Patent GB-1077279-A (GB)
Title: Improvements in or relating to navigational systems
Inventor: WEEDON MICHAEL GEORGE BEESTON
IPC: F41G7/22
Filed: 1963-12-05
Abstract: 1,077,279. Radio direction finders. MINISTER OF AVIATION. Nov. 23, 1964 [Dec. 5, 1963], No. 48177/63. Heading H4D. Navigational apparatus for guiding an aircraft or missile towards a source of radiation includes two directional receivers having main lobes of their polar sensitivity diagrams overlapping and having their directions of maximum sensitivity inclined to one another, two receiving channels each arranged to receive signals from a separate one of the directional receivers and controllable to act upon the said signals with an adjustable relative gain which determines the orientation of an equisignal plane which is the locus of those possible positions of the source which would result in equal outputs from the two receiving channels, differencing means for generating a target bearing signal dependent on the difference in amplitude of the outputs of the receiving channels and gain control means responsive to the target bearing signal for adjusting the relative gain acting on the signals in the receiving channels so that the equisignal plane is moved towards the actual position of the source of radiation. The apparatus is particularly applicable to a known type of missile guidance system in which the rate of change of the angle between the missileto-target sight line and a space reference axis is measured, a desired rate of turn of the missile flight path is computed therefrom, and controls are automatically operated to achieve the desired rate of turn of the missile flight path. As shown, a missile guidance system arranged to guide yawing components of the missile motion comprises two receiver aerials 1, 2 respectively connected to independent variable gain amplifiers 3, 4, the outputs of which are applied to detectors 5, 6. The detector outputs are applied to a subtracting circuit 7 and to a summing circuit 8, the outputs of 7, 8 being applied to a dividing circuit 9. The output of 9 is passed through a filter 10 to a high gain amplifier 11 and thence to one input of an adding circuit 12. The output of a turn rate signal generator 13 is passed through a filter 14 and integrator 15 to another input of the adding circuit. The output of the adding circuit 12 is applied to a phase splitter 16 whose outputs control, respectively, the gains of the amplifiers 3, 4. The output of 12 is also applied through a filter 17 to a differentiating circuit 18 and thence to one input of an adding circuit 19. The output of 13 is also passed through a filter 20 to a second input of 19, the output of 19 being applied to an autopilot 21. In Fig. 3, OM is the missile axis, OT is the projection on the yaw plane of the missile-to-target sight line, OR is the projection on the yaw plane of a fixed space reference axis, and OE is the intersection of the yaw plane and the equisignal plane. Fig. 3 also shows the missile axis displaced by a sudden yaw from its original direction OM to a direction OM<SP>1</SP>. The system operates as follows. The aerials 1, 2 are rigidly mounted on the missile and together with parts 3. . . 11 and 16 form an automatic direction finder. Radio signals from a distant transmitter (not shown) received by the aerials 1, 2 produce a difference signal D in the circuit 7 and a sum signal S in the circuit 8. The circuit 9 produces an output proportional to D/S, this output being a function of angles # and # in Fig. 3. The D signal - is applied to the phase splitter 16 S which produces two equal and opposite outputs so that the gain of one of the amplifiers 3, 4 is increased and the gain of the other reduced, thereby bringing the difference between the outputs of the amplifiers to zero. Thus the equisignal plane OE is moved to be almost coincident with the target sight line OT, the angle α being exaggerated in Fig. 3. The direction of OE and its rate of turn can be used as a measure of the direction and rate of turn of OT. Components 12 . . . 15 are provided to minimize the effect of sudden yawing oscillations, these components and phase splitter 16 having a fast response. The turn rate signal generator 13, which may be a rate gyroscope produces a signal proportional to the rate of d# turn -of the missile axis OM relative to the dt fixed space reference axis OR. This signal is integrated and inverted by the integrator 15 to produce a signal which is applied to the phase splitter 16 through circuit 12 so that when the missile yaws suddenly through an angle # the equisignal plane is moved through an angle - #. As shown in Fig. 3, the angles # and # assume the values B<SP>1</SP> and #<SP>1</SP>. The outputs of circuit 12 and integrator 15 then represent #<SP>1</SP> and #<SP>1</SP>. The signal #<SP>1</SP> is differentiated by circuit 18 and then added to the turn rate signal d#<SP>1</SP>/dt by the circuit 19 to produce a signal representing the rate of change d#/dt of the angle # between OE and OR. The signal d#/dt is applied to the autopilot 21 which controls the missile flight path so that as far as possible the rate of turn of the missile flight path in the yaw plane is proportional to the rate of turn d#/dt. A duplicate system may be provided to guide the pitching components of the missile motion.