WO2003067351A2

WO2003067351A2 - Stabilisation and control of aircraft and other objects

Info

Publication number: WO2003067351A2
Application number: PCT/GB2003/000588
Authority: WO
Inventors: Phil Jermyn
Original assignee: Levitation Technologies Ltd.
Priority date: 2002-02-07
Filing date: 2003-02-07
Publication date: 2003-08-14
Also published as: WO2003067351A3; AU2003244415A1

Abstract

Disclosed is a method of control and stabilisation of aircraft particularly, but not exclusively, to control of miniature aircraft such as model or toy aircraft (102). In one aspect pitch and roll may be controlled by a simple method of signalling to or from a reference location orbeacon (104,106) in a number of different ways. Both height and lateral position may be controlled. In a further aspect, polarised light may be used to determine the relative rotation of the aircraft. Additionally, preferred aircraft and associated features are disclosed.

Description

STABILISATION AND CONTROL OF AIRCRAFT AND OTHER OBJECTS

The present invention relates to the control and stabilisation of aircraft and other objects. The present invention relates particularly, but not exclusively, to the control of miniature aircraft. Control and stabilisation of aircraft is a well known and well documented area of study. Many methods of aircraft stabilisation however, are complex and costly, and are not suitable for small, low cost aircraft such as model or toy aircraft. This is particularly the case for helicopters and other vertical take off aircraft. United States Patent 5 971 320 discloses a helicopter with a gyroscopic rotor assembly, having thrust generators mounted to a rotating support structure. This arrangement is suitable for model aircraft but requires expert control and continuous manual adjustment in flight.

It is an object of one aspect of the present invention to provide a method for controlling an aircraft in stable flight. It is another object of the present invention to provide for an aircraft that is capable of stable flight with minimal manual intervention. In a first aspect the invention provides a method for controlling the pitch and roll attitude of a hovering aircraft relative to a reference location, said aircraft having sensor means and said reference location emanating a beacon signal, the method comprising the steps of:

using said sensor means on the aircraft to detect the angle of incidence of a beacon signal from said reference location, relative to an aircraft fixed reference plane;

producing an angle output signal representative of the angle of incidence detected; and

using the angle output signal in a feedback control circuit, and adjusting the attitude of the aircraft using pitch and roll controls so as to obtain a desired flight characteristic.

This method of controlling an aircraft has the advantage that flight can be controlled using only a single signal beacon, or reference location. This makes the method simple and easy to set up, and has the further advantage that the sensing and controlling steps involved can be simple and inexpensive. In a preferred embodiment the beacon signal emanates from a point source emitter at said reference location. In another preferred embodiment, an emitter is located on the aircraft, and the beacon signal emanating from the reference location is a reflection of a signal from the aircraft mounted emitter. In certain embodiments it may be advantageous to derive a measure of distance of the aircraft from the reference location by using the detected amplitude of the beacon signal.

Preferably the sensing means can detect the angle of incidence of said beacon signal relative to said reference plane in two orthogonal component planes, and also preferably, two corresponding angle output signals can be produced. This is desirable since it allows a simple sensor arrangement, and the control means may be realised by using two single parameter control channels, or by using a dual parameter control channel.

This aspect of the invention can very advantageously be applied to an aircraft which develops lift thrust in a direction generally fixed relative to said aircraft fixed plane, and wherein said reference location is above said aircraft, and said sensor means is positioned on top of said aircraft. This configuration can be arranged such that the feedback control circuit responds to lateral deviation of the aircraft away from a demanded lateral position so as to cause the aircraft to move towards the demanded lateral position.

An embodiment according to such a configuration has great benefit since the control of pitch and roll attitude now further allows the aircraft to be controlled in a horizontal flight plane. An aircraft in such an embodiment will usually have different lateral response dynamics and attitude response dynamics, and typically it will be possible to optimise the feedback control circuit for either attitude control or lateral response control. The feedback control circuit is preferably optimised for attitude control. The lateral response of the aircraft may be stabilised by adding physical damping means such as sails.

Preferably, a user may control the lateral position of the aircraft by inputting an angle demand signal (which can represent a 2D horizontal position) into the feedback control circuit. In this way a user may control the position of the aircraft while the aircraft remains, to a degree, under automatic control. This is very advantageous in allowing the user to get a feel for controlling the aircraft, which for some aircraft can be very difficult. More preferably the feedback gain in the feedback control circuit can be user controlled. Thus the amount of automatic pilot assistance can be gradually reduced, and the pilot can advance progressively towards unassisted flight.

Typically, the sensing means will comprise an array of three or more sensors configured as two pairs. More typically each pairs of sensors will be aligned in a plane, and the alignment planes of each pair of sensors will be orthogonal. In this way each pair of sensors can be used to detect the angle of incident light in one plane. Furthermore, each pair of sensors will typically produce a differential output signal representative of the angle of light detected. This differential signal may be derived from either the amplitude or phase difference of a signal as detected by each of a pair of sensors. An advantage of using phase difference measurement is that angular sensitivity can be made independent of distance from the beacon.

Preferably the angle output signal produced is proportional to the angle of incidence detected over a range of angles. In one embodiment the output signal is proportional to the detected angle of incidence over a range of +/- 45 degrees.

In one embodiment, the point source emitter comprises an ultrasound emitter. Preferably the point source emitter emits ultrasound in pulses having a pulse length of approximately 1 ms, and at a frequency of approximately 50 Hz. This technique allows any echoes to die away between pulses, but provides a frequency of pulses sufficient to control the aircraft. An ultrasound emitter may be advantageous in certain applications since interfering sources of ultrasound are unlikely, and an ultrasound emitter can be small and battery operated.

In another embodiment, the point source emitter comprises an infra-red emitter. In order to overcome any potential variation in sensitivity with distance, in certain embodiments wherein infra red sensing means comprises an array of detectors, electrical automatic gain control is employed whereby signals taken from said detectors are multiplexed through a common amplifier. Infra red signalling also has advantage in that ambient interference would be unlikely in certain applications.

Although an infra red emitter has been given as a possibility, a large number of light sources distinguishable from ambient light could be used for signalling. A tungsten filament lamp, for example, could be used since the emitted spectrum contains a high proportion of infra red light.

In a second aspect the invention provides a method for controlling the height of a hovering aircraft relative to a reference location by determining the rate of change of distance of said aircraft from said reference location, and controlling the lift thrust of said aircraft so as to reduce the rate of variation in height of the aircraft.

This method is particularly applicable to model aircraft where maintaining a constant height manually can be extremely difficult. This aspect of the invention advantageously provides a simple method of stabilising the variation in height of an aircraft without the need for a demanded reference height, or even for any measurement of absolute height.

In one possible embodiment, the rate of change of distance of said aircraft from said reference location is preferably derived by differentiation of a measure of the distance of said aircraft from said reference location. In another possible embodiment, the rate of change of distance of said aircraft from said reference location is preferably derived directly, by appropriate sensing means.

In a third aspect the invention provides sensor means for detecting the direction of incident light comprising a plurality of light sensing components connected in pairs, whereby each pair is connected in a cascade arrangement such that the common node voltage varies according to the relative states of the two light sensitive components. This is a simple and effective arrangement for sensing the direction of incident light. In certain embodiments the light sensing components comprise phototransistors, which have a good light response characteristic. In an embodiment which detects the intensity of a received signal, there may be a variation in sensitivity of the sensor means with mean signal intensity. In some embodiments therefore, the common node is biased by a load impedance which comprises a light sensitive component. By selecting a certain variation of load impedance with mean intensity, the angular sensitivity of the sensor array can be made substantially constant over a wide range of mean intensities, and hence over a range of distances from the point source emitter.

In an embodiment which uses phototransistors as light sensitive components, it is advantageous to arrange four phototransistors spaced equally apart and facing radially outwards, and inclined to a mounting surface at approximately 30 degrees. This arrangement provides good linearity of incident angle vs. output.

In order to reduce interference from ambient signal sources the light sensing components may be partially enclosed in a light shielding housing. Also the sensor means may further comprise a light filter having a certain frequency response, to shield the sensors from ambient light. In a fourth aspect the invention provides for a method for controlling the pitch, roll and lateral location of a hovering aircraft relative to a reference location below the aircraft, said aircraft having first sensor means and a first point source emitter, and said reference location having second sensor means and a second point source emitter, the method comprising the steps of:

using first sensor means on the aircraft to detect a first angle of incidence of a beacon signal from said second point source emitter, relative to a reference plane on the aircraft;

producing a first signal representative of said first angle of incidence detected, which is dependent on aircraft attitude;

using second sensor means at said reference location to detect a second angle of incidence of a beacon signal from said first point source emitter, relative to a reference ground plane;

producing a second signal representative of said second angle of incidence detected, which is dependent on lateral position of the aircraft; and

using said first signal and said second signal in a feedback control circuit, and adjusting the attitude of the aircraft using pitch and roll controls so as to achieve a desired flight characteristic.

Since the reference location is below the aircraft, this novel method provides the advantage that it may be used in a large number of situations such as outdoors or from a desk mounted base. The system uses only two emitters and two sensing means, which can be robust and use components which are widely available. This method is therefore relatively simple and can be operated inexpensively. The feedback control circuit for this aspect of the invention will typically comprise two feedback loops, one loop having a lateral position signal feedback gain, and the other having an attitude signal feedback gain. In a preferred embodiment the feedback control circuit can be arranged so that the lateral position feedback gain and the attitude feedback gain can be adjusted independently. Such an arrangement is advantageous since it allows both lateral position and attitude control to be optimised independently.

In preferred embodiments of this aspect of the invention, the first point source emitter and the second point source emitter may comprise a light source, an ultrasound emitter, or a combination thereof.

Some aircraft may be prone to yaw variations, and in order to control such an aircraft according to this aspect of the invention it would be advantageous to be able to derive a measure of the yaw orientation of the aircraft.

In certain embodiments therefore one of said point source emitters further comprises a rotating polarising filter, and the corresponding receiving sensor means further comprises a fixed polarising filter; said rotating filter rotating about a generally vertical axis such that a measure of yaw orientation of the aircraft may be derived from the signal detected at said receiving sensor. This novel method is a simple and effective way of modulating one of the signals to allow it to contain aircraft yaw information which may be extracted at the adapted sensing means. In a further possible embodiment there is provided a third sensor means comprising a fixed polarising filter, said third sensor means fixed relative to said polarising emitter and adapted to provide a reference yaw signal. In this way the polarised signal detected at the polarised sensor means may be compared to derive a measure of yaw attitude of the aircraft.

In this way it can be seen that the combination of an emitter-sensor pair having a rotating polarising filter at the emitter together with a means for providing a reference signal for comparison, allows a measure of rotation of the sensor (and hence of an aircraft on which it is mounted) about an axis substantially parallel with the axis of rotation of the filter. In an alternative arrangement capable of performing the same determination of rotation of the sensor, but without the need for a separate reference signal from the emitter, there is provided an emitter with a fixed polarising filter, and a sensor arrangement comprising two light sensing components each having a fixed polarising filter. By arranging the filters at the sensor arrangement to have different orientations of polarisation, the angle or rotation about an axis substantially perpendicular to the plane of the filters can be determined by comparing the intensity of light arriving from the emitter at the two sensors.

It has been mentioned that certain aspects of the invention can advantageously be applied to an aircraft which develops lift thrust in a direction generally fixed relative to said aircraft fixed plane. Whereas the invention might be used to control a wide variety of aircraft, it will be particularly suitable for use in association with model helicopters and other hovering aircraft. Although the invention may be used to provide simple and effective control of aircraft with no inherent stability, in some applications it may be desirable that the aircraft which is to be controlled is stabilised to a degree without the need for any external input.

In one aspect the invention can be used to control a mechanically stabilised aircraft comprising:

an airframe;

one or more thrust devices mounted to said airframe, oriented to provide generally vertical thrust;

a first gyroscope rotatably mounted to said airframe, with its spinning axis oriented generally vertically and fixed relative to the airframe;

drive means to drive said gyroscope;

said gyroscope providing negligible thrust, but being sufficiently massive to provide, when spinning, direct gyroscopic stabilisation for said aircraft.

This novel arrangement provides the advantageous characteristic that such an aircraft is stabilised to a degree against pitch and roll variations without the absolute need for any orientation sensors or feedback control. Additionally the stability does not depend on the rate at which vertical thrust can be changed (feedback loop bandwidth), allowing a wide choice of propeller types.

Preferably the aircraft further comprises means for tilting the thrust direction of one or more of said thrust devices about a horizontal axis in response to a yaw disturbance of the airframe. Typically the means for effecting this tilt will comprise mechanical apparatus to provide self-stabilising negative feedback. This is desirable since it avoids using any stabilisation sensors or electronics, thus making the aircraft simpler and less costly. In a preferred embodiment for example, there is included a second gyroscope with its spinning axis oriented generally horizontally; said second gyroscope coupled to one or more of said thrust devices, such that the direction of thrust of the thrust device may be rotated about a horizontal axis in response to a yaw disturbance of the airframe. In this embodiment, any yaw disturbance experienced by the aircraft will cause the second gyroscope to precess about an axis in the horizontal plane, and this precession can be used to tilt those thrust devices to which it is coupled. The precession of the second gyro can be used advantageously to drive the thrust devices so as to provide a correcting torque to the airframe to oppose any yaw disturbances. In another preferred embodiment there is included one or more air vanes coupled to one or more of said thrust devices, such that the direction of thrust of the thrust device may be rotated about a horizontal axis in response to a yaw disturbance of the airframe. In this embodiment the air vanes will typically be pivoted such that any yawing of the airframe causes the vanes to be deflected due to the relative motion of the air. This deflection of the vanes about their pivots may be used advantageously to rotate the thrust devices so as to provide a correcting torque to the airframe to oppose any yaw disturbances.

In a variation of this embodiment, the thrust devices are fixed to the airframe, and one or more thrust vanes are disposed in the exit flow of the thrust device. In this embodiment the air vanes will typically be pivoted such that any yawing of the airframe causes the vanes to be deflected due to the relative motion of the air. This deflection of the vanes about their pivots causes one or more thrust vanes disposed in the exit flow of one or more thrust devices to tilt the direction of thrust of that device so as to provide a correcting torque to the airframe to oppose any yaw disturbance. The flight direction can typically be controlled by varying the thrust generated by one or more of said thrust generators. The aircraft will typically be controlled remotely, either via an umbilical wire or by wireless means.

The thrust devices will preferably comprise propellers, and more preferably they will be driven by electric motors, however the invention might be realised with any thrust device, such as a miniature gas turbine engine for example.

A propeller drive system however, tends to generate a propeller torque reaction. In an aircraft having multiple propellers aligned in a common direction, the standard way to cancel the propeller torque reactions is to use contra rotating propellers. This necessarily involves using two different types of propeller, which increases the number of individual components used and hence increases cost. It is desirable therefore to use propellers which all rotate in the same sense, and to cancel the propeller torques in another way.

In another aspect the invention can be used to control an aircraft comprising an airframe and two or more propellers arranged to provide vertical thrust, wherein said propellers all rotate in the same sense.

In a preferred embodiment, at least one of said propellers has its thrust direction oriented at an angle to the vertical during steady state motion so as to apply a torque to the airframe about a vertical axis. This novel arrangement is an advantageous way to cancel the propeller torque reactions without using contra rotating propellers or a tail rotor, both of which add complexity. A further advantage of this arrangement is that it allows propellers to overlap one another by virtue of the fact that the tips of the blades are displaced up or down due to the tilt, and may pass over or under each other. In still another aspect the invention can be used to control a remote controlled miniature aircraft comprising:

a generally planar airframe;

at least three propellers, oriented to provide thrust in a generally vertical direction;

a gyroscope rotatably mounted to said airframe, with its spinning axis oriented vertically and fixed relative to the airframe;

drive means to drive said propellers and said gyroscope;

a remote controller;

It is preferable that the dimensions of the airframe are significantly larger than the diameter of the propellers. In this way the propellers can be made be less exposed, and therefore are less likely to come into contact with any foreign objects. This is of great advantage in making the miniature aircraft very safe, and suitable for. indoor use for example. It is further preferable, for increased safety, that the airframe itself does not spin significantly. It is still further preferable that the propellers may be ducted or caged, again for increased safety. In this way, if the aircraft does come into contact with a foreign object it may simply bounce off, without suffering catastrophic damage, as would often be the case in prior art miniature aircraft.

Further preferred embodiments of the invention utilise a combination of some or all or the features described in the aforementioned embodiments. An embodiment of the invention will now be described by way of example with reference to the accompanying drawings in which:

Figures la to Id are schematic side views illustrating the operation of a method of flight control according to an embodiment of the present invention; Figure 2 is a circuit diagram of a preferred sensor configuration;

Figures 3 a and 3b are two views of a preferred sensor arrangement;

Figure 4 shows a preferred emitter-sensor pair arrangement;

Figure 5 shows an alternative emitter-sensor pair arrangement;

Figures 6 to 10 illustrate aspects of preferred aircraft which the present invention can be used to control.

Figures la to Id illustrate schematically a method, according to an embodiment of the preset invention, for controlling the attitude of an aircraft 102 relative to a reference location 104 having a point source emitter 106. In Figure 6a the aircraft is oriented horizontally and positioned below the reference location, with its sensing means 108 directly below the point source emitter 106. In this embodiment, when light strikes the sensor normal to the plane of the aircraft, as shown in Fig. la, the control output is zero and the attitude is not altered. Figure lb shows the aircraft positioned directly below the sensor, but at an angle to the horizontal. The sensing means 108 detects the angle at which light from the emitter strikes the plane of the aircraft, and the feedback control produces a non zero control signal. The pitch and roll controls of the aircraft are adjusted so as to restore the attitude of the aircraft plane to be perpendicular to light from the emitter 106 (i.e. to restore the aircraft to the orientation shown in Figure la) Figure lc shows the aircraft in a horizontal attitude, but moved laterally to a position such that the sensor means is not directly below the emitter. The sensor means detects the angle at which light strikes the plane of the aircraft and the control circuit uses the pitch and roll controls to adjust the attitude of the aircraft plane to be perpendicular from light from the emitter (i.e. to tilt the aircraft to the orientation shown in Figure Id). As the aircraft tilts towards the position shown in Figure Id, a sideways thrust component is generated which acts to move the aircraft laterally towards a position directly below the reference location.

Hence Figures la to Id show how in this embodiment, the control means operates to maintain the incident angle of light from the emitter to be normal to the plane of the aircraft. In this embodiment the desired flight characteristic is level flight directly below the reference location. It should be noted that in the embodiment described above, light from an emitter has been used as a beacon signal, however, any beacon signal emanating from the reference location could be used, and this could include a signal emitted from the aircraft, and reflected off the reference location, typically a ceiling. Alternatively the beacon signal could be a signal emitted from a source below the aircraft and again reflected off the reference location.

Referring to Figure 2, which is a circuit diagram of a preferred sensing arrangement, two matched phototransistors 202, 204 are connected in cascade between supply rails 206, 208, and are both connected to a common node 210. The common node is connected to a local earth via a load impedance 212. The supply rails are balanced either side of the local earth. When both phototransistors 202, 204 are illuminated equally, they both pass the same amount of current, and therefore no current passes through load impedance 212. The output 214 is at zero volts. If the two phototransistors are not equally illuminated, they will operate in different states of conductance, and will pass different currents. In this case, some current will pass through load 212, and the output 214 will swing either positive or negative by an amount dependent on the relative illumination levels of the phototransistors. The output swing as described above will depend on the difference in illumination of the two phototransistors, and their absolute illumination. Consequently there will be a variation in sensitivity with distance from a light source. This effect may be reduced by including a light dependant component, such as a light dependant resistor, in the load impedance 212. The response of the load impedance with light variation may be tailored to produce a circuit for detecting the direction of incident light, which is substantially independent of distance from the light source over a wide range of distances.

Figure 3 a is a plan view of a preferred sensor arrangement comprising four phototransistors 302. The phototransistors point radially outwards, and are equally spaced in a diamond like arrangement. Figure 3b is a side view of said preferred sensor arrangement, showing the phototransistors to be mounted on a base plate 310, and inclined at an angle of approximately 30 degrees to the base plate. Thus the phototransistors are arranged in two orthogonal planes, both of which are orthogonal to the base plate 310. Figure 4 illustrates an arrangement for determining the relative angle of rotation of a sensor arrangement 402 about an axis Y-Y, relative to an emitter arrangement 404. Emitter arrangement 404 comprises an LED 406, and a rotating polarising filter 408 mounted to a drive spindle 410, such that light emitted from the LED (indicated by arrow 412) passes through the filter. An alternative light emitting component such as a tungsten element could be used instead of LED 406.

Sensor arrangement 402 simply comprises a light sensing component 414, with a fixed polarising filter 416 through which incident light arriving at the component 414 passes.

The output from the sensor in response to the incident (rotating polarised) light will be periodic with peaks substantially every 180 degrees, corresponding to the direction of polarisation of the two filters aligning every half rotation of the emitter filter. In order to determine the relative angle of rotation of the sensor arrangement, this output signal can be compared in phase to a reference signal representative of the degree of rotation of the rotating filter, produced at the emitter arrangement. The reference signal could be obtained by, for example, providing a further sensor having a fixed polarising filter on the emitter arrangement , located opposite the rotating filter to the emitter (418). Alternatively a shaft encoder 420 could be used in conjunction with the spindle to provide a reference signal. The difference in phase of the two signals will provide a measure of the angle of rotation, and whether the sensor signal leads or lags the reference signal will determine the direction of rotation.

It can be seen that in this arrangement it is necessary to compare a signal obtained at the sensor arrangement with a signal obtained at the emitter arrangement. In order to overcome this limitation, the arrangement of figure 5 can be used to provide a similar function. An emitter arrangement 502 simply comprises a light emitting component 504 with a fixed polarising filter 506. A sensor arrangement 508 comprises two light sensing components 510, each having a fixed polarising filter 812, 814, the directions of polarisation of the two filters 512 & 514 being at different angles as shown in the figure. Since filters 512 & 514 have different directions of polarisation, the sensors 510 will detect different intensities of light depending on the orientation of the emitter arrangement about generally horizontal axis Y'-Y'. When the direction of polarisation of filter 506 lies midway between the directions of polarisation of filters 512 & 514, the sensors will detect equal intensities of light. If the emitter arrangement rotates from this position about axis Y'-Y' then one of sensors 510 will detect an increase in intensity, while the other sensor will detect a decrease in intensity. By determining the ratio of sensed intensities of the two sensors 510 therefore, a measure of rotation of the emitter arrangement about axis Y'-Y' can be determined.

Referring to Fig. 6, four electric motors 608 are arranged in a square formation, and are fixed to a rigid airframe 600. The airframe shown is one possible design, and has a cross shaped structure comprising four equal length arms. The airframe will typically be a lightweight structure and may be formed for example from a polymer material. A gyroscope 604 is rotatably mounted to the centre of the airframe 600, with its axis oriented generally vertically. The gyroscope has sufficient rotational inertia to provide direct stabilisation to the aircraft at its operational spinning speed, which might typically be about 20000 rpm. Here the gyroscope is shown enclosed by the airframe, however many different mounting configurations would be possible. The gyroscope 604 is driven by an electric motor 606, also mounted to the centre of the airframe 600. Each of the four electric motors 608, drives a propeller 610 oriented to provide generally vertical thrust. Here two bladed propellers are shown, however a wide variety of propellers may be employed. Power and control signals are supplied to the aircraft by an umbilical wire 612.

Figure 7a illustrates the principle of a first possible yaw stabilisation arrangement. One of the vertical thrust providing propellers 702, and its drive motor 704 are mounted to a portion of the airframe 706 which is rotatably connected to the rest of the airframe (represented by numeral 708) by a pivot arrangement 710. This allows the portion 706 to rotate about the generally horizontal axis X. A second gyroscope 718, comprising two inertial discs 712, 714 is attached to the portion 706 with its gyroscopic axis aligned along the generally horizontal axis Y, perpendicular to axis X. The second gyroscope is driven by its own motor 716. Any yaw disturbance of the airframe 708 (and hence also portion 706) about the generally vertical axis Z will cause gyroscope 718 to precess, and thus will cause portion 706 to rotate about axis X. This will tilt the direction of thrust of the propeller and, by choosing the correct spin direction of gyroscope 718, the propeller will generate a horizontal thrust component to oppose the yaw disturbance.

Figure 7a simply demonstrates the principle of using a horizontally mounted gyroscope to tilt a propeller in response to a yaw disturbance. It should be understood that this principle could be applied to one or more propellers using one or more gyroscopes in a variety of arrangements, within the scope of the claims.

Figure 7b illustrates the principle of a second possible yaw stabilisation arrangement. One of the vertical thrust providing propellers 742, and its drive motor 744 are again mounted to a portion of the airframe 746 which is rotatably connected to the rest of the airframe (represented by numeral 748) by a pivot arrangement 750. This allows the portion 746 to rotate about the generally horizontal axis X. In this arrangement however, the propeller 742 is a 'pusher' type propeller, and is mounted below its drive motor 744. An air vane 752 is rigidly mounted to the top of the rotatable portion 746, in the plane of the X and Z axes. Any yaw disturbance of the airframe 748 about the generally vertical axis Z will cause the air vane 752 to move along its normal direction, and the relative motion of the air (aerodynamic drag) against the vane will cause the portion 746 to rotate about axis X. This will tilt the direction of thrust of the propeller and, since the propeller is arranged as a 'pusher' type, the propeller will generate a horizontal thrust component to oppose the yaw disturbance. Figure 7c illustrates the principle of a third possible yaw stabilisation arrangement. One of the vertical thrust providing propellers 782, and its drive motor 784 are fixed relative to the rest of the airframe (represented by numeral 788). Propeller 002 is a 'pusher' type propeller. Two vanes 790, 792 are rigidly mounted to rotatable pivots 794, 796, with air vane 790 above the propeller and thrust vane 792 mounted in the exit airflow of the propeller. The pivots allow the vanes to rotate in unison about the generally horizontal axis X. Any yaw disturbance of the airframe 788 about the generally vertical axis Z will cause the air vane 790 to move along its normal direction, and the relative motion of the air (aerodynamic drag) against the air vane will cause both vanes to rotate about axis X. Thrust vane 792 will tilt and alter the direction of thrust of the propeller and, since the propeller is arranged as a 'pusher' type, the propeller will generate a horizontal thrust component to oppose the yaw disturbance. The area of the air vane is of course substantially greater than that of the thrust vane 5 Figure 8 shows a side view of the invention according to one embodiment wherein two propellers are oriented at an angle to the vertical during steady state flight. Propellers 802 and 803 are oriented to provide purely vertical thrust. Propellers 804 (foreground) and 805 (partially obscured) are oriented at a slight angle to the vertical. In this way propeller 804 generates a thrust component to the right, and 10 propeller 805 a corresponding component to the left, as viewed in the figure. Thus the airframe 806 experiences a torque acting clockwise as viewed from below the aircraft. It can be arranged therefore, for this torque to be equal and opposite to the propeller torque acting on the airframe to the propellers all being driven clockwise as viewed from beneath. 15 Figure 9 shows, for a preferred embodiment, a schematic plan view of an overlapping propeller arrangement wherein the propellers are all oriented at an angle to the vertical. Propellers A and B are tilted in opposite senses about axis C, and propellers P and Q are tilted in opposite sense about axis R such that the tips of propeller A pass over the tips of propeller Q in region 1. Vertical clearance between 20 blade tips is achieved similarly in regions 2, 3 and 4. In this way, the size, and therefore weight, of the airframe (shown broken line) can be advantageously reduced from a corresponding non overlapping arrangement having the same propeller swept area.

Figure 10 shows an example of a simple 'bedstead' type aircraft with no 25. inherent stability, which the present invention may be used to control.

The features described above can be used usefully either alone or in any combination. Although specific embodiments of the invention are provided here, it should be understood that further embodiments are included in the scope of the claims. For example, a wide variety of different signal emitters and sensors may be 30 used other than those specifically mentioned. The invention may be used to control a wide variety of different aircraft, and whilst particularly suited to miniature aircraft, many features of the present invention may be useful in full size craft. Indeed, the term 'aircraft' is used not only to include vertical thrust craft, but also winged craft, and further extends to cover hovering structures serving a wide variety of functions.

Claims

1. A method for controlling a hovering aircraft relative to a reference location, said aircraft having sensor means and said reference location emanating a beacon signal, the method comprising the steps of: using said sensor means on the aircraft to detect the angle of incidence of a beacon signal from said reference location, relative to an aircraft fixed reference plane;

using the angle output signal in a feedback control circuit so as to obtain a desired flight characteristic.

2. A method according to Claim 1, wherein said angle output signal is used in a feedback control circuit to maintain the direction of thrust of said aircraft along a line extending between the aircraft and said reference location.

3. A method according to Claim 1 or Claim 2, wherein said aircraft develops lift thrust in a direction generally fixed relative to said aircraft fixed plane, and whereby said output signal is used in a feedback control loop to adjust the attitude of the aircraft.

4. A method according to any one of Claims 1 to 3, wherein said beacon signal emanates radially over a substantial solid angle from a point source emitter at said reference location.

5. A method according to any previous claim, wherein said beacon signal emanating from said reference location is a reflection of a signal from a point source emitter.

6. A method according to Claim 5, wherein said point source emitter is mounted on said aircraft.

7. A method according to any previous claim, wherein said sensor means can detect the angle of incidence of said beacon signal relative to said reference plane in two orthogonal component planes.

8. A method according to Claim 7, wherein two angle output signals are produced; one representative of the angle of incidence detected in each component plane.

9. A method according to any previous claim, wherein said aircraft develops lift thrust in a direction generally fixed relative to said aircraft fixed plane, and wherein said reference location is above said aircraft; the configuration arranged such that the feedback control circuit responds to horizontal deviation of the aircraft away from a demanded horizontal position, so as to cause the aircraft to move towards the demanded horizontal position.

10. A method according to Claim 9, wherein the horizontal response dynamics and attitude response dynamics of the aircraft are different, and whereby the feedback control circuit is optimised for attitude control.

11. A method according to Claims 9 or 10, wherein a user can control the demanded horizontal position of said aircraft by inputting an angle demand signal into the feedback control circuit.

12. A method according to Claim 11 , wherein said demand signal represents a two dimensional horizontal position.

13. A method according to any previous claim, wherein said sensor means comprises an array of three or more sensors configured as two pairs.

14. A method according to Claim 13 wherein each pair of sensors detects the angle of incidence of said beacon signal in a component plane, the respective component planes being orthogonal.

15. A method according to Claim 13 or 14 wherein each pair of sensors produces a differential output signal.

16. A method according to any previous claim wherein the feedback gain in the feedback control circuit can be user controlled.

17. A method according to any previous claim, wherein said angle output signal produced is representative of the angle of incidence detected over a range of angles.

18. A method according to any previous claim, wherein said point source emitter comprises an ultrasound emitter.

19. A method according to Claim 18, wherein said point source emitter emits ultrasound in pulses having a pulse length of approximately 1 ms, and at a frequency of approximately 50 Hz.

20. A method according to any previous claim, wherein said point source emitter comprises an infra-red emitter.

21. A method according to any previous claim, wherein said point source emitter comprises a light source distinguishable from ambient light.

22. A method according to any previous claim, wherein said sensor means comprises an array of detectors, and where the signals taken from said detectors are multiplexed through a common amplifier.

23. A method according to any previous claim, wherein a measure of distance of said aircraft from said reference location may be derived from the detected amplitude, phase or timing of said beacon signal.

24. A method for controlling the height of a hovering aircraft relative to a reference location by determining the rate of change of distance of said aircraft from said reference location, and controlling the lift thrust of said aircraft so as to reduce the rate of variation in height of the aircraft.

25. A method according to Claim 24, wherein the rate of change of distance of said aircraft from said reference location is derived by differentiation of a measure of the distance of said aircraft from said reference location.

26. A method according to Claim 24, wherein the rate of change of distance of said aircraft from said reference location is derived directly by appropriate sensing means.

27. A method for controlling the pitch, roll and lateral location of a hovering aircraft relative to a reference location, said aircraft having first sensor means and a first point source emitter, and said reference location having second sensor means and a second point source emitter, the method comprising the steps of:

producing a first signal representative of said first angle of incidence detected, which is dependent on aircraft attitude, and upon lateral position of the aircraft;

using second sensor means at said reference location to detect a second angle of incidence of a tracking signal from said first point source emitter, relative to a reference ground plane; producing a second signal representative of said second angle of incidence detected, which is dependent on lateral position of the aircraft; and

28. A method according to Claim 27, wherein said feedback control circuit comprises two feedback loops, one loop having a lateral position signal feedback gain, and the other having an attitude signal feedback gain.

29. A method according to Claim 28, wherein the lateral position feedback gain and the attitude feedback gain can be adjusted independently.

30. A method according to any one of Claims 27 to 29, wherein at least one of the beacon signal and the tracking signal is plane polarised with changes in the angle of polarisation or in the angle of polarisation sensitivity of the sensor being used in a measure of aircraft yaw.

31. A method according to any one of Claims 27 to 30, wherein for one of said emitter-sensor pairs, a rotating polarising filter, rotating about a horizontal axis, is mounted to one of said emitter or said sensor, and a fixed polarising filter is mounted to the other, such that a measure of yaw orientation of the aircraft may be derived from the phase or frequency or the signal detected by said adapted sensor.

32. A method according to Claim 31, further comprising a third sensor means comprising a fixed polarising filter, said third sensor means fixed relative to said polarising emitter and adapted to provide a reference yaw signal.

33. A method according to any one of Claims 27 to 32, wherein said first point source emitter and said second point source emitter may comprise a light source, an ultrasound emitter, or a combination thereof.

34. A method for detecting the rotation about an object rotation axis of an object located at a direction from a remote location; comprising the step of establishing between the object and the remote location a beam of radiation that is plane polarised in the plane orthogonal to said direction and detecting said radiation with a polarisation sensitive detector.

35. A method according to Claim 34, wherein two polarisation sensitive detectors are provided to produce a differential detection signal which represents the degree of rotation of the object about the object rotation axis.

36. A method according to Claim 34 or Claim 35, wherein the angle of the beam polarisation or of the polarisation sensitivity of the detector, rotates about said direction to produce a detection signal the phase of which represents the degree of rotation of the object about the object rotation axis.

37. A method according to Claim 36, wherein the angle of the beam polarisation rotates about said direction and wherein a reference polarisation sensitive detector serves to produce a reference signal the phase of which is independent of the degree of rotation of the object about the object rotation axis

38. Apparatus for detecting the rotation about an aircraft rotation axis of an aircraft located at a direction from a remote location, comprising:

an emitter capable of emitting radiation plane polarised in the plane orthogonal to said direction, which emitter is fixed relative to the reference location; a polarisation sensitive detector mounted on said aircraft; and means for producing a signal from said detector in response to radiation from said emitter, representative of said rotation of the aircraft.

39. Apparatus according to Claim 38, wherein said polarisation sensitive detector comprises two sensors having different sensitivities to plane polarised radiation in a given polarisation direction, and wherein said signal is a differential signal from said two sensors, representative of the degree of rotation of said aircraft about said rotation axis.

40. Apparatus according to Claim 38, wherein said emitter further comprises means for emitting radiation, the direction of polarisation of which rotates about said direction.

41. Apparatus according to Claim 38, wherein the polarisation sensitivity of said detector rotates about said direction.

42. Apparatus according to Claim 40 or Claim 41, further comprising a reference polarisation sensitive detector capable of producing a reference signal which is independent of the degree of rotation of the object about the object rotation axis.

43. Apparatus for determining the yaw attitude of an aircraft relative to a reference location, wherein one of a transmitter and sensor pair is mounted on said aircraft and the other is fixed relative to said reference location; characterised in that a rotating polarising filter, rotating about a horizontal axis, is mounted to one of said transmitter or said sensor, and a fixed polarising filter is mounted to the other, such that a measure of yaw orientation of the aircraft may be derived from the phase or frequency of the signal detected by said adapted sensor.