Stabilization of radio controlled aircraft

A stabilization system is disclosed in which stabilization pitch and roll signals effective for stabilizing the aircraft are combined with respective pilot provided elevator and aileron position demand signals in accordance with a function which reduces the effects of the stabilizing signals in dependence on increasing values of the respective elevator and aileron position demand signals.

This invention relates to stabilisation of the attitude of a radio 
controlled aircraft in pitch and roll. 
Flying radio controlled model aircraft requires training and the 
acquisition of skill. Special easy to fly training aircraft are used for 
novices in order to reduce the risks of a crash. Especially during early 
training there is a risk that the novice pilot will create a situation 
which s/he has insufficient skill to recover from. Even training models 
will merely continue at the existing attitude if the controls are 
neutralised so positive action needs to be taken to recover. Radio 
failure, or interference will, or can, lead to a neutral setting of the 
controls so that the model just continues as it was before the 
interference. Even training models can be expensive to replace and a crash 
always carries an element of risk to personnel. 
U.S. Pat. No. 2,828,930 describes a system for stabilising the attitude of 
an aircraft. If the system were applied to a radio controlled aircraft, 
however, the stabilisation system would continually resist attempts by the 
pilot to control the aircraft's attitude to perform manoeuvres. 
Against this background, in accordance with the invention, there is 
provided a radio controlled aircraft, including a radio receiver for 
receiving an elevator position demand signal for specifying a required 
position for the elevator and at least one aileron position demand signal 
for specifying required positions for the or one of the ailerons; 
stabiliser means for generating pitch and roll stabilisation signals 
dependent on differences in the aircraft's attitude from level in pitch 
and roll respectively, mixer means for combining the stabilisation pitch 
and roll signals with the elevator and aileron position demand signals in 
a sense to stabilise the aircraft in accordance with a function which 
reduces the effect of the stabilisation signals in dependence on 
increasing values of elevator and aileron position demand signals 
respectively; and control means for controlling the positions of the 
elevator and ailerons in accordance with respective combined signals. 
Importantly, if the pilot removes all input by neutralising the position of 
the stick, the stabilising signals will bring the aircraft into level 
flight. Small stick inputs will still be resisted by the stabilisation 
signals. However, large stick inputs will be less affected. 
In one form, the values of the modified pitch and roll stabilisation 
signals are proportional to those of the (unmodified) pitch and roll 
stabilisation signals below respective thresholds and equal to zero above 
said thresholds. 
More preferably, the values of the pitch and roll stabilisation signals are 
reduced in proportion to any increase in the respective elevator and 
aileron position demand signals from a maximum when the respective 
position demand signal is zero, to zero when the respective position 
demand signal has a predetermined value larger than zero, e.g. 
corresponding to maximum deflection of the elevator or ailerons. 
Most preferably, the radio receiver is adapted for receiving a gain setting 
signal to which the means for modifying the pitch and roll stabilisation 
signals is responsive to provide a level of gain dependent on the gain 
setting signal. 
In one convenient form, two pairs of directional light sensors are 
provided, the sensors of each pair being responsive to light input from 
two different directions on opposite sides of a respective axis to provide 
respective light level signals indicative of the level of light sensed 
said stabiliser means being responsive to difference between the light 
level difference signals to provide difference signals. 
The signal from each sensor will vary in value dependent on whether it is 
directed above or below the horizon. When the aircraft is level and the 
two sensors are directed at the horizon, the difference should be zero. 
The light sensors are preferably photo conductive light sensors connected 
in series across a reference voltage. The light sensors may be aligned 
approximately normal to respective pitch and roll axes. 
In order to reduce the possibility of a sensor being incapacitated by a 
deposit of oil or other pollutant it would be preferable if none of the 
sensors faced directly in the direction of motion. In a preferred 
arrangement, the light sensors are aligned in directions approximately 
bisecting the angles between the pitch and roll axes. 
In this arrangement, in order to derive suitable stabilisation signals 
there is preferably included means to add the differences in order to 
provide one of the pitch and roll stabilisation signals and to subtract 
one difference from the other in order to provide the other of the pitch 
and roll stabilisation signals.

Referring to the drawings, the apparatus is intended to stabilise the 
attitude of a model aircraft about two axes so as to control the pitch and 
the angle of bank of the aircraft. A sensor unit 2 shown in FIG. 1 is has 
the external semblance of a model radome. The unit has a body 4 and a 
cover 6. Four small holes 8 are provided in the periphery of the body 
aligned with corresponding holes 10 on the cover. Two of the holes 8,10 
are aligned diametrically opposite on one axis A--A. The other two holes 
8,10 are aligned diametrically opposite on an axis B--B normal to the axis 
A--A and generally in the same plane. 
The body 4 has a central chamber 12 into which the holes 8 extend. A 
photoconductive light sensor (photo resistor) 14 is mounted in the inner 
end of each hole 8 so as to sense light entering from the outer end of the 
hole. In the particular example the holes are about three times as long as 
their diameter, more preferably 2 to 21/2 times, so that each photo 
resistor responds to light from the general direction of the axis A--A or 
B--B. In an even more preferred example the outer end of the holes is not 
circular but wider in the horizontal plane than in the vertical plane. For 
example, the holes may be 4 mm diameter round at the sensors and 6 mm wide 
by 2.5 mm high at their outer ends. More light is thus collected from a 
wider arc of the horizon. 
The photo resistors are more sensitive to the blue and ultra violet end of 
the spectrum than to the red and infra red, so that the background 
illumination level normally increases above the horizon and reduces below 
the horizon. This is to avoid hot ground masking, or counteracting, the 
increase of illumination level above the horizon. Either the photo 
resistors may be selected to be more sensitive to the blue/ultra violet 
end of the spectrum or, if light sensors are used which are undesirably 
sensitive to the red/infra red end of the spectrum, filters may be used to 
filter out the undesired part of the spectrum. 
The sensor unit 2 is installed on the aircraft so that when it is in flying 
in a level attitude the axes A--A and B--B are directed at the horizon. If 
the aircraft's attitude is different from that, then one or both of the 
axes A--A and B--B will no longer be aligned with the horizon. Supposing 
the axis A--A is no longer aligned with the horizon, then one of the 
respective photo resistors will be directed above the horizon and will 
receive more light than the other which will be directed below the 
horizon. 
A cable 16 and connector 18 (FIG. 1) connect the photo resistors as shown 
in FIG. 2. The two photo resistors 14 on one axis, e.g. A--A are connected 
in series across a voltage supply. The node between the photo resistors is 
connected to signal processing means 19 at an analog to digital convertor 
20 in which an analog to digital conversion is performed. When the axis 
A--A is horizontal, the respective photo resistors 14 are directed at the 
horizon and receive nominally the same level of light so that the node 
between them is at a voltage level half way between the supply rails. If 
the axis A--A is inclined in one direction to the horizon, the upwardly 
directed photo resistor receives more light than the downwardly directed 
photo resistor. The resistance of the upwardly directed photo resistor 
falls and that of the downwardly directed photo resistor rises so that the 
voltage at the node between them rises. 
The voltage at the node between the two photo resistors therefore rises and 
is converted into a digital light level difference signal S1, S2 by the 
analog to digital converter 20. The light level signal is transmitted to a 
signal shaping unit 22. With switches 24 set in the position shown, the 
light level signals S1 and S2 are each input directly with a digital 
reference to a subtractor 26. The digital reference signal is equivalent 
to equal light being received by the photo resistors. The output 
difference signal Rs represents angular deflection of the aircraft about 
the roll axis and the output difference signal Ps represents the angular 
deflection of the aircraft about the pitch axis. 
The sensor unit 2 should be placed on the aircraft so that it is nominally 
horizontal in level flight. Provided the background light level is uniform 
in a plane parallel to the horizon, the digital reference signal could be 
set permanently. However, in reality, the background light level will not 
be uniform and the sensor unit 2 may not be completely accurately aligned. 
To accommodate these practical problems, the digital reference signals may 
be adjusted on the ground before flight and in flight via a respective 
radio channel. 
If the axis A--A is inclined in the other direction to the horizon, the 
voltage at the node falls. 
The direct comparison of the light level signals S1 and S2 with the 
reference signal is appropriate when the axes A--A and B--B are aligned 
with the pitch and roll axes of the aircraft. Especially in the form 
illustrated where the sensing unit has open holes 8 to direct the light 
sensors, it may be found that a forwardly directed hole is subject to 
blocking, e.g. with engine oil. In this case it would be an advantage if 
none of the holes is aligned in the direction of flight. 
To this end the axes A--A and B--B may be aligned at 45.degree. to the 
pitch and roll axes of the aircraft, i.e. in directions parallel to the 
bisectors of the angles between the pitch and roll axes. With this 
arrangement, the switches 24 are operated to their alternative positions 
so that the signals input to the subtractors 26 result from an adder 28 
and a subtractor 30. The adder 28 produces a signal having a value one 
half of the sum of S1 and S2. The sum signal is input to the respective 
subtractor 26 to produce the signal R which represents the attitude of the 
aircraft about the roll axis. The subtractor 30 produces a signal which 
has a value one half the difference between S1 and S2. The difference 
signal is input to the respective subtractor 26 to produce the signal P 
which represents the attitude of the aircraft about the pitch axis. 
The pitch and roll representative signals P and R, are dynamically shaped 
by respective differentiators 32 and adders 34 which add a differental 
component to produce proportional/differential stabilising signals 
##EQU1## 
where A is gain. 
The model is controlled by radio signals received by a receiver 36. The 
receiver decodes the received signals and places pulse width modulated 
signals on output lines 38 to 46, one for each channel. Position demand 
signals Pd on line 38, Rd1 on line 42 and Rd2 on line 46 are intended to 
control the position of respective servos 48, 50 and 52 which in turn 
control the aircraft's elevator (controlled by servo 48) and ailerons 
(controlled one each by servos 50 and 52) to control the attitude of the 
aircraft. Separate controls for each aileron allow their use also as 
flaps. Conventionally the output lines 38 and 42 and 46 would be connected 
direct to the respective servo. As shown in FIG. 2, however, the lines are 
connected to the controller 19 which converts the pulse width modulated 
signals to digital signals in convertors 54. 
The stabilising signals 
##EQU2## 
may be adjusted, as hereinafter described, by gain setting multipliers 60 
and by invertors 62 under control of switches 64, and the adjusted signals 
are applied to the mixers 56. 
Mixers 56 act, broadly, to modify the digital demand signals Pd, Rd1 and 
Rd2 increasing or reducing them in a sense to reduce the value of the 
respective stabilising signals 
##EQU3## 
The modified demand signals Pm, Rm1 and Rm2 are converted to pulse width 
modulated signals by convertors 58 and drive the respective servos 48, 50 
and 52. 
In use, if the control sticks on the transmitter are neutralised, so that 
in the prior art arrangements the aircraft would continue doing what it 
had been doing, the modification to the pulses produced by the controller 
19 will return the aircraft to a level attitude in both axes and the right 
way up. However, if the mixing function is mere summation, every demand 
input intended by the pilot will be resisted or counteracted by the 
stabilising signals. 
To improve the response of the system to pilot inputs without reducing the 
effectiveness of the stabilisation system if the pilot looses control, the 
mixing function of the mixers 56 broadly reduces the effect of the 
stabilisation signals in dependence on increasing values of elevator and 
aileron position demand signals respectively. 
Broadly, for each mixer: 
EQU U.sub.output =U.sub.input1 +U.sub.input2 +F(U.sub.input1,U.sub.input2,P) 
where: U.sub.output is the output signal from the mixer 
U.sub.input1 is the stabilisation signal appearing at one input to the 
mixer 
U.sub.input2 is the pilot demand signal appearing at the other input to the 
mixer 
P is a parameter defining the mixing law and which may be preset or derived 
from a further channel output of the receiver 36 on one of lines 40, or 
44. 
The effect of the stabilisation signal may be reduced proportionally, for 
example: 
EQU F(U.sub.input1,U.sub.input2,P)=-{U.sub.input1 *(U.sub.input2 
/maxU.sub.input2)} 
where max U.sub.input2)} is the value of U.sub.input2)} which gives maximum 
control deflection. 
In another example, the function may be defined as follows: 
EQU F(U.sub.input1,U.sub.input2,P)=0 if U.sub.input2 &lt;maxU.sub.input2 /2 
EQU F(U.sub.input1,U.sub.input2,P)=(a negative sign)U.sub.input1 if 
U.sub.input2 &gt;maxU.sub.input2 /2 
The gain set by the multipliers 60 may be derived from an on board 
potentiometer 66 via analog to digital convertor 68, or from a signal 
transmitted over a further channel (not shown). 
If the aircraft is inverted, the modifications to the pulse width modulated 
signals will be in the wrong sense to return it to an inverted level 
attitude so it flies back the right way up. 
The sensor unit may be mounted above or below the aircraft. However the 
effect of the control unit 19 would need to be reversed according to where 
the sensor unit is mounted. Supposing the sensor extends the servo pulse 
width if the sensor unit is mounted above the aircraft in the orientation 
illustrated in FIG. 1, then it would need to produce a reduction in pulse 
width if the sensor were mounted below the aircraft inverted compared with 
FIG. 1. Switches 64 and invertors 62 are provided order to preset the 
controller to accommodate either orientation of the sensor unit. 
The functions of the signal processing means may be conveniently 
implemented using a programmed microprocessor. 
In alternative arrangements, the stabilisation signals may be derived from 
conventional gyros.