are a clarification and inventive-step above and beyond the author's previously-published patent applications.

Furthermore, the tuning capacitor 3s3 and inductor 3s4 are a clarification and inventive-step further than the author's previously-published patent applications. Specifically, these act to keep any induced electrical current density vector in the non-paramagnetic terminal material 3sl, 3s2, in a good phase relationship with the magnetic flux density vector permeating said material, such that the resulting Lorentz force vector - generated orthogonally to both of the aforementioned vectors - has the optimum magnitude. Although only two thruster assemblies 3 are depicted in the accompanying drawings, the number of thruster assemblies may differ in practice .

Figure 2 shows two rays of EM radiation emitted from the transmitting antenna 2. In practice, an antenna as depicted in figure 1 will radiate in all directions. In the "far zone", the radiation flux will tend to zero parallel and anti-parallel to the electric dipole moment as depicted in figure 2. In an alternative invention embodiment, a magnetic dipole antenna, or other EM antenna configurations are envisaged.

For an ideal antenna, the radiation in the far zone case; is emitted radially outwards such that any EM momentum flux in a particular direction is balanced and nullified by EM momentum flux in the opposing direction. This symmetry prevents the antenna from thrusting itself by virtue of its radiation.

For an non-centred antenna enclosed within a cavity 4 such as the fuselage of a spacecraft, the reflected momentum vectors may allow for minutely oscillating thrust vectors on the craft, but not for propulsion over large spatial displacements .

For partially-enclosed cavities or cavities composed of material with different properties, such as an antenna at the focus of a parabolic reflector, or an antenna enclosed within a surface composed of reflecting and transmitting materials; definite, thrust can be observed to move the craft over large distances: This is photonic propulsion and the thrust-levels are low.

For the present invention, photonic propulsion thrust to power ratio is amplified by absorbing and concentrating the magnetic field energy of an EM field incident on a non-paramagnetic material. The same material carries an electrical current induced by said EM wave and the resulting Lorentz force, although parallel in direction to the aforementioned photonic thrust, is actually much greater in magnitude. Figure 3 shows this. To maximise thrust, ferromagnetic terminals 3sl,3s2 are used to increase the magnetic flux density vector's magnitude of the impinging EM wave 2q4. Said terminals are separated such that the electric field component of the EM wave will induce a current density within the magnetised ferromagnetic material.

Said current can be phase-shifted with respect to the magnetic flux density's phase by the control capacitor 3s3 and control inductor 3s4 in the thruster assembly 3. As such, both the orthogonal current and magnetic flux density vectors maxima can be synchronised so as to yield the maximum, mutually orthogonal Lorentz force vector. The net Lorentz force vector sum can be used to thrust the craft. Alternatively, phase misalignment of at least one thruster assembly 3 can re-orient or "steer" the craft 1.

If the thruster assemblies 3 are positioned only on part of the surfaces 4 enclosing the antenna 2, then Lorentz forces directing radially away from the antenna 2 and generated at the thrust plates 3 will not necessarily be balanced and nullify. See figure 3. Said unbalanced forces will permit the spacecraft 1 to thrust and travel through space without the counter-movement of matter external to said craft.

The spacecraft could be spherical 4, or of any other geometry, with the transmitting antenna 2 dipole-moment at the spacecraft's geometric or mass centre. Additionally, the thruster assemblies 3 could be distributed as evenly as possible around the entire surface 4 enclosing said antenna 2. In this scenario, the spacecraft 1 would still be able to thrust as a whole in a preferred direction if at least one of the thruster assembly's magnetic flux and current density vectors were in a different phase relationship to the remainder of the thrust assemblies 3. This scenario was not explicitly depicted in the accompanying drawings.

The accompanying drawings of figures 1 to figure 3 inclusive are meant to convey the use of capacitive and inductive structures capable of being energised by electromagnetic radiation. Thrust assembly 3 construction could be carried out as depicted, or by the use of standard electrical components whose paramagnetic electrically conductive materials are replaced by ferromagnetic electrically conductive materials. The depictions are a non-exhaustive example.

From figure 4 onwards, the second embodiment of the present invention is depicted in the accompanying drawings. Here, an earlier invention of the author's is adapted from its thermal radiation focusing applications to EM radiation focusing at typically longer wavelengths. For the present invention, two elliptical reflectors are arranged such that one focal axis of each of the elliptical trough reflectors 4sl,4s2 coincides 2 and runs parallel to one another. The other focal axis of each cylindrical trough 3s4,3s4 lies in the same plane as, and parallel to the aforementioned coinciding elliptical reflector axes 2. Hence, when either high-current or rapidly varying electrical current 2ql or both high-current or rapidly varying electrical current 2ql passes through the transmitting antenna 2 - which corresponds to the aforementioned coinciding focal axes; EM radiation is emitted radially as depicted in figure 6, reflected as depicted in figure 7 and figure 10 and focused onto the barrels 3s4 which both correspond to the non-coinciding elliptical reflector axes.

The rocket-antenna receiver or RAR 1 as depicted in figure 8 may either be at rest in the breech of the barrel 3s4 or travelling through the breech of the barrel as the EM radiation penetrates said barrel and impinges upon the RAR. In the literature, it is understood that rocket engines are relatively inefficient at low velocities and relatively efficient at high velocities. As such, the RAR may be "fired" into the breech at velocities technologically available at present by devices such as light gas guns, EM rail guns, magnetic launchers or conventional firearm propellant as a few non-exhaustive examples. What is essential is for the transmitting antenna 2 to emit EM radiation which will impinge upon the RAR 1 at or near the breech 3s4. The barrel 3s4 must have high transparency of the EM radiation at the applied frequency. This maximises the energy available for absorption within the RAR as the EM wave impinges said RAR 1.

What is essential is for the EM wave energy transmitted from the transmitting antenna 2 to be converted into the kinetic energy lql of the rocket receiving antenna 1 as depicted in figure 12. The mechanisms by which this can be effected are outlined in a non-exhaustive list in the priority application to the present patent application. To briefly re-visit some of the methods:

An electric arc across two electrodes within the RAR 1 will produce thermal plasma which vaporise propellant 3s2,3sl for expansion through a nozzle 3sl into the barrel and surrounding medium. The thrust-type derived is similar to arc-jet propulsion as is described in the art.

An electric arc with current density within the RAR has orthogonal components to a magnetic field such that the Lorentz force acts to flush out the propellant at high velocity anti-parallel to the desired flight direction of the RAR. Alternatively, the Lorentz force can accelerate and heat plasma 3sl, 3s2, for thermal expansion through a nozzle as described in the previous paragraph .

The nozzle medium 3sl may be a plug that acts as a sabot by which expanding gases are used to propel the SAR into the breech 3s4 of the RPPS by light-gas gun or conventional firearm technology. It may also act as a tamper to allow the pressure and temperature of the propellant 3s2 to rise whilst it is energised by the EM wave electric field 2q2. The tamper may act as an electrode or receiver antenna terminal. The tamper disintegrates or is expelled when the propellant 3s2 temperature and pressure are sufficiently high .

Inductively coupled plasmas are also an option provided that the plasma is thermal in its nature. In general, existing technologies typically used for spacecraft propulsion with dense, thermal plasma as the working fluid are considered with the present invention. Different transmitting and receiving electrode configurations may be deployed to exploit different electromagnetic heating and thrusting mechanisms .

At this point, the author would like to point out that the Remotely Powered Propulsion System first and second embodiments or RPPS mark I and RPPS mark II are similar in their use of EM radiation to induce electrical current and generate thrust. However, whereas the RPPS mark I generate thrust via the Lorentz force with heat emitted as an un-useful by-product, RPPS mark II uses heat to generate thrust and can augment said thrust with the Lorentz force.

In short RPPS mark II is a higher-electrical current application of RPPS mark I .

Returning to RPPS mark II figure 12, the rifled jacket 3s2 of the RAR 1, the rifled barrel 3s4 and the thrust lql of the RAR causes the RAR 1 rotation lq2 about the thrust axis lql, conferring upon the RAR angular momentum lq2.

As an aside, the rifled surfaces 3s3, 3s4 indirectly optimise EM energy to thermal energy to RAR kinetic energy transfer for RPPS mark II; as the tuning capacitor 3s3 and inductor 3s4 more directly optimise EM energy to RAR kinetic energy transfer for RPPS mark I.

Ideally; returning to RPPS mark II, the angular acceleration to angular velocities and centripetal forces and pressures just below material stress limits are applied just as the RAR 1 emerges from the muzzle. Like a gyroscope, the RAR orientation is stabilised in flight. Some precession is anticipated, and this "wobbling" effect about the intended flight axis may result in an undulating or wave-like trajectory of the RAR whilst it is still thrusting by evacuating plasma 3s2 from its interior or cavity. The nature of the oscillation is required knowledge so that the target position and size down-range of the RPPS mark II can be ascertained. For maximum hypervelocities, the RAR may have to travel from the RPPS muzzle 3s4 into an evacuated chamber, whose dimensions will be determined by deviations from the intended flight-path.

The elliptical reflectors 4sl,4s2 behave as a waveguide in which unabsorbed EM energy can reverberate by repeatedly converging upon and diverging away from the barrels 3s4. Ideally, the energy is absorbed mostly by the RAR 1 within the barrel 3s4. The high electrical permittivity and high magnetic permeability of the RAR 1 components, within the aforementioned waveguide 4sl,4s2 endows the RAR with high EM energy density. In essence, the RAR 1 sweeps through the barrel 3s4, collecting reverberating EM energy, converting it into thermal energy. Said RAR 1 thermal energy is converted into RAR kinetic energy within the barrel 3s4 and in flight away from the barrel's muzzle. The majority of the kinetic energy is acquired once the RAR leaves the barrel to prevent frictional forces from retarding its acceleration.

The applications for RPPS mark I are for vehicular propulsion; particularly in the aerospace sector. Whereas RPPS mark II can be utilised for terminal ballistic applications ranging from spacecraft fuselage testing to hypervelocity interception of threatening extra-terrestrial objects, long-range missile interception, and finally, thermonuclear impact fusion for electrical power generation.

End of Description Section

References [1] A search with patent application numbers GB1605639.2, GB1116356.5 filed and patent serial number GB2494941 on the United Kingdom of Great Britain's intellectual property office website www.ipo.gov.uk or the European Patent Office website www.epo.org should yield the relevant prior art and supporting documents .

[2] A visit to the author's website www.quaws. name and a link provided on each page directs the browser to the Massachusetts Institute of Technology "Climate CoLab" "Energy Supply" contest and the author's profile on that platform where the relevant proposal "Small Guided Missile or Dart for Thermonuclear Impact Fusion: Vast, Clean Energy" is available.

Claims

1. A remotely powered propulsion system, which comprises energy transmitting means, for emitting electromagnetic or EM energy, a receiver which can receive said energy for the propulsion means, and said propulsion means for converting the transmitting means' EM energy into the propulsion means' kinetic energy.

2. A remotely powered propulsion system according to claim 1 in which the transmitting means is provided by either or both of the following: a high electrical current, a rapidly time-varying electrical current; either or both passing through or moving along or across an electrical conduit.

3. A remotely powered propulsion system according to claim 2 in which the receiving means is provided by the following conditions: At least one ferromagnetic body, suitable electrical capacitive properties, suitable electromagnetic inductive properties; to facilitate the generation of Lorentz forces; the associated Lorentz pressure can be orders of magnitude greater than the photonic pressure incident on surfaces without the aforementioned conditions in the present claim.

4. A remotely powered propulsion system according to claim 3 in which a body such as payload, cargo, a vehicle or craft can experience the noticeable, useful and non-trivial force magnitudes to control or prevent said body's motion by a remotely-located transmitting means and the selection of driving frequencies suitable for the remoteness and its ferromagnetic component's characteristics, which is not adjoined to, or mechanically or materially connected with the body.

5. A remotely powered propulsion system according to claim 4 in which a body such as payload, cargo, a vehicle or craft can experience the noticeable, useful and non-trivial force magnitudes to control or prevent said body's motion by a transmitting means and the selection of driving frequencies suitable for the body's dimensions and its ferromagnetic component's characteristics, which is adjoined to, or mechanically and materially connected with the body.

6. A remotely powered propulsion system according to claim 2 in which the electrical conduit coincides with one focal axis of a reflective elliptical tube and the receiver is able to move or is projected along the other axis of said tube.

7. A remotely powered propulsion system according to claim 2 in which the electrical conduit coincides with the common focal axes of a plurality of reflective elliptical troughs and at least one receiver is able to move or is projected along the other axes of said troughs.

8. A remotely powered propulsion system according to claim 6 or claim 7 in which the receiver is projected along the aforesaid axis by preexisting technologies prior to its illumination by the transmitting means as a rocket-engine efficiency-enhancing measure.

9. A remotely powered propulsion system according to claim 6 or claim 7 or claim 8 in which the receiver converts the absorbed EM energy into thermal internal energy of the propellant into kinetic energy by the expulsion of hot expanding plasma through the propulsion means' nozzle.

10. A remotely powered propulsion system according to claim 9 in which the receiver also converts a portion of the absorbed EM energy directly into plasma particle kinetic energy by the exploitation of EM fields within the propulsion means.

11. A remotely powered propulsion system according to claim 10 in which the aforesaid directly converted particle kinetic energy is converted into thermal internal energy, augmenting the thermal energy contribution to thrust for rocket propulsion as was described for claim 9 .

12. A remotely powered propulsion system according to claim 9 or any of the subsequent claims preceding the present claim in which the receiver travels through a tube or barrel due to the propulsion means so as to guide the receiver until it exits said barrel's muzzle.

13. A remotely powered propulsion system according to claim 12 in which either or both of the following are rifled: The aforementioned barrel, the projectile; where said projectile comprises an EM radiation receiver and propulsion means; such that projectile rotation is induced to guide the projectile in flight subsequent to its departure from the barrel muzzle.

14. A remotely powered propulsion system according to claim 9 in which the resulting hypervelocity projectile can be utilised for the following non-exhaustive applications; thermonuclear impact fusion for electrical power, heat generation, radioisotope production, nuclear reaction product generation, astronomical body impacting for earthcrossing threats, rapid missile interception for national defence, terminal ballistic testing of spacecraft fuselage and impact physics research .