rexresearch.com


Roger J. SHAWYER

Electromagnetic Space Drive







SPR Ltd

Unit 40, Broadmarsh Centre, Havant, Hampshire, United Kingdom, PO9 1HS
Tel : 01243 377783
http://www.emdrive.com
sprltd@emdrive.com


SPR Ltd.

Theory

Development

Benefits

"EM Thrust Drive Technology will Dominate Space"

Tom Sheelley : "No-Propellant Drive Prepares for Space and Beyond"

Criticism

GB # 2399601 : Thrust Producing Device using Microwaves

GB # 2334761 : Microwave Thruster for Spacecraft

GB # 2229865 : Electrical Propulsion Unit for Spacecraft

USP # 5543801 :  Digitally Controlled Beam Former for a Spacecraft

SPR Ltd : Theory Paper [ PDF ]

Roger Shawyer : The Dynamic Operation of a High Q EMDrive Microwave Thruster [ PDF ]

Yang Yuan, et al. : Net Thrust Measurement of Propellantless Microwave Thrusters [ PDF ]

Yang Juan, et al. : Effectively Calculating Performance of Microwave Radiation Thruster [ PDF ]


http://emdrive.com

SPR Ltd.

Satellite Propulsion Research Ltd (SPR Ltd) was formed in October 2000 as the corporate vehicle for progressing the development of a new form of electric propulsion. This electromagnetic “EmDrive” technology provides direct conversion of electrical energy to thrust, using radiation pressure at microwave frequencies in a tapered, high Q, resonant cavity.

The first UK government funded programme was a feasibility study completed in 2002. This work confirmed the theoretical predictions in a large series of tests using an experimental thruster. In addition, the huge potential savings for the space industry were identified during the preparation of a business model.

Following an independent review of the feasibility study report, a Demonstrator programme was authorised. This covered the design, manufacture and test of an S Band Demonstrator Engine. The Engine successfully demonstrated viable performance in both Static and Dynamic test programmes, and provided the basic knowledge to design the Flight Model Thruster. Once again a full technical report was prepared and reviewed before acceptance by the UK government.

A feasibility study is currently underway to investigate second generation superconducting technology. This includes the design, build and test of an experimental thruster operating at liquid nitrogen temperature.

A flight model development programme has started on a 300W, C band thruster, specified to produce 85mN thrust.
 
 


http://emdrive.com/principle.html

Theory

Principle of Operation

At first sight the idea of propulsion without propellant seems impossible. However the technology is firmly anchored in the basic laws of physics and following an extensive review process, no transgressions of these laws have been identified.

The principle of operation is based on the well-known phenomenon of radiation pressure. This relies on Newton’s Second Law where force is defined as the rate of change of momentum. Thus an electromagnetic (EM) wave, travelling at the speed of light has a certain momentum which it will transfer to a reflector, resulting in a tiny force.

If the same EM wave is travelling at a fraction of the speed of light, the rate of change of momentum, and hence force, is reduced by that fraction. The propagation velocity of an EM wave, and the resulting force it exerts, can be varied depending on the geometry of a waveguide within which it travels. This was demonstrated by work carried out in the 1950’s. (CULLEN, A.L. ‘Absolute Power Measurements at Microwave Frequencies’ IEE Proceedings Vol 99 Part 1V 1952 P.100)

Thus if the EM wave travelling in a tapered waveguide is bounced between two reflectors, with a large velocity difference at the reflector surfaces, the force difference will give a resultant thrust to the waveguide linking the two reflectors. If the reflectors are separated by a multiple of half the effective wavelength of the EM wave, this thrust will be multiplied by the Q of the resulting resonant cavity, as illustrated in fig 1.

Fig 1. Diagram of an engine concept.

The inevitable objection raised, is that the apparently closed system produced by this arrangement cannot result in an output force, but will merely produce strain within the waveguide walls. However, this ignores Einstein’s Special Law of Relativity in which separate frames of reference have to be applied at velocities approaching the speed of light. Thus the system of EM wave and waveguide can be regarded as an open system, with the EM wave and the waveguide having separate frames of reference.

A similar approach is necessary to explain the principle of the laser gyroscope, where open system attitude information is obtained from an apparently closed system device.

Video : http://www.youtube.com/watch?v=57q3_aRiUXs


http://emdrive.com/demonstratorengine.html

The Development of a Demonstrator Engine

Although the experimental thruster had verified the static thrust equation, it became apparent that the concept would not become generally accepted until a viable engine could be demonstrated. Accordingly, a proposal for the design, manufacture and test of a complete demonstrator engine was submitted to DTI. A Research and Development grant was awarded in September 2003 and the work started with a mission analysis phase.

This work enabled the specification of the demonstrator engine to be optimised against the requirements of a typical commsat mission. Unlike the experimental thruster, the engine would be rated for continuous operation and extensive design work was required to increase the specific thrust by raising the design factor and unloaded Q.

The engine was built with a design factor of 0.844 and has a measured Q of 45,000 for an overall diameter of 280 mm. The microwave source is a water cooled magnetron with a variable output power up to a maximum of 1.2 kW.

To obtain the predicted thrust the engine must maintain stable resonance at this high Q value. Major design challenges have included thermal compensation, tuning control and source matching.

The engine was tested in a large static test rig employing a calibrated composite balance to measure thrust in 3 directions, up, down and horizontal. A total of 134 test runs were carried out over the full performance envelope, with a maximum specific thrust of 214mN/kW being measured.


http://emdrive.com/benefits.html

Benefits

Cost Performance Improvement

The EmDrive makes possible a big improvement in the cost performance of the Next Generation of Satellites.

The EmDrive offers a more elegant solution to satellite propulsion than any other form that exists today.

Satellites are not burdened with heavy propellant subsystems. The satellite platform configuration can dispense with the tanks, pipes and valves and the propellant itself.

Launch site and programme costs are reduced as propellant procurement and handling costs are eliminated.

LEO to GEO Security and more effective Mission Control

Satellites launched into LEO can then be positioned into their allocated orbit using the existing solar array generated electrical power in 30 days.

The satellite is more maneuverable and errors in positioning can be corrected from Mission Control without damage or loss to the satellite.

Launch mass savings of 60% plus can be made per satellite launch.

Overall cost savings in fuel, orbit management and satellite design can save up to 70% of the total mission cost.

At least two satellites can be launched using the same launch vehicle.

Further satellite redesign can take advantage of the reduction in hardware required, which could enable three satellites of similar mission to replace the one satellite using conventional thruster combinations.

Longer Satellite Lifetime

The specified operating life of the most critical component, the microwave source cathode is 15 years.

Standard space industry cathode technology can be employed.

Test cathodes have given an accelerated lifetime of up to 30 years.

With the increased deployment of space stations, a satellite can be moved from its orbit to a space station for both scheduled and unscheduled maintenance.

The simpler satellite platform layout will enable maintenance as well as upgrades to the payload to be made more quickly and efficiently.

A typical satellite using such maintenance procedures could have an effective lifetime of as much as 30 to 45 years.

Flexible repositioning of Orbits

An almost unlimited energy source from the solar panels, via the onboard batteries, offers unlimited orbit adjustment.

Orbit adjustment can be made at any time and can be made on a continuous basis as required.

Payloads and Missions can be Enhanced

Existing launchers can be used to launch larger satellites into Geostationary orbit.

Satellites of 20 tonnes per launch will not be uncommon.

Deep space probes can be made to go deeper into the outer reaches of space.

Scientific Missions can stay in operation longer and can be manipulated to view ad hoc situations without fear of loss of fuel for the thruster.

The Commercial and Social Benefits of Lower Satellite Costs

Lower satellite launch and operational costs can be used to reduce transponder prices to satellite service providers typically by 50-70%.

1. Increased use of satellites for two way broadband communications for data, image and phone will open up new markets.

2. New markets and lower costs will provide more competition amongst SSPs and increase technology investment in new applications.

3. Technological innovation, which to date has been slower in space related endeavours will now be accelerated.

4. This will lead to new markets for satellite data transmission.

5. It will provide a more universal coverage of information for all countries and regions on this planet.

Lower all-round costs and longer effective satellite lifetimes will ease budgetary pressures on governments in the race in space.

In the emerging countries new social services can be opened up on a wider scale than are currently available.

1. Broadband multimedia can be used for education, health awareness, field hospitals and instruction on the maintenance of consumer goods.

2. It will enable schools and communities in different countries to talk to each other and share their cultures and ideas.

3. Wildlife can be monitored more cost effectively on a wider scale.

4. Weather monitoring and mineral prospecting, to name two examples can be employed more cost effectively and on a more timely basis.

5. Global emergency planning in the event of natural disasters can have a wider more cost effective coverage.



http://news.softpedia.com/news/Emdrive-Thrust-Technology-Will-Dominate-Space-94224.shtml

EM Drive Thrust Technology Will Dominate Space

 China's space attempt may be fueled by the newly-emerging technology that the majority of scientists are now contesting. Chinese researchers are currently trying to build the "Emdrive" (electromagnetic drive), which would change all space propulsion systems in case it becomes functional.

Basically, the technology relies on converting electricity into thrust by means of microwaves, without contradicting the laws of physics. The original idea belongs to British scientist Roger Shawyer and, after the British government seemed to have lost interest in it, it has been bought (read financed) by a Chinese company. Currently, the project is undergoing at the Northwestern Polytechnical University (NPU) in Xi'an, China, under the command of Professor Yang Juan who has gathered a lot of experience in the field of microwave plasma thrusters, a similar hardware technology based on a different theory.

The engines built by Shawyer's company (Satellite Propulsion Research – SPR) for demonstrative purposes produce thrust by means of a microwave-filled tapering resonant cavity. Australian physicist John Costella says: "It is well known that Roger Shawyer's 'electromagnetic relativity drive' violates the law of conservation of momentum, making it simply the latest in a long line of 'perpetuum mobiles' that have been proposed and disproved for centuries. His analysis is rubbish and his 'drive' impossible." That's an example of how Shawyer's theory is perceived by other experts. But no great idea came to happening without a strong public disbelief and opposition. Some even claim that, although SPR's work is based entirely on Einstein's principles, this particular theory and device type violates Einsteinian physics laws and, as such, it must be false or wrong at some point.

In reply, Shawyer told Danger Room that "NPU started their research program in June 2007, under the supervision of Professor Yang Juan. They have independently developed a mathematical simulation which shows unequivocally that a net force can be produced from a simple resonant tapered cavity. The thrust levels predicted by this simulation are similar to those resulting from the SPR design software, and the SPR test results."

This kind of thrust is sought to be replaced.

Comparisons between the C-Band Emdrive and NASA's NSTAR ion thruster clearly demonstrate the superiority of the former: Emdrive obtains 85 mN of thrust using about 25% of the power that NSTAR uses to produce 92 mN (about a third of an ounce or 9 grams). It also weighs only 7 kg, compared to NSTAR's 30, not to mention the propellant: while NSTAR uses only 10 grams an hour, Emdrive uses... well, none whatsoever, as it relies on energy.

 If the technology is proved to work, its applications are fantastic. The endurance of satellites would be enhanced, as well as their maneuverability, eliminating the toxic risk factor along the way. The probes sent into the deep space would exhibit faster speed, longer trek abilities, and they would be able to stop at any time, like when they reach their target or meet something interesting along their journey. Based on Shawyer's calculations, a solar-powered Emdrive thrust would be capable, in theory, to carry a manned mission to Mars in 41 days' time.


http://www.eurekamagazine.co.uk/article/9657/No-propellant-drive-prepares-for-space-and-beyond.aspx

No-Propellant Drive Prepares for Space and Beyond

by

Tom Shelley

Tom Shelley reports on progress with the controversial Emdrive and its potential applications in space and on the ground

The Emdrive – originally revealed in Eureka’s December 2002 edition as a way of driving satellites and spacecraft using microwaves – is now demonstrating its ability to produce thrust on a consistent basis and is scheduled to be ready for space use by May 2009.

Meanwhile studies are underway into the design of a superconducting version, with a possible thrust of more than 30kN/kW. While it couldn’t be used to accelerate a rocket, it might well be able to provide enough static lift for a flying vehicle propelled forward by other means.
Despite massive controversy, the project continues to be backed by DTI and private investors, and has now been shown to work. Roger Shawyer, who spent 20 years at Marconi Space Systems (now EADS Astrium), revealed details of the prototype engine and its development plans to an IEE chapter meeting in Portsmouth, where the audience included many ex-colleagues from his former company.

Shawyer stated that many of the claims he is alleged to have made about the Emdrive are untrue, particularly suggestions that it defies the principles of conservation of momentum, or Newton’s Laws of Motion.

The crucial part, as he explained, is that it is a relativistic effect that arises because the waves being reflected at the two ends of the conical cavity into which the microwaves are injected have different effective velocities, and thus different frames of reference, and that a closed microwave wave guide is an ‘open system’ in terms of relativity. According to Einstein, all moving frames of reference are equivalent. Why this should be so, whether one is standing still or going at half the speed of light, nobody knows, and in effect Shawyer’s engine could be chucking Dark Energy out of the back of it and functioning as a conventional rocket. On the other hand, there may be no such thing as Dark Energy, and Shawyer may have stumbled on what is really driving the galaxies apart.

But as he pointed out: “I am just a microwave engineer and all that matters is that it works.”
In the present experimental engine and its immediate predecessor, the cavity is made in the form of a copper cone closed off by flat plates at the wide and narrow ends. The net thrust is proportional to the Q value of the cavity, where Q is the ratio of the amount of stored energy to the amount of energy lost per cycle. Acceleration extracts energy from the system and Q decreases. The development engine has a Q value of 50,000 and produces a specific thrust of 0.315N/kW. The original engine produced a thrust of 1.6 grammes, but could only run for tens of seconds at a time before the magnetron overheated and burned out. The present engine produces 9 grammes of thrust from 300W of microwave power and is continually water-cooled. The internal power density is about 17MW.

A video of the demonstration of the engine in its test cell involved mounting the engine and its cooling system on a beam, and supporting it on an air bearing. The test was undertaken in October last year, producing a thrust of 9.8 grammes, maximum speed of 2 cm/s and a movement distance of 185cm. According to Shawyer, the tests had involved accelerating from rest, deceleration to rest, forward and reverse engine mounting, energising at different start angles and using different input powers.

While 9.8 grammes of thrust from a 100kg of machinery may not sound very much, it is a much better power-to-weight ratio than the best competitive satellite and spacecraft propulsion technology, which involves using an ion engine.

For a 1500W DC input power, an ion drive produces 92mN thrust, whereas an Emdrive, based on current technology, should produce 330mN thrust. Furthermore, an ion engine of this size would weigh 112.5kg plus propellant, whereas an Emdrive would weigh 9kg. And while life under power for an ion engine is about six months, an Emdrive should run for 15 years – or virtually forever, if the microwaves were generated by some solid state device.

The big application is for commercial communication satellites – where ‘Hotbirds’ have a take-off weight of 3 tonnes, of which 1.7 tonnes is propellant – both to get them from Low Earth Orbit to Geostationary and to keep them pointing the right way once they get there. Using Emdrives, says Shawyer, should save £15 billion in launch costs over 10 years. While the thrust from the Emdrive would be small, it should get the satellite from Low Earth to Geostationary orbit in 36 days. Power would be from 6kW of solar cells fed to Travelling Wave Tube Amplifiers.

But what really captivated the audience was Shawyer’s proposal for the next stage, which would be to use a superconducting cavity with a Q value of 5 billion and a thrust of 3 tonnes/kW. Unfortunately, one could not use the device accelerate without “causing the Q value to collapse, losing thrust in that vector”, he conceded.

One serious consideration is to develop the technology so that it could be used gently to divert a large asteroid in danger of colliding with the Earth. In fact, prior to Shawyer’s address, David Hall from EADS Astrium had discussed the use of microwave technology to study the internals of Near Earth Object asteroids. He said blowing up such a threat, Hollywood-style, was not really practical – the parts would still be likely to hit the Earth. Current ideas were mostly about finding ways of nudging asteroids into a safe trajectory. As acceleration would be so low, a superconducting Emdrive would be a possible option. A 1kW engine would require 24kW to keep it cool and shifting the asteroid would take somewhere in the order of 10 years, depending on its size.

Shawyer said his team was thinking of using superconducting cavities of a type already being developed and manufactured for a major accelerator project, and cooling them with hydrogen. If applied to lifting a vehicle, the boiling off hydrogen could be used to provide horizontal motion by feeding it to conventional turbofan engines, if in the atmosphere, or to rocket engines for use in space. Whether the technology will ever be used to produce hydrogen-propelled air cars or other wonders, only time will tell.

Pointers

Novel non-propellant microwave drive has reached the point where it can be run continuously

Demonstration engine and cooling system weighs about 100kg and produces just under 10 grammes of thrust

Next engine for satellite propulsion should weigh a little less than 10kg and produce 330mN thrust

A superconducting design able to deliver tonnes of lift thrust (but no acceleration) is being studied



Criticism

"Why Shawyer’s Electromagnetic Relativity Drive is a Fraud" --- John P. Costella: "It is well known that Roger Shawyer’s ‘electromagnetic relativity drive’ violates the law of conservation of momentum, making it simply the latest in a long line of ‘perpetuum mobiles’ that have been proposed and disproved for centuries."

PDF :  http://www.assassinationscience.com/johncostella/shawyerfraud.pdf

Roger Shawyer  replies --

"The momentum exchange is between the electromagnetic wave and the engine, which is attached to the spacecraft. As the engine accelerates, momentum is lost by the electromagnetic wave and gained by the spacecraft, thus satisfying the conservation of momentum. In this process, energy is lost within the resonator, thus satisfying the conservation of energy.

"The emdrive concept is clearly difficult to comprehend without a rigorous study of the theory paper, which is available via emdrive.com or the New Scientist website. This paper, which has been subjected to a long and detailed review process by industry and government experts, derives two equations: the static thrust equation and the dynamic thrust equation.

"The law of the conservation of momentum is the basis of the static thrust equation, the law of the conservation of energy is the basis of the dynamic thrust equation. Provided these two fundamental laws of physics are satisfied, there is no reason why the forces inside the resonator should sum to zero.

"The equations used to calculate the guide wavelengths in the static thrust equation are very non-linear. This is exploited in the design of the resonator to maximise the ratio of end plate forces, while minimising the axial component of the side wall force. This results in a net force that produces motion in accordance with Newton’s laws."

Penny Gruber ( 20:23, 29 September 2008 (PDT) --  AFAIK COM has to apply in any inertial frame of reference. Assuming that the microwave cavity is well sealed as it must be for the high Q's Shawyer's system needs, then no microwaves escape. The magnetron, the waveguide to the cavity, the cavity and all the waves that bounce around inside of it are intrinsically in the same frame of reference, with no ejected mass or energy other than heating from the dielectric and conduction losses of the cavity materials. The thruster ejects nothing and so by COM cannot experience any accelerating force in an external FOR.


Patents & Applications


Thrust Producing Device using Microwaves
GB # 2399601
Roger J. Shawyer

Abstract -- A microwave engine, which produces high thrust, may be used to propel spacecraft where the thrust vector is at ninety degrees to the main velocity vector. It may also be used in an airborne vehicle to counteract gravitational force. The engine comprises a gimbal mounted matrix of a number of superconducting microwave thrusters 11 which are supplied with pulses of microwave energy via an array of switches 15 and enclosed in a Dewar 19 which is maintained at low temperature by liquefied gas. The engine may include an automatic control system to maintain the correct frequency of the microwave generator 7, a means 17 of dissipating the stored microwave energy, and a gyroscopic instrument 21 and motors 22,23 for maintaining the axis of thrust parallel to the direction of gravitational acceleration for an airborne vehicle.

   ... 



Microwave Thruster for Spacecraft
UK Patent Application GB # 2334761
Roger J. Shawyer

Abstract -- The thruster comprises a tapered waveguide comprising a section 1, that is evacuated or filled with air, and a section 6 containing a dielectric resonator or ferrite material whose relative permeability or relative permittivity (or both) have values greater than unity. Microwaves may be introduced into the guide via a slot 2, or a probe. It is stated that the force 9, on the end wall 5, due to reflection of the microwaves, is less than the force 4, exerted on the end wall 3, thereby generating a resultant propulsive thrust. The thruster may be used to enable the orbit of a spacecraft to be maintained or changed over a period of time.



Electrical Propulsion Unit for Spacecraft
UK Patent Application GB # 2229865
Roger J. Shawyer

Abstract -- A unit which will generate thrust when provided with electrical energy at the appropriate frequency. This will enable the orbit of a spacecraft to be maintained or changed when applied over a period of time. The thrust is generated as a result of the difference in the forces obtained when electromagnetic waves are reflected at the end walls 3 and 5 of a resonant waveguide assembly. This assembly comprises an air or vacuum filled end section 1 together with a transition section 6 and an end section 8 containing an electrical material 7.



Digitally Controlled Beam Former for a Spacecraft
USP # 5543801
Roger J. Shawyer

Abstract --  A digitally controlled beam former for a spacecraft which includes means for periodically calibrating the feed paths of the spacecraft's antenna array by measuring the apparent movement of the center of a reference signal and a nominal signal and utilising the measured data to compensate for at least the phase drift in the antenna feed paths. The measured data may also be used to compensate for amplitude and phase drift in the antenna feed paths.

Also published as: EP0642191 (A1) // GB2281660 (A)

Current U.S. Class:  342/354 ; 342/174; 342/372
Current International Class:  H01Q 3/26 (20060101); H04B 007/185 ()

References Cited [Referenced By]
U.S. Patent Documents

3964065 June 1976 Roberts et al.
4280128 July 1981 Masak
4628321 December 1986 Martin
4947176 August 1990 Inatsune
4983981 January 1991 Feldman
5038146 August 1991 Troychak et al.
5093667 March 1992 Andricos
5184137 February 1993 Pozgay
5353031 October 1994 Rathi

Foreign Patent Documents

 0452970A3  Oct., 1991  EP
 4218371A1  Dec., 1992  DE

Description

BACKGROUND OF THE INVENTION

The invention relates to a digitally controlled beam former for a spacecraft.

There is a requirement in spacecraft for active arrays for both beam forming and null operation. The key component of these active array subsystems is a digitally controlled beam former in which variation of amplitude and phase of the individual antenna elements of the spacecraft's antenna array is effected under digital control.

Experience gained from existing spacecraft highlights the difficulties of maintaining phase and amplitude calibration over the life and temperature of x-band digitally controlled beam formers. The requirements of null generation gives rise to a tight specification for these parameters and thereby temperature control within the limits .+-.2.degree. C.

With a relatively large number of antenna array elements and spot beams, thermal control of the beam formers will be difficult to attain and will probably not, therefore, be an acceptable method of controlling phase and amplitude calibration of the beam forming elements.

SUMMARY OF THE INVENTION

It is an object of the present invention to provide a digitally controlled beam former for a spacecraft in which each of the N-paths of the beam former for each element of the spacecraft's receive and transmit antenna arrays is periodically calibrated against a secure tracking telemetry and command (TT&C) uplink. This calibration process not only addresses the major design problem of amplitude and drift in the antenna element feed paths but can also provide the spacecraft with a secure pointing reference which can be utilised to provide back up attitude and orbit control system (AOCS) data in the event that the main optical sensors are disabled for any reason.

The invention provides a digitally controlled beam former for a spacecraft having a multi-element antenna array and a control processor for the antenna array, the beam former including means for periodically calibrating the feed paths of the spacecraft's antenna array by measuring the apparent movement of the centre of a reference signal and a nominal signal and utilising the measured data to compensate for phase drift in the antenna feed paths. The measured data may also be used to compensate for amplitude and phase drift in the antenna feed paths.

According to one embodiment of the invention a digitally controlled beam former is provided wherein the spacecraft has N-paths containing amplitude and phase control elements for each element of the spacecraft's antenna array, wherein the antenna array processor has a number of outputs, each one of which is connected to a separate one of the N-paths for controlling the weighting applied to the amplitude and phase signals of the respective paths; and wherein the beam former includes N-beam former channels, each one of which is connected to a corresponding one of the N-paths of each of the antenna elements; and means for sequentially selecting and calibrating each of the N-channels while the other channels are operational, the weightings of the signals applied to the amplitude and phase elements of the corresponding one of the N-paths of each of the antenna elements being varied in dependence upon the difference between the initial weightings and the weightings required for a reference beam.

According to a further embodiment of the present invention a digitally controlled beam former is provided wherein the antenna array is a receive array and wherein each of the sequentially selected N-channels is calibrated in response to the receipt of a reference uplink signal from a ground transmitter of known location, the reference signal being applied to the corresponding one of the N-paths of each of the antenna elements and causes reference amplitude and phase signals indicative of the location of the source of the reference signal, to be applied thereto, any offset in both the X and Y phases of the reference beam relative to a nominal beam position being detected and applied to the antenna array processor for causing the weightings of the output signals thereof to be varied in dependence upon the level of the detected offset.

The calibration procedure for the receive array is a two stage process, wherein the reference beam for the first stage is a spread spectrum uplink signal which is received by sweeping a wide receive beam in both X and Y co-ordinates by the receive antenna to establish a coarse boresight for nominal signal weightings, and wherein the same reference beam is used for the second stage and is received by sweeping a narrow beam in both X and Y co-ordinates by the receive antenna to obtain characteristic slopes and offsets for storage by the antenna array processor and thereby variation of the corresponding signal weightings. The narrow beam may incorporate a coarse fixed offset corresponding to the offset in the X and Y phases for the coarse boresight.

According to another embodiment of the present invention a digitally controlled beam former is provided wherein the antenna array is a transmit array, wherein a reference channel is established to provide nominal coverage over a ground station, wherein a reference signal is transmitted from the spacecraft, through the reference channel, to the ground station, the reference signal being modulated by a recognition code, wherein the reference signal is swept over the ground station by the application of control signals to the amplitude and phase control elements of the N-paths of the reference channel by the antenna array processor, and wherein the signal level data received by a calibration beacon of the ground station is uplinked to, and stored by, the antenna array processor for effecting optimisation of the signal weightings applied to the reference channel and the sequential calibration of the other channels of the transmit array utilising the calibration beacon.

The calibration means for the receive and transmit arrays include switching means for each of the N-beam former channels, the switching means being adapted under the control of the antenna array processor to sequentially connect each of the channels to the reference uplink signal for calibration while the other channels are operational. The switching means for the operational channels are change over switches and the switching means for the calibration channel is a n-way switch. The switching means can be provided by high speed switch diodes, preferably in the form of monolithic microwave integrated circuits.

According to another embodiment of the present invention a digitally controlled beam former is provided which includes means for switching operation of the attitude and orbit control system (AOCS) for the spacecraft to the receive antenna array calibration means in the event of failure of the AOCS sensors, the reference channel of the calibration means being used as the AOCS channel, wherein the correlator ensures that only a spread spectrum tracking telemetry and command uplink signal from the ground station is monitored by the detector and wherein the X and Y co-ordinate data for the AOCS is provided by the antenna array processor.

The foregoing and other features according to the present invention will be better understood from the following description with reference to the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 diagrammatically illustrates a digital beam former for a spacecraft, in the form of a block diagram;

FIG. 2 diagrammatically illustrates, in the form of a block diagram, a digital beam former according to the present invention for the receive antenna array of a spacecraft;

FIG. 3 diagrammatically illustrates, in the form of a block diagram, a digital beam former according to the present invention for the transmit antenna array of a spacecraft; and

FIG. 4 diagrammatically illustrates, in the form of a block diagram, the digital beam former illustrated in FIG. 2 adapted for operation in AOCS mode.

DETAILED DESCRIPTION OF THE INVENTION

As is diagrammatically illustrated in FIG. 1 of the drawings, a digital beam former includes a beam forming network 1 having N-paths (1,2,3 . . . N) for each element (A, B and C) of the antenna array 2 of the spacecraft. Corresponding ones of the N-paths of each of the antenna elements (A,B,C) are connected to separate ones of a number of beam former channels 6. Each of the N-paths is connected to a separate one of the outputs (A.sub.1 . . . A.sub.N, B.sub.1 . . . B.sub.N, C.sub.1 . . . C.sub.N) of an antenna array processor 3 for controlling the weighting of the signals applied to the amplitude (4) and phase (5) control elements of the respective paths (A1,A2 . . . AN, B1,B.sub.2 . . . BN, C1,C.sub.2 . . . CN). The signal weightings for each element of a beam are indicative of the location on the Earth to which the antenna array is pointing. Hence, calibration of the N-paths can be effected using these weightings for a specific location or region of the Earth. Only three paths are illustrated for each of the antenna elements A, B and C but, it will be directly evident to persons skilled in the art, that any number of paths, channels and antenna elements could be employed in dependence upon the specific requirements of the spacecraft's antenna array.

The antenna elements (A,B and C) include either low noise amplifiers (LNA's) for the receive arrays, or solid state power amplifiers (SSPA's) for the transmit arrays. The phase and gain of each of these elements together with their connecting cables must be calibrated.

The antenna array elements (A, B and C) are adapted to establish beams or nulls for each of the channels 6 which may then be allocated to particular uplink, or downlink, users by the on board switching subsystem (not illustrated). The beamwidths, or null depths, and their position on the Earth are generated by the different weightings applied to the amplitude and phase control signals.

Thus, a reference uplink will require reference weightings to be applied to achieve maximum received signal level. Variation of these weightings in a calibration routine will enable the reference beam on the spacecraft to be shifted in both X and Y phases. The variation in signal level will then enable on-board software to establish any offset from the nominal beam position that is required to counter drift in the amplitude and phase of the elements in the reference path.

These offsets can then be applied to any other beam or null requirements, either as a fixed offset, or as a function derived from the slope of the characteristic obtained during the calibration routine.

As, and when, one `reference` channel is calibrated, it can be switched, in turn, to carry the traffic on each of the operational channels, whilst the elements of that channel are calibrated.

The calibration process referred to above is continuous with each channel being cycled through the calibration routine periodically, enabling short term temperature variations to be compensated.

The periodic calibration arrangement for a receive array 2A is diagrammatically illustrated, in the form of a block diagram, in FIG. 2 of the drawings. The basic structure of the beam forming network 7 of FIG. 2 is the same as the beam forming network 1 of FIG. 1 but, for the purposes of the description, only some of the connections are illustrated. In addition, one of the three channels is designated as a reference channel `R`.

As with FIG. 1, the receive array beam former is, for the sake of simplicity, shown with three N-path channels and three corresponding antenna array elements (A, B and C) only.

As is illustrated in FIG. 2, the operational channels 1 and 2 respectively include change over switches SW1 and SW2 for connecting the channel output terminals 8 and 9 either to the N-paths (AR, BR and CR) of the reference channel `R`, or one of the other channels. In the case of channel 1, the N-paths are (A1, B1 and C1) and in the case of channel 2, the N-paths are (A2, B2 and C2).

In practice, the change over switches SW1 and SW2 can be provided by high speed switch diodes, i.e. PIN diodes, in the form of monolithic microwave integrated circuits (MMIC).

The reference channel R is switched through an n-way selector switch SWR to a simple correlation/detector unit 10 comprising a filter 10A, correlation circuit 10B and detector circuit 10C connected in series between the reference channel R and an input of an antenna array processor 12. The correlation circuit 10B is connected to an input terminal 11 and the switches SW1, SW2 and SWR are each connected to separate outputs of the processor 12.

In operation, a synchronised key code from a secure processing system (not illustrated) is applied to the unit 10 via the input terminal 11 to enable correlation with the x-band command signal to be effected. The output of the detector circuit 10C is applied to the processor 12 which controls the calibration routines and the application of control signals AR, A1, BR, B1 etc to respective ones of the amplitude (13) and phase (14) control elements of the N-paths of each of the antenna elements (A, B and C).

The processor 12 also controls the calibration cycle by providing switching signals to the switches SW1, SW2 . . . SWR.

In practice, the processor 12 will, as part of the onboard autonomy of the spacecraft, contain stored data for beam forming and null pattern generation in the form of sets of control words for each channel, for example, A1, B1, C1 etc for channel 1. The control word values are varied according to the null or beam required.

In operation, the initial calibration of the reference channel R is carried out by processor 12 causing switch SWR to be set to position R, SW1 to be set to position 1, SW2 to be set to position 2 etc.

A coarse measurement is made at the commencement of the calibration routine using a spread spectrum uplink signal centred on the nominal position of the control ground station. A wide receive beam is swept in both X and Y co-ordinates by the receive antenna and a coarse boresight is established for the nominal control words, i.e. nominal signal weightings. A narrow beam is then set up incorporating, if necessary, a coarse fixed offset. The X and Y sweeps by the receive antenna are then repeated and characteristic slopes and offsets are stored. Control word offsets are then determined for each beam, or null, and are designated .DELTA.AR .DELTA.BR etc. The control words for the reference channel would, therefore, become:

On completion of the reference channel calibration process, the calibration of the first operational channel, i.e. channel 1 of FIG. 2, is then started by changing the reference channel control words for those used for the nominal channel 1 i.e. A1 B1 C1 etc.

Thus, having set up the reference path to Channel 1, the processor 12 causes switch SW1 to be switched to position R to maintain traffic, whilst switch SWR is switched to position 1 to enable channel 1 calibration to take place. The calibration procedure for channel 1 is exactly the same as the procedure used for the calibration of the reference channel R. The resulting offsets and slopes are stored in the array processor 12.

Based on this stored data, the corrections needed for the actual channel 1 operational settings are then determined and the control words are set up as follows:

The switch SW1 is then returned to position 1 by the processor 12 with traffic now being allowed to flow through the calibrated pathway whilst channel 2 is set up and calibrated in a similar manner.

The calibration procedures outlined above can be used to calibrate beam forming networks with any number of channels and antenna elements. The cycle time of the calibration process increasing with system complexity.

The periodic calibration arrangement for a transmit antenna array 2B is diagrammatically illustrated, in the form of a block diagram, in FIG. 3 of the drawings. The basic structure of the beam forming network 15 of FIG. 3 is the same as the beam forming network 7 of FIG. 2 and, as with FIGS. 1 and 2, only three N-path channels and three corresponding antenna array elements (A, B and C) are shown for the sake of simplicity.

The transmit beam former calibration procedures are basically the same as the calibration procedures for the receive beam former, but involve active participation of the control ground station (not illustrated) and the detector is part of the ground station equipment.

The transmit antenna array processor 16 is used to effect operation of the switches SW1, SW2 and SWR and to apply the weighted signals (AR, A1 . . . etc) to the corresponding amplitude (17) and phase (18) control elements of the N-paths of each antenna element (A, B and C).

A reference channel R is first set up to provide nominal coverage over the ground station. A beacon signal is then transmitted from the spacecraft to the ground station. This signal which is transmitted through the reference channel is modulated by a simple recognition code. The beam is swept by on-board generated control signals to the amplitude (17) and the phase (18) control elements, with detection data being measured on the ground. The received signal level data is then uplinked over the secure command link 19 to the processor 16 and the reference channel is optimised.

As with the receive beam former of FIG. 2, the reference channel path is then cycled, in turn, through the operational channels (1, 2, . . . etc). The operational channel paths are then calibrated using the calibration beacon with the resulting slope and offset data being calculated and stored in the array processor 16. As the required beams are selected, the appropriate offsets are calculated for the control words and the beams set up accordingly.

The transmit calibration routine will of necessity be slower than the receive calibration routine, due to the time delay inherent in transmitting signal level data from the ground station. Since a spot transmit beam can be used, the total transmit power required for the calibration beacon will be minimal, the control ground station will have a good Gain/Temperature performance and the beacon is narrow band.

The AOCS, referred to above, normally relies on input data from optical sensors, typically infra red sensors, to provide a reference to establish the attitude of the spacecraft. With infra red sensors, the edge of the Earth is detected and used as a reference point for the AOCS.

However, in the event that such sensors are disabled for any reason, then control of the spacecraft would be seriously impaired, if not, totally lost. It is, for these reasons, that much effort is being directed towards overcoming these problems.

It has been recognised that it may not be possible to make spacecraft completely immune from laser attack and alternative spacecraft altitude and orbit control systems have been proposed.

Since the calibration procedure of the present invention effectively measures the movement of boresight from the uplink transmitter position, for whatever reason, it can, therefore, be used to continuously update the AOCS with X and Y co-ordinate data. The beamwidth of this control beam can be extended to beyond Earth cover for coarse positioning data, or reduced to the minimum spot size for fine position control.

Thus, the periodically calibrated receive beam former of FIG. 2 can be modified in the manner diagrammatically illustrated, in the form of a block diagram, in FIG. 4 of the drawings for operation in the AOCS mode. The reference channel R is used as the AOCS channel.

As stated above, the basic application of the periodically calibrated beam former of FIG. 2 is to compensate for amplitude and phase drift in the antenna feed paths by measuring the apparent movement of the centre of the TT&C uplink beam from its transmitter position on the Earth. This movement could equally be caused by a change in the altitude of the spacecraft if the normal AOCS sensors are subject to interference.

Thus, in the event that the AOCS sensors are disables for any reason, the apparent shift resulting from the calibration routine being applied to the designated AOCS channel of FIG. 4, would provide the X and Y co-ordinate data for the AOCS system at the X and Y outputs of the processor 12. During this period, the accuracy of the spacecraft altitude will be dependent upon the stability of the amplitude (13) and phase (14) control elements which form part of the antenna array feed paths for the designated channel.

In order to cater for extended AOCS mode, some of the control elements of the designated AOCS beam would be temperature controlled. The number of such elements would be limited to a sub-set of those required to solely place the AOCS spacecraft receive beam over the transmitter position on Earth.

With the arrangement of FIG. 4, the use of the correlation circuit 10B of the unit 10 will ensure that only the spread spectrum TT&C uplink is monitored by the detector because any interfering signal will be reduced to insignificant levels by the narrow bandwidth of the detector.

Whilst the calibration procedures outlined above effect compensation for both amplitudes and phase drift in the antenna feed path, it may, with some systems, only be necessary to compensate for phase drift.

The primary objective of periodic calibration is to compensate temperature and life drifts of the active and passive elements in each beam forming path. As stated above, the achievement of the required stability for the paths on existing spacecraft gives rise to a temperature control requirement of .+-.2.degree. C.

Assuming that there will be a continuing requirement for similar phase and amplitude stabilities and using a maximum rate of change of temperature for payload equipments of 2.degree. C./Min, it is considered that a minimum calibration cycle time of one minute will be required.

It should be noted that 2.degree. C./Min is the normal design restraint applied to a thermal subsystem for an eclipse/sunlight change and therefore represents a worst case condition.

For a 12 channel beam former feeding a 200 element antenna array, each Complete calibration cycle represents less than 200 KBits of data, or a data processing rate of 3.3 KBits/sec for the array processor.

The transmit beam former calibration requires less than 20 KBits of signal level data per cycle. This leads to a maximum uplink data rate of 333 bits per sec on the secure command link.

For most of the operational life of the system, rates of change of temperature will be very much lower than the maximum, and hence calibration cycle times can be significantly extended. The calibration procedure could also make use of variable cycle time dependent on measured drift rates or orbital timing.



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