Note: Descriptions are shown in the official language in which they were submitted.
13 41555
STEERING OF MISSILES
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,
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(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
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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
Figure 1 is a perspective view of a missile according to
the invention;
Figure 2 is a view from one side of a forward part of
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the missile of Figure 1, partly cut away to reveal
details of a slip clutch; and
Figure 3 is a block diagram of the steering means which
controls the slip clutch.
DESCRIPTION OF THE PREFERRED EMBODIMENT
Figure 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.
13 4 15 5 5
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It will be appreciated from the foregoing that flight
of the missile is controlled in canard fashion.
Figure 2 shows in somewhat more detail the connection
5 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 er 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 Figure 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 Figures 1 and 2 is guided
along a plane polarised, pulsed laser beam emanating
4
= 4 3
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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 PoLGfilters. These
photodiodes respond to the laser beam and produce
electrical signals used in steering the missile, as
shown schematically in Figure 3.
In Figure 3, a first photodiode 30 and second photo-
diode 31, each having a sensitive area of 51mm 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 cos2e (where A 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 0.5A/W, the maximum output
is 3 x 10-4A and the minimum is 5 x 10-$A. When 6= 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
and 31 within the laser beam. More particularly,
the signals from the photodiodes carry information
sufficient to establish a radial discrepancy R of
30 the longitudinal axis of the missile from a notional
guidance axis at the centre of the laser beam and
an error angle 9 E representative of the direction
in which the axis of the missile lies relative to
the notional guidance axis.
- 7 -A~ An analogous arrangement is shown in United States
Patent Specifleatien No. 3957377.
As shown in Figure 3, 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 AB 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
eE'
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 0 B of the missile
body relative to space. It does not identify a unique
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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 9 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:gu.idance circuitry, (typically of complexity
4 op-amps),.which examines the error angle 9E and
determines what angle 9G 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 ADNB relative to the body of the missile 11 having
regard to the space roll angle 9B 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 9E and the radial
error R. From these inputs it determines what is
the required missile nose space angle 8G and provides
this as input to an adding circuit 40.
The body angle logic circuit 38 examines how the instant-
aneous radial error R and the rate of change, R, in
0
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 unambig-
uous missile body space roll angle 9B as input to
the adding circuit 40. This last circuit provides,
-9= 1341555
as an analogue output from the guidance circuitry,
a signal representative of a demanded relative angle
9DNB 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 9NB
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 6B, 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 9 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
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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.
