GB1605347A - Missile guidance systems - Google Patents

Description

(54) MISSILE GUIDANCE SYSTEMS VWe, BsH AERosPAcr PUBLIC LIMITED COMPANY, a British Company organised under British Aerospace (Nominated Company) of 100 Pall Mall, London SWIY 5HR, do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to missile guidance systems, and is concemed with sighting and tracking arrangements for such systems.

In its broadest concept, the presenl invention comprises a guidance system for guided missile and includes an optical sight; a manually.

operable control for generating manual control signals in response to manipulation thereof by an operator; a tracker having an optical system for tracking a missile in flight in the field of view of the tracker optical system, the tracker being constructed and ananged to generate error signals representative of the radially-offset position of the missile with respect to the projected tracker axis, and being incorporated in a flight control channel including means for transmitting flight control signals to a receiver in the missile for guiding the missile in flight; whereby a closed control loop is formed which includes an optical feedback path from the missile to the tracker; means for introducing the manual control signals into the closed control loop; a common support which supports the tracker and the sight, the tracker being so constructed and arranged that the direction of its projected axis is angularly movable relatively to the support; and a gyroscope mounted on the support and arranged to stabilise and control the direction of the projected tracker axis independently of movement of the support Error signals from the tracker may be used directly to provide the flight control signals for transmission to the missile whilst the manuallyoperable control is used to impose manual control on the gyroscope to cause it to steer the tracker axis. The manually-operable control may then be arranged to control the gyroscope to steer the tracker axis, thereby introducing the manual control signals into the closed control loop at the tracker stage rather than independently of the operation of the tracker.

Alternatively the error signals from the tracker may be used to precess the gyroscope so that it causes the tracker axis to follow the missile automatically whilst the error signals and the manual control signals together are used to form the flight control signals.

In another form of the invention the gyroscope is used to stabilise both the sight and the tracker, for use for example where the guidance system is mounted on a moving vehicle or craft The gyroscope may be arranged to provide a signal representing the direction and rate of angular displacement of the sight when tracking a moving target immediately prior to the launching of a missile, means being provided for introducing into the closed control loop, after launching of the missile, a prediction signal corresponding to the signal provided by the gyroscope immediately prior to launching for controlling the flight of the missile after launching.

In one such arrangement, the said means for introducing the prediction signal into the closed control loop comprises a rate compensation signal generator connected to receive the signal provided by the gyroscope immediately prior to launching. The rate compensation signal generator may also supply an additional signal for introduction into the closed control loop.

Another feature of the present invention comprises the integration of the optical systems for the tracker and sight into a single optical system which provides the field of view both for the tracker and for the operator viewing through the sight, and whose axis is the aforesaid tracker axis, the tracker and the stabilising gyroscope being connected in a quick-response autofollower control loop whereby the gyroscope steers the tracker axis to keep it aligned on the missile being tracked, the output of the tracker being also supplied to an integrator whose integrated output signal provides secondary control signals for controlling the missile during the guidance phase after gathering, during which phase the manual control signals from the manually-operated control are superimposed on the secondary control signals. The term "gathering" means the operation of bringing the missile into the guidance field of view after it has been launched. During the prelaunch and gathering phases the system is switched to a condition in which the tracker output is disconnected from the flying control signal transmitter and the output of the manually operated control is supplied directly to the torque motors of the gyroscope to drive the latter for the purpose of laying-on (aiming) the sight at a target This arrangement is intended for applications, such as a helicopter-borne weapons system, in which a gyro-stabilised optical sight is necessary for the observation of targets from a moving vehicle or crafL No prediction control is provided in this arrangement, The invention may be carried into practice in various ways, but certain specific embodiments will now be described by way of example only and with reference to the accompanying drawings, in which: Figure 1 is a general view of a fire control unit for controlling a missile, in which a stabilising gyroscope forms an integral part of an automatic missile tracker; Figure 2 is a schematic block diagram of one arrangement of the control system of the fire control unit of Figure 1, in which the error signal from the tracker is used directly to control the missile, and a joystick is used to alter the point direction of the tracker; Figure 3 is a diagram showing in longitudinal section the schematic arrangement of the tracker 2 and gyroscope 3 of the fire control unit; Figures 4,5 and 6 are diagrams similar to Figure 3 showing the tracker and gyroscope in different states of operation for the system of Figure 2; Figure 7 is a diagram similar to Figure 2 of a modified control system in which the error signal from the tracker is used to steer the tracker by means of the gyroscope to point at the missile, and control signals are generated from the tracker and the joystick to steer the missile; Figure 8 is a diagram of a third embodiment of the invention having a gyro-stabilised optical sight and a tracker whose optics are integrated with those of the sight, the single gyroscope stabilising both the tracker and the sight, the system being illustrated in the guidance phase; and Figure 9 is a diagram of the system of Figure 8 but illustrating the conditions in the prelaunch and missile-gathering phases.

Referring to Figure 1, a telescope sight 1 and an integrated tracker 2 and stabilising gyroscope 3 are mounted together on a housing 4 which contains all necessary electrical equipment The telescope 1 and tracker 2 are generally aligned, and the tracker 2 and gyroscope 3 are so integrated that the direction, relative to the housing, of the forwardly-projected optical axis of the tracker is variable and is determined exclusively by the position of the gyroscope flywheel, the direction of the optical axis being controlled by the action of an optical element attached to the gyroscope flywheel. A handgrip 5 is attached to the housing 4, and includes a joystick 6 to supply control signals. The handle 5 also carries a firing trigger 7 in order that an operator may fire a missile at the appropriate time. The whole equipment is designed in such a way that it can be used by an operator in a variety of ways, either supported on the ground by means of a support structure, or hand-held.

One purpose of the gyroscope 3 is to ensure that however the system is supported, the projected opocal axis of the tracka 2 is stabilised in space and does not respond to angular movements of the complete equipment Another purpose is to act as a device for steering the projected optical axis of the tracker by means of control signals applied to precessing torque motors contained within the gyroscope assembly, at the same time stabilising the projected optical axis of the tracker when the whole equipment is subjected to angular movement A third purpose of the gyroscope 3 is to act as a device for measuring angular velocities of the whole equipment at a time when stabilisation and steering action as described above are not required.

The tracker 2 is of the kind having a photo electric detector cell array with a diffuser screen on which a real image of a flare carried by a missile 8 is focussed by the tracker's optical system, the tracker producing electrical signals corresponding to the coordinates of the position of the missile image with respect to the electrical centre or datum point of the screen.

These signals are therefore representative of the direction and extent to which the missile is offset from the forwardly-projected optical axis of the tracker. These signals are suitably modified by the electrical equipment contained in the housing 4 and thereby produce control signals which are transmitted to the missile 8 via a trailing cable 9. These control signals are received by a receiver in the missile 8 and are employed to control the operation of steering actuators used to steer the missile. Olher types of tracker may also be used.

Referring now also to Figure 2, this shows diagrammatically one arrangement of the control system incorporated in the apparatus of Figure 1. In use, before launching the missile the operator first aims the sighting device 1 at the target by means of the handgrip 5, and tracks the target using an aiming mark in the sight, and thereby also directs the tracker 2 approximately at the target When it is required to launch a missile the operator depresses the firing trigger 7 whilst tracking the target with the sight The resultant firing signal is used to run-up and uncage the gyroscope 3 which, operating in the angular velocity measuring mode, passes a signal proportional to target angular velocity to a target rate compensation generator 10. After a shon delay the missile is then automatically launched, and an automatically-operated switch causes the gyroscope 3 to commence operating in the stabilising mode, the optical axis of the tracker 2 being thereby stabilised in the direction pointing to the target position at the instant of missile launching. After launch the tracker 2 detects the missile flare and generates signals representing the radial direction and distance of the missile position with respect to the stabilised tracker axis. These signals pass to a shaping unit 11 where they are modified and passed on to a transmitter 12 which in turn supplies command signals to the missile 8 via the trailing cable 9. Displacements 13 of the missile 8 in flight from the projected tracker axis are detected by the tracker 2 and compensating corrections are fed to the transmitter 12 by the error read-out system of the tracker so that the missile 8 is controlled by the tracker to stay on the axis of the tracker.

Also after launch, the target rate compensation generator passes a signal, representing the angular velocity of the sighting axis at the firing instant, to a shaping unit 14 which modifies the signal and passes it to the precessing torque motors in the gyroscope 3 to cause the optical axis of the tracker 3 to start and continue turning at an angular velocity corresponding to that of the sight which was established when tracing the target before launch. In addition the target rate compensation generator provides signals which pass to the shaping unit 11 for onward transmission to the missile. These signals correct a natural tendency of the missile to lag behind the tracker axis when the latter is moving.

During missile flight the operator observes through the sight 1 the positions of the target and the missile flare, and can maintain the missile night path along his line of sight to the target by use of the joystick 6, without necessarily turning the whole housing 4.

Movement of the joystick 6 provides an electrical signal which is modified by the shaping unit 14 and passed to the precessing torque motors in the stabilising gyroscope 3 to direct the optical axis of the tracker so that the missile flare and target are superimposed in the field of view of the sight 1. The joystick 6 thus provides primary directional control of the missile, superimposed on the automatic control by the tracker and the predictive control by the target rate compensation generator.

After the missile has been launched it may be necessary during target engagement for the operator to rotate the whole fire control unit so as to retain the target in the field of view of the sight 1. The rotation of the fire control unit after launching does not influence the flight direction of the missile because of the stabilising influence of the gyroscope 3.

Figure 3 shows the tracker 2 and the gyroscope 3 which share a housing 30. The gyroscope 3 is a twoaxis-fiee, light-weight, spring-start, re-usable instrument having a rotor 31 mounted in a gimbal assembly 32 with an associated torque motor and picksff assembly 33 whose peripheral torque motors can be utilised to position the inner gimbal of the assembly 32 with respect to orthogonal axes (azimulh and elevation). The torque motors of the assembly 33 have small-angle magnetic pick-offs integrated with them. A tensator spring assembly 34 with a wind-up lanyard 35 and a run-up and sequencing gear box 36 can be coupled to the rotor 31 by means of a caging clutch 37 controlled by a cage/uncage pin 38 protruding from the rear of the housing 30.

The tracker 2 in the forward end of the housing 30 includes an objective lens system 39 which contains in its centre the tracker detector assembly 40. The latter comprises a static split detector array41 mounted in front of a diffuser screen 42 and coupled to an amplifier 43. The objective lens system 39 has an optical axis 44.

A mirror 45 having a plane reflecting surface is mounted on the forward end of the gyro rotor 31 with the reflecting surface 46 at right angles to the axis of rotation of the rotor 31 and facing towards the diffuser screen 42. The reflecting surface 46 of the mirror 45 is set at a distance equal to half the back focal length of the objective lens assembly 39 from the diffuser screen 42, so that a real optical image of a distant object viewed by the tracker is formed by the lens assembly 39 on the diffuser screen 42 after reflection by the mirror 45.

The principles of operation of the optical system of the tracker will now be briefly described with reference to Figures 4 to 6.

Figure 4 shows the situation for basic tracker operation with the mirror 45 normal to the tracker axis 44, i.e., the gyro is either caged or undeflected. The lens assembly 39 and mirror 45 produce on the diffuser screen an image A' of a missile at A which is on the projected axis 44 of the tracker, the image A' being at the centre of the screen 42. A missile at B which is offset from the projected axis 44 of the tracker is imaged at B', offcentre with respect to the screen 42, and the edge of the screen 42 thus sets the limiting off-axis angle i; of the tracker field. The off-centre image B' on the screen 42 will generate a signal from the detector which is related to the off-centrc coordinates of the image B', and this signal is used as a demand signal by which the missile at B is controlled to fly back to position A on the projected tracker axis.

Figure 5 shows the situation when the tracker housing 30 has moved through an angle Bm as a result of a launch disturbance or a wander of aim. The gyro rotor 31 and mirror 45 have remained orientated in the original direction in space. The principal light ray from the missile at A still strikes normal to the mirror 45, and the image of the missile thus remains at A' at the centre of the detector screen 42. No error signal is produced by the detector, and so the missile is not demanded to change its position in space.

In other words the optical axis of the tracker has been stabilised in space by the gyroscope 3 and mirror 45 against the effect of disturbances of the tracker housing 30.

Figure 6 shows the situation when the mirror 45 is driven off-axis by means of a demand signal from the joystick 6 to the torque motors, as described with reference to Figure 2. A missile off-axis at B will be imaged at B' which is now at the centre of the screen 42, so that no signal will be produced by the detector 41. it follows that a missile at A which is responding to the output signal from the detector will be driven off-axis to B as the mirror moves.

It will be seen that the gyroscope 3 has three distinct modes of operation; (a) it is uncaged at the the start of a targettracking operation before launching, and performs the function of a rate gyro. Deflection of the gyrohracker housing 30 around the spacestabilised rotor 31 is measured by the small angle pick-offs, and the signal so generated is amplified and used to power the peripheral torque motors, which drive the rotor back into alignment with the housing 30. The rotor is thus effectively held on-axis with the flying control unit' but the signal on the pick-offs gives a measure of the torque required to hold the rotor in alignment with the housing and hence of the rate at which the tracker is being turned in following the moving target This signal is processed to provide target rate compensation as described with reference to Figure 2; (b) when the torque motors are decoupled from the pick-offs, as occurs at the instant when the missile is launched, the rotor is free to retain a space-fixed reference, and performs a normal stabilising function for the tracker axis; and (c) if a signal is applied to the torque motors from some source other than the pick-offs, the rotor will be deflected as a rate proportional to that signal, but will continue to perform a stabilising function against motion of the fire control uniL In practice the joystick 6 in the missile guidance system is used to drive the torque motors to provide positional changes in the rotor, as previously described.

An alternative arrangement of the control system and method of operation of the apparatus shown in Figure 1, is shown diagrammatically in Figure 7. In this arrangement the output of the tracker 2 is used to steer the gyroscope 3, and thereby the projected optical axis of the tracker 2, so as to align the tracker axis with the missile flare as observed through the sight In addition the tracker signals, which represent to the stabilising gyroscope 2 angular rate demands, are passed to an electrical integrator 15 which converts the angular rate signals to an integrated signal proportional to the angle turned through by the tracker axis. This integrated signal is passed to the shaping unit 11 and converted therein to command signals for onward transmission to the missile 8. The effect of this integrated signal is to control the missile 8 to move towards a sight line originally determined by the initial direction of the tracker axis. The axis of the tracker 2 remains pointed al the missile flare because of the flight control loop existing between the tracker 2 and the gyroscope 3, and as the tracker axis follows the missile flare and turns back to its original direction the signal from the integrator 15 returns to zero. in this arrangement the target rate compensation generator 10 operates as in the previous arrangement except that the signal delivered to the gyroscope 3 does not pass through the shaping unit 14. Signals from the joystick 6 pass to the shaping unit 14 where they are modified and passed direct to the shaping unit 11 as primary control signals for onward transmission to the missile 8. Joystick movement now controls the missile to move relative to the line of sight to the target, the axis of the tracker 2 being caused to follow the missile until the tracker output signal having passed through the integrator 15 is sufficient to null the joystick signal. Thus in this variant of the system the joystick 6 still provides primary directional control of the missile 8.

superimposed on the automatic control by the tracker 2 and the predictive control by the target rate compensation generator 10.

In both these arrangements the flight of the missile is not affected by any movement of the complete hand-held (or hand-guided) fire control unit after launching, owing to the space stabilising effect of the gyro.

Both arrangements embody predictive compensation for moving target rates. Prior to launch, the gyro is used to measure the approximate angular rate of the target as followed by the sight, and this rate is automatically maintained imposed on the tracker after launch. This feature is provided to assist in very short range engagements of crossing targets.

The third embodiment shown in Figures 8 and 9 relates to the application of the invention to a fire-control unit for a helicopter-borne weapon system. The main differences from the first two embodiments arise from the association of the system with a gyro-stabilised optical sight which is necessary for the observation of targets from the moving helicopter.

In this case the tracker optics are integrated with the optical sight and the gyro stabilises both sight and tracker. A single joystick is used both to steer the sight and, during engagements, to control the missile.

The general arrangement during the guidance phase is illustrated diagrammatically in Figure 8.

The tracker 2, gyro 3 and stabilising mirror 3A are connected in a fast auto-follow loop. This keeps the tracker axis (which is also the optical axis of the sight) aligned with the missile flare and the missile in the centre of the optical fieldof-view presented to the operator by the integrated optical system. Any tendency of the missile to drift causes the gyro to follow up and a signal to appear at the output of the integrator 15, this signal representing the measured drift of the missile. This causes commands to be fed to the missile so as the eliminate the drift Coursecorrecting movements of the missile and tracker sight can be effected by operating the joystick 6.

The arrangement before launch and during missile gathering is illustrated in Figure 9.

Before launch the same joystick 6 is connected to the torque motors of the gyro 3 and is used by the operator to lay the sight on the target using an illuminated aiming mark in the sighL Immediately after launch, switch S1 is closed and open loop commands may be sent through it to the missile to bring it towards the target sight-line which, in some missile systems, can be at large angles from the axis of the helicopter body on, which the missiles are mounted. When the missile enters the field of view of the trackerlsight which is initially operating in a wide-angle gather mode, switch S2 is closed and commands from the tracker 2 are sent to the missile to bring it onto the tracker axis. During the gathering phase the operator continues to lay the sight on the target using the illuminated aiming mark and the joystick. When the missile has been gathered, the tracker is changed to its narrow-field accurate-guidance mode, the illuminated marker is extinguished and the system is switched to the arrangement shown in Figure 8.

In this system, the target rate compensation feature described with reference to Figure 2 is not required since, in following the target with the illuminated marker, the operator establishes the required position of the joystick to give the correct sightline rate and, apart from small adjustments, the same joystick position will apply during the guidance phase.

WHAT WE CLAIM IS: 1. A guidance system for guided missiles, which comprises: an optical sight; a manuallyoperable control for generating manual control signals in response to manipulation thereof by an operator; a tracker having an optical system for tracking a missile in flight in the field of view of the tracker optical system, the tracker being constructed and arranged to generate error signals representative of the radially-offset position of the missile with respect to the projected tracker axis, and being incorporated in a flight control channel including means for transmitting flight control signals to a receiver in the missile for guiding the missile in flight, whereby a closed control loop is formed which includes an optical feedback path from the missile to the tracker; means for introducing the manual control signals into the closed control loop; a common support which supports the tracker and the sight, the tracker being so constructed and arranged that the direction of its projected axis is angularly movable relatively to the support; and a gyroscope mounted on the support and arranged to stabilise and control the direction of the projected tracker axis independently of movement of the support 2. A system as claimed in Claim 1 in which the optical systems of the sight and of the tracker are integrated as a single optical system which provides a common field of view for both the tracker and for the operator viewing Lhrough the sight, the axis of the integrated optical system comprising the coincident axes of the tracker and of the sight and in which the gyroscope is arranged to stabilise and control the direction of the projected axis of the integrated optical system.

3. A system as claimed in Claim 1 or Claim 2 in which the tracker error signals are used directly to provide the flight control signals.

4. A system as claimed in Claim 1 or Claim 2 in which the error signals of the tracker are utilised to cause the gyroscope to steer the axis of the tracker system in such a way as to maintain the said projected axis aligned on the missile being guided 5. A system as claimed in Claim 4 in which the tracker error signals are caused to precess the gyroscope to steer the tracker axis to follow the missile in flight, the tracker error signals being also utilised to provide the flight control signals.

6. A system as claimed in Claim 5 including means for integrating the tracker error signals and for utilising the integrated signals to form the flight control signals.

7. A system as claimed in Claim 6 in which the integrating means is so arranged that the integrated signals are representative of the deviation of the missile from the initial line of the projected tracker axis, whereby the effect of the flight control signals is to move the missile towards alignment with the said line.

8. A system as claimed in any of the preceding claims in which the gyroscope is arranged to provide a signal corresponding to the directional rate of angular displacement of the sight when tracking a moving target immediately prior to the launching of a missile, means being provided for introducing into the closed control loop, after launching of the missile, a prediction signal corresponding to the signal provided by the gyroscope immediately prior to launching, for controlling the flight of the missile after launching.

9. A system as claimed in Claim 8 in which the said means for introducing the prediction signal into the closed control loop is effective after launching of the missile to precess the gyroscope to cause it to steer the tracker axis in a direction and at an angular rate corresponding to those of the sight immediately before launching.

10. A system as claimed in Claim 8 or Claim 9 in which the said means for introducing the prediction signal into the closed control loop comprises a rate compensation signal generator connected to receive the signal provided by the gyroscope immediately prior to launching.

11. A system as claimed in Claim 10 when dependent on Claim 9, including means for introducing into the closed control loop an additional signal derived from the output of the rate compensation generator.

12. A system as claimed in any of Claims 8 to 11 in which the gyroscope is provided with torque motor means by which the axis of the

**WARNING** end of DESC field may overlap start of CLMS **.

Claims (17)

**WARNING** start of CLMS field may overlap end of DESC **. missile gathering is illustrated in Figure 9. Before launch the same joystick 6 is connected to the torque motors of the gyro 3 and is used by the operator to lay the sight on the target using an illuminated aiming mark in the sighL Immediately after launch, switch S1 is closed and open loop commands may be sent through it to the missile to bring it towards the target sight-line which, in some missile systems, can be at large angles from the axis of the helicopter body on, which the missiles are mounted. When the missile enters the field of view of the trackerlsight which is initially operating in a wide-angle gather mode, switch S2 is closed and commands from the tracker 2 are sent to the missile to bring it onto the tracker axis. During the gathering phase the operator continues to lay the sight on the target using the illuminated aiming mark and the joystick. When the missile has been gathered, the tracker is changed to its narrow-field accurate-guidance mode, the illuminated marker is extinguished and the system is switched to the arrangement shown in Figure 8. In this system, the target rate compensation feature described with reference to Figure 2 is not required since, in following the target with the illuminated marker, the operator establishes the required position of the joystick to give the correct sightline rate and, apart from small adjustments, the same joystick position will apply during the guidance phase. WHAT WE CLAIM IS:

1. A guidance system for guided missiles, which comprises: an optical sight; a manuallyoperable control for generating manual control signals in response to manipulation thereof by an operator; a tracker having an optical system for tracking a missile in flight in the field of view of the tracker optical system, the tracker being constructed and arranged to generate error signals representative of the radially-offset position of the missile with respect to the projected tracker axis, and being incorporated in a flight control channel including means for transmitting flight control signals to a receiver in the missile for guiding the missile in flight, whereby a closed control loop is formed which includes an optical feedback path from the missile to the tracker; means for introducing the manual control signals into the closed control loop; a common support which supports the tracker and the sight, the tracker being so constructed and arranged that the direction of its projected axis is angularly movable relatively to the support; and a gyroscope mounted on the support and arranged to stabilise and control the direction of the projected tracker axis independently of movement of the support

2. A system as claimed in Claim 1 in which the optical systems of the sight and of the tracker are integrated as a single optical system which provides a common field of view for both the tracker and for the operator viewing Lhrough the sight, the axis of the integrated optical system comprising the coincident axes of the tracker and of the sight and in which the gyroscope is arranged to stabilise and control the direction of the projected axis of the integrated optical system.

3. A system as claimed in Claim 1 or Claim 2 in which the tracker error signals are used directly to provide the flight control signals.

4. A system as claimed in Claim 1 or Claim 2 in which the error signals of the tracker are utilised to cause the gyroscope to steer the axis of the tracker system in such a way as to maintain the said projected axis aligned on the missile being guided

5. A system as claimed in Claim 4 in which the tracker error signals are caused to precess the gyroscope to steer the tracker axis to follow the missile in flight, the tracker error signals being also utilised to provide the flight control signals.

6. A system as claimed in Claim 5 including means for integrating the tracker error signals and for utilising the integrated signals to form the flight control signals.

7. A system as claimed in Claim 6 in which the integrating means is so arranged that the integrated signals are representative of the deviation of the missile from the initial line of the projected tracker axis, whereby the effect of the flight control signals is to move the missile towards alignment with the said line.

8. A system as claimed in any of the preceding claims in which the gyroscope is arranged to provide a signal corresponding to the directional rate of angular displacement of the sight when tracking a moving target immediately prior to the launching of a missile, means being provided for introducing into the closed control loop, after launching of the missile, a prediction signal corresponding to the signal provided by the gyroscope immediately prior to launching, for controlling the flight of the missile after launching.

9. A system as claimed in Claim 8 in which the said means for introducing the prediction signal into the closed control loop is effective after launching of the missile to precess the gyroscope to cause it to steer the tracker axis in a direction and at an angular rate corresponding to those of the sight immediately before launching.

10. A system as claimed in Claim 8 or Claim 9 in which the said means for introducing the prediction signal into the closed control loop comprises a rate compensation signal generator connected to receive the signal provided by the gyroscope immediately prior to launching.

11. A system as claimed in Claim 10 when dependent on Claim 9, including means for introducing into the closed control loop an additional signal derived from the output of the rate compensation generator.

12. A system as claimed in any of Claims 8 to 11 in which the gyroscope is provided with torque motor means by which the axis of the

gyroscope rotor can be steered in response to signals supplied to the torque motor means, the torque motor means being provided with electromagnetic pick-off means, for measuring deflection or rate of deflection of the rotor axis.

13. A system as claimed in any of the preceding claims in which the means for introducing the manual control signals into the closed control loop is arranged to utilise the manual control signals to precess the gyroscope to steer the axis of th tracker.

14. A system as claimed in any one of the preceding claims in which the tracker and the gyroscope are mounted in a common housing, the gyroscope rotor being mounted by means of a gimbal arrangement behind the tracker optical system in the housing, the tracker having a rearwardly-facing detection screen positioned behind its objective lens assembly, and in which the gyroscope carries a mirror so mounted that its reflecting surface is maintained radial with respect to the rotor axis and faces towards the tracker detector screen, the mirror being arranged to reflect back onto the tracker detector screen a real optical image formed by the tracker optical system of the scene in the field of view in front of the tracker optical system.

15. A system as claimed in Claim 14 in which the distance of the reflecting surface of the mirror from the tracker screen equals half the back focal length of the objective lens assembly of the tracker optical system.

16. A system as claimed in any of the preceding claims in which the common support for the sight and the tracker comprises a movable vehicle or craft

17. A guidance system for guided missiles, the system being substantially as herein described with reference to Figures 1 to 6, or Figures 1 and 3 to 7, or Figures 8 and 9, of the accompanying drawings.