Abstract:
A beam-riding missile ( 10 ) has a freely rotating control portion ( 11 ) forming its nose and carrying a pair of fixed ailerons ( 13 ) and a pair of fixed elevators ( 14 ). Detecting means (not shown) gather information indicative of the location of the missile in the beam and steering logic circuitry (not shown) provides signals to a clutch ( 18 ) which interfers with the free rotation of the nose in such a way that the elevators are effective to maintain the chosen flight path. 
     The clutch can be electromagnetic, piezo-electric or function on the Johnson-Raebeck effect. 
     The combination of fixed control surfaces and steering by a single actuator leads to the possibility of useful reductions in the size, weight and complexity of the missile.

Description:
FIELD OF THE INVENTION 
   This invention relates to the steering of missiles. It is particularly, but not exclusively, concerned with small, aerial missiles having an elongate body with main, fixed flight surfaces which cause the body to rotate in one direction during flight of the missile, and a relatively small nose portion which tends to rotate in the opposite direction during the flight of the missile. 
   SUMMARY OF THE INVENTION 
   According to the present invention there is provided a missile suitable for controlled flight through a fluid medium having an elongate body portion of relatively high inertia and a control portion of relatively low inertia which can rotate freely on the body portion about the longitudinal axis of the missile, wherein:
         (1) the control portion has an aileron which is fixed at a predetermined and constant angle of incidence so that, in flight of the missile, the force of reaction between the aileron and the fluid medium gives to the control portion a tendency to rotate within the fluid medium,   (2) the body portion is provided with control means which induce in the body portion a rate of change of roll angle of the body portion relative to the fluid medium which is different from that of the control portion,   (3) the control portion includes an elevator which is fixed at a predetermined and constant angle of incidence to react at all times during the flight of the missile against the fluid medium incident upon it to impose an instantaneous lateral force on the missile,
 
and the missile includes
   (1) detecting means for generating an error signal indicative of a discrepancy between an instantaneous flight path of the missile and a chosen flight path, and   (2) steering means comprising steering logic responsive to said error signal for generating a missile steering signal and a clutch responsive to the steering signal for limiting the free rotation between the body portion and the control portion of the missile such that, in response to the error signal, the steering means biases the control portion towards that roll angle at which the transverse force imposed on the missile by the elevator is such as to reduce said discrepancy.       

   In the miniature missiles for which the present invention has particular application, it may be convenient for the control portion to be embodied as a relatively small nose section of the missile, which nose may essentially comprise a pair of fixed ailerons at opposite ends of a first transverse diameter, a pair of fixed elevators at opposite ends of a second transverse diameter, itself transverse to the first and a mass of dense metallic material as a charge to be delivered to the target by the missile. 
   The body of the missile, on the other hand, may contain one or more gyroscopes for maintaining the missile stable and possibly assisting in its guidance. One convenient way of defining the chosen flight path is to provide a beam, such as a laser beam, emanating from a missile guidance station. Although a laser is preferred, other coherent, electromagnetic radiation may be suitable for the beam. Sensors on a rearward-facing surface of the missile feed sufficient information about the position of the missile within the beam to steer the missile and keep it within the beam. 
   Missiles according to the invention may be employed as sub-missiles in the invention disclosed in our co-pending British Patent Application No. 8132088, in which Application the small size which can be achieved in the missiles of the present invention is of prime importance. 
   For a better understanding of the invention, and to show more clearly how the same may be carried into effect, reference will now be made to the accompanying drawings in which: 

   
     BRIEF DESCRIPTION OF THE DRAWING 
       FIG. 1  is a perspective view of a missile according to the invention; 
       FIG. 2  is a view from one side of a forward part of the missile of  FIG. 1 , partly cut away to reveal details of a slip clutch; and 
       FIGS. 3   a  and  3   b  are a block diagram of the steering means which controls the slip clutch. 
   

   DESCRIPTION OF THE PREFERRED EMBODIMENT 
     FIG. 1  shows a missile having a body  10  and a nose section  11 . Four main flight surfaces  12  are provided at the rear of the body  10  and are so oriented that the body  10  has a tendency to rotate in a clockwise direction (viewed from the front of the missile) during normal flight, as indicated by arrow F. 
   A nose section  11  of the missile is freely rotatable relative to the body portion  10  about the longitudinal axis of the missile. It carries a pair  13  of fixed ailerons at opposite ends of a transverse diameter of the nose, these giving the nose  11  a tendency to rotate in normal flight of the missile in a direction shown by arrow f counter to that of the body portion  10  in normal flight of the missile. A pair of elevators  14  fixed on the nose section at a small angle of incidence and located at opposite ends of a diameter transverse to that containing the ailerons  13  imposes on the missile a transverse force i.e. one in directions transverse to that of its flight. During such time as the rotation of the nose  11  is free there is no resultant unidirectional transverse steering force on the missile. However, when the free rotation is interrupted, the resultant force will accelerate the missile in a direction transverse to its length. 
   It will be appreciated from the foregoing that flight of the missile is controlled in canard fashion. 
     FIG. 2  shows in somewhat more detail the connection of the nose  11  and the body  10 . An axial shaft  15  of the nose extends rearwardly into the body  10  and is carried therein by a forward ball race  16  and a rearward ball race  17 . A conventional electromagnetic clutch, referenced generally  18 , is employed to interfere with free rotation of the nose  11  relative to the body  10  in a manner known per se. The clutch  18  comprises an annular coil  19  through which an electric current may be flowed to generate an electromagnetic field which interacts with an armature  20  on the nose  13  to resist rotation of the nose  13  relative to the coil  19 . Electrical current is supplied to the coil  19  by a steering means, not shown in  FIG. 2 , which varies this current with time in such a way as to interfere with the free rotation of the nose at times when a steering correction of the missile is required. This interference introduces a disparity between the length of time which the elevator surfaces  14  occupy in one angular position of the nose and the time during which they occupy other angular positions i.e. it biases the elevators towards a selected angular position thereby to accelerate the missile transversely as necessary to correct the path of its flight. In the limiting case, the current through the coil is such as to maintain the angular position of the nose fixed in relation to the environment of the missile for long enough to achieve the necessary steering connection. 
   The missile illustrated in  FIGS. 1 and 2  is guided along a plane polarised, pulsed laser beams emanating from a missile control station. The length of each of the laser pulses is conveniently 100 ns. On a rearward-facing surface of the missile are provided pin photodiodes having crossed polarising filters. These photodiodes respond to the laser beam and produce electrical signals used in steering the missile, as shown schematically in  FIGS. 3   a  and  3   b.    
   In  FIG. 3   a , a first photodiode  30  and second photodiode  31 , each having a sensitive area of 5½ mm diameter, generate electrical signals when the laser beam is incident upon them, these signals constituting inputs to the remaining components of the steering means of the missile. The transmittance of the polarisers when crossed with the laser beam is 3% and when parallel is 45%. The output current from each photodiode is proportional to cos 2  θ (where θ is the angle between the plane of polarisation of the laser beam and that of the polariser on the photodiode). The responsivity of each photodiode cell is 0.5 A/W, the maximum output is 3×10 −4  A and the minimum is 5×10 −8  A. When θ=45° for each of the two polarisers, the transmittance of each is the same, at 25%. 
   The laser beam is modulated in such a way that the inputs vary according to the position of the diodes  30  and  31  within the laser beam. More particularly, the signals from the photodiodes carry information sufficient to establish a radial discrepancy R of the longitudinal axis of the missile from a notional guidance axis at the centre of the laser beam and an error angle θ E  representative of the direction in which the axis of the missile lies relative to the notional guidance axis. 
   An analogous arrangement is shown in U.S. Pat. No. 3,957,377. 
   As shown in  FIG. 3   a , the photodiodes  30  and  31  have crossed polarising filters and so, as shown in the drawing, with a polarised laser beam, a comparison of the signals emanating from the photodiodes establishes a roll angle θ B  of the missile body  10  relative to the plane of polarisation of the laser beam. 
   The diodes  30  and  31  provide inputs to amplifiers  32  and  33  respectively, these constituting photodiode bias and pre-amp circuitry which typically has a complexity in a range of from 2 to 4 op-amps. The analogue outputs from the amplifiers  32  and  33  provide two inputs to each of an adding circuit  34  and a roll angle circuit  35 , these two circuits together performing a function of missile roll angle and pulse train extraction and typically having a complexity of 2 op-amps. 
   The adding circuit  34  provides as a digital output a series of pulse trains  36  which series is representative of the pulsed laser guidance beam received by the diodes  30  and  31 . The laser beam is so modulated that the duration of the pulses  36  which the adding circuit  34  produces as its output is representative of the said radial error R. The frequency of repetition of the pulses  36  is representative of the error angle θ E . 
   The pulses  36  are fed to a pulse decoding circuit  37 . The output from the roll angle circuit  35  provides information as to the roll angle θ B  of the missile body relative to space. It does not identify a unique roll angle but rather one of two roll angles spaced apart by 180°. The output of the roll angle circuit  35  is fed to a body angle logic circuit  38 . The roll angle circuit  35  also generates an automatic gain control (AGC) signal which is fed to the amplifiers  32  and  33  where it serves to ensure their linear operation. So long as the amplification is linear, the body roll angle θ B  is determinable by comparison of the magnitude of the outputs of the amplifiers  32  and  33 . 
   The circuits  37  and  38  are components of digital logic, beam-riding guidance circuitry, (typically of complexity 4 op-amps), which examines the error angle θ E  and determines what angle θ G  of the nose section  11  of the missile in space is needed to rectify the error. The desired space angle of the nose section  11  is achieved by securing a desired angle of the nose section  11  θ DNB  relative to the body of the missile  11  having regard to the space roll angle θ B  of the missile body. 
   Thus, the guidance circuitry comprises beam-riding guidance shaper circuitry  39  which receives from the pulse decoding circuit  37  an input signal indicative of the missile body error angle θ E  and the radial error R. From these inputs it determines what is the required missile nose space angle θ G  and provides this as input to an adding circuit  40 . 
   The body angle logic circuit  38  examines how the instantaneous radial error R and the rate of changes,            , in R vary in consequence of a guidance command and, from this information, inverts the signal from the roll angle circuit  35  when necessary, to provide an unambiguous missile body space roll angle θ B  as input to the adding circuit  40 . This last circuit provides, as an analogue output from the guidance circuitry, a signal representative of a demanded relative angle θ DNB  between the body of the missile and the nose of the missile.
   This output is fed to analogue nose roll loop circuitry comprising a shaper circuit  41  (which is typically of 3 op-amps complexity) which compares the demanded angle with a signal derived from a voltage divider  42  which is representative of the actual angle θ NB  between the nose and the body of the missile. At such times when the longitudinal axis of the missile is coincident with the notional guidance axis there will be zero output from the guidance shaper circuit  39  so that the adding circuit  40  will merely feed to the nose roll loop shaper  41  a cyclical signal indicative of θ B , i.e. the steady rotation in space of the missile body  10 . In these circumstances the circuit  41  produces zero output. 
   On the other hand, whenever there is a radial discrepancy R between the axis of the missile body  10  and the notional guidance axis, the adding circuit  40  will produce a signal which causes the circuit  41  to produce an output amplified by a drive amplifier  43  for operating the clutch  18  between the missile body  10  and the nose  11  to procure a demanded nose body angle θ DNB . 
   The clutch need not be an electromagnetic device such as is shown in the illustrated embodiment. It can be, for example a piezo-electric device which responds to the passage of electric current therethrough to expand along one axis and thereby exert a frictional resistance to the free rotation of the nose portion on the body of the missile. Again, a clutch member may utilise the Johnson-Raebeck effect whereby a material such as agate undergoes a change in its coefficient of friction when it is subject to electrical stress. A suitable device for use as the clutch  18  which utilises this effect is made by M.L. Aviation Limited, whose address is White Waltham Aerodrome, Maidenhead, Berkshire. 
   Information about the roll angle of the missile body in space can be obtained from a roll gyroscope on board the missile instead of from a laser beam guidance signal.