JPS5928840B2 - Aircraft guidance system - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to an aircraft guidance system that guides an aircraft in a command guidance mode during part of its flight and in a homing guidance mode during other parts of its flight.

More particularly, the present invention relates to an air vehicle guidance system that requires the provision of a reference element within the air vehicle having a known attitude orientation.

As is known in the art, an air vehicle, such as a missile, is directed toward a target by guidance signals generated from tracking data obtained at a remote radar station or by a radar located on the missile. be guided.

The former is generally called a command guidance system, and the latter is called a homing guidance system.

For example, in a command and guidance system where a missile is used to attack a flying target, a large, high-resolution radar device and a high-speed digital computer at a remote location select one target from among multiple targets, and the missile The system tracks this target and uses the tracking data to calculate and send appropriate guidance signals to the missile.

As is known, it is generally necessary to provide the missile with a reference element, such as an attitude-stabilized platform with an angular orientation known at a remote base, to convert the transmitted guidance signals into missile control signals. .

In homing guidance systems, a small, lightweight, low power tracking radar system is provided to generate target tracking data and guidance signals.

Such low power tracking radar systems or semi-
As in the case of active radar, at least its receiver part) is relatively lightweight and can therefore be placed inside the missile.

Such homing guidance systems typically include a target tracking antenna.

The target tracking antenna is gimballed to substantially eliminate the effects of missile vehicle angular velocity on the tracking data.

Some conventional missile systems combine command and homing guidance features.

During the early part of the missile's flight, guidance signals are generated by a digital computer operated in response to signals obtained by tracking the missile and specific targets with a high-resolution radar system, and during the later part of the flight, the missile The guidance signal is obtained by tracking the target with a receiver section of the radar system that is fed the signal by a tracking element gimbaled within the radar system.

For the reasons discussed above, such a missile will require an attitude reference element during at least the early part of its flight.

One way to provide a baseline for this postural stabilization is to use a postural stabilization platform.

This additionally requires a gimbaled radar receiving antenna with two degrees of freedom on the missile vehicle to obtain tracking data during the latter portion of the flight.

Attitude stabilization platforms generally have (a) angular velocities about each of three mutually orthogonal axes;
(b) three angular rate sensing gyros arranged to measure (b) the attitude stabilization platform relative to the missile vehicle in a manner that compensates for any angular rotation; It includes three drive means corresponding to the three gyros.

A common mechanism providing this drive means is a mechanical servo.

However, these servos are relatively expensive and often relatively heavy.

It is an object of the present invention to provide an air vehicle guidance system in which an air vehicle is selectively responsive to guidance signals obtained from tracking data generated at a remote base and to tracking data generated by means contained within the air vehicle. It is an object of the present invention to provide an improved apparatus for use in a system contained within a vehicle that provides both a reference element having a known angular orientation at a remote base and a gimballed tracking element. .

Another object of the invention is to control the reference element without the use of three relatively expensive and heavy mechanical servos.

These and other objects of the present invention are generally achieved by having a gimballed tracking element included within the vehicle also function as an attitude reference element.

This tracking element is used for pitching and yaw (y
Awing is stabilized by a mechanical servo drive and rolling is stabilized by aerodynamic means.

In a preferred embodiment, the target tracking antenna gimballed on the missile with two degrees of freedom relative to the airframe of the vehicle includes a third angular rate sensing gyro in addition to the normal pair of angular rate sensing gyros. is attached.

The three gyros are arranged to detect any angular change in the inertial direction of the antenna.

The output signal from each of a conventional pair of angular velocity sensing servos is used as an input signal to one of the mechanical servo drives coupled to the antenna.

The output signal from the third angular rate sensing gyro is used as an autopilot input signal for missile rolling.

These two servo drives and the rolling autopilot thus function to stabilize the attitude of the antenna, which stabilizes the antenna during at least a portion of the missile's flight. It acts as a reference element, allowing a correct conversion of the guidance signals transmitted to and from the base, and as a target tracking antenna.

The two servo drives and rolling autopilot are also adapted to respond to signals sent to the missile from a remote base to orient the reference element to any desired attitude.

The present invention will be explained in more detail below with reference to the drawings.

FIG. 1 shows that the missile 10 receives either a guidance signal transmitted to the missile 10 from a remote radar station 13 (during a command guidance system) or a guidance signal generated by a radar contained within the missile 10 (during a homing guidance system). This shows a state in which the aircraft is guided to attack a target 12 in response to a signal.

When operating in command and guidance mode, remote radar system 13 tracks both missile 10 and target 120.

Tracking data obtained from this system is converted into guidance signals by a digital computer C (located at the remote V-der base, but not shown).

This guidance signal is transmitted to the missile 10, and the missile 10
The signals are received by an antenna 15 supported below and sent to a receiver and processing unit 17 where these signals are converted into control signals for the flight control section 19 of the missile.

The guidance signal at the remote radar base 13 is the missile 10
It should be clear that this is a signal requesting the driver to steer in the desired direction.

This guidance signal is converted by the flight controller 19 to control the missile 10.

In particular, when the missile 10 has a cruciform shape as in this example, the pitching control vane and the yawing control vane 20 may be individually controlled along two orthogonal axes (commonly referred to as the pitching axis and the yawing axis). It can be steered in any lateral direction.

These guidance signals, which request maneuvering of the missile in a desired direction, must be properly resolved between the pitching and yaw control wings 20 to accomplish the requested maneuver.

Resolving this signal is possible if the angular orientation of the wing 20 is known.

The angular direction of the wing 200 can be known because the missile 10 is equipped with a reference element, which will be described later.

Suffice it to state here that the angular orientation (ie attitude) of the reference element is known at the remote radar station 13.

When operating in the homing guidance mode, the missile 10 is oriented such that its nose at the beginning of the mode is approximately on a collision course with the target 12.

Remote radar station 13 targets radar frequency waves 12
A portion of the wave is reflected by the target 12 and received by the missile's tracking antenna 22.

A signal representative of the angular deviation of the target 120 from the central axis of the target tracking antenna 22 is generated in a conventional manner and transmitted to the target tracking antenna controller 2 via a receiver and processor 17.
4 and the flight control unit 190.

This signal drives tracking antenna 22 to maintain tracking of target 12 and steers missile 10 to maintain a collision course with target 12.

In other words, the missile 10 is on a collision course with the target 12 and the center axis of the target tracking antenna 22 is aligned with the target 12.
2, the antenna 22 will be pointing toward the target 12 as long as the missile 10 is on this course.

However, if missile 10 deviates from this collision course (eg, if target 12 moves), the central axis of target tracking antenna 22 is no longer oriented toward target 12.

That is, an axis deviation is detected at this time, which is proportional to the amount of change in the line-of-sight angle between the missile 10 and the target 12.

It will be clear that the target tracking antenna 22 needs to be driven to zero this axis offset so as not to lose sight of the target 12.

Furthermore, as is known in proportional navigation guidance, it is desirable to turn the missile 10 from its current course at a speed proportional to the rate of this line-of-sight angle change.

This missile rotation speed is obtained by accelerating the missile laterally in a direction that makes the line-of-sight angle change zero.

That is, missile 10 must be maneuvered back onto a collision course with target 12.

The lateral acceleration of this missile 10 is caused by the selected control wing 2.
This is possible by deflecting 0.

In particular, the tracking of the target 12 by the tracking antenna 22 and the control to maintain the missile 10 on a collision course generate a signal representing the deviation of the antenna axis and apply this signal to the flight control section 19 and the target tracking antenna control section 24. This is achieved by

It is noted here that the target tracking antenna 22 is gimballed with two degrees of freedom relative to the missile vehicle.

This gimbal support is conventionally accomplished by attaching two velocity sensing gyros to the antenna 22 that sense the inertial angular velocity of the antenna and providing drive means to move the antenna 22 relative to the missile 100 vehicle.

The target tracking antenna 22 is gimballed to prevent missile body motion from generating false off-axis indication signals.

FIG. 2 shows a flight control section 19 including pitching, yaw and rolling autopilots 26, 27 and 28, each controlled by a receiver and processor 17.
shows.

Output signals generated from each of the autopilots are applied to the actuator section 30.

pitching-yaw and rolling autopilots 26 , 27 .

28 and actuator portion 30 may be of conventional technology.

The actuator section is mechanically connected to the control vane 20.

The control wing 20 is rotatably supported relative to the missile body.

An opposing pair of trapezoids (pitching control vanes, i.e. those driven by the pitching autopilot) cause maneuvering of the missile about the pitching axis P, and the other pair of control vanes (yawing control vanes, i.e. The yaw autopilot (driven by the yaw autopilot) causes the missile to maneuver around the yaw axis Y.

Differential deflections of a pair of control vanes, such as a pitching control vane, cause the missile 10 to roll about its longitudinal axis.

Deflection with such a difference is caused by the actuator section 30
occurs in response to a command signal from rolling pilot 28.

Target tracking antenna control section 24 has a structure 32 to which target tracking antenna 22 is mounted.

Structure 32 is designed to allow target tracking antenna 22 to move within the missile with two degrees of freedom.

In particular, the structure 32 has a base 34 suitably attached to the missile vehicle.

The outer member 36 is rotatably supported by the base 34 about an antenna pitching axis P' that is parallel to the pitching axis P, and the inner member 38 is rotatable about an antenna yawing axis Y' that is perpendicular to the antenna pitching axis P'. Possibly outer member 36
is supported by

Target tracking antenna 22 is attached to inner member 38 .

The central axis of this antenna is perpendicular to the antenna pitching axis P' and the antenna yawing axis Y'.

Target tracking antenna 22 may be of conventional technology, but here it is a monopulse antenna.

Therefore, of the pair of signals generated by this antenna, one represents a component of axis deviation with respect to antenna pitching axis P', and the other represents a component of axis deviation with respect to antenna yawing axis Y'.

Outer member 36, inner member 38 are coupled in any conventional manner (shown here in dotted lines) to pitch drive 42 and yaw drive 40, respectively.

That is, the yawing drive device 40 rotates the inner member 38 about the antenna yawing axis Y', and similarly the pitching drive device 42 rotates the outer member 36 about the antenna pitching axis P'.

The yawing drive device 40 and the pitching drive device 42
is an electrical or mechanical prime mover responsive to electrical signals provided by the receiver and processor 17.

Three antenna angular velocity detection gyros 44, 46, 4
8 is supported by the inner member 38.

The input axes of these three antenna angular velocity detection gyros are arranged to be orthogonal to each other.

In detail, the antenna angular velocity detection gyro 44, 46.4
8 indicates that each gyro's output signal is a target tracking ano2.
The antenna is arranged to express angular velocity around the antenna yawing axis Y', the antenna pitching axis p/, and the antenna center axis.

It is therefore clear that these output signals are adapted to provide a measurement of the inertial angular velocity of the structure 32.

FIG. 3 shows a receiver and processor 17 having a hetrodyne receiver 50 fed by a target tracking antenna 22 and a hetrodyne receiver and controller 52 fed by a downwardly supported antenna 15. shows.

The hetrodyne receiver and control device 52 includes the missile 1
0 generates a gate signal on either line H or line C according to the required induction mode.

That is, a gate signal is applied to the line H during the homing guidance mode, and to the line C during the command guidance mode.

Suppose now that missile 10 is launched in command-guided mode.

It is first noted that the signal from the hetrodyne receiver 50 (ie, from the target tracking antenna 22) is not communicated to the flight restriction section 19 or the target tracking antenna control section 24.

Furthermore, it is noted that the structure 32 is stabilized in attitude and is therefore considered a reference element.

That is, in the command guidance mode, the inertial angular velocity received by the body structure 32 is determined by the antenna angular velocity detection gyro 44, 46.4.
8, and signals representative of these inertial angular velocities are communicated to means (described below) for returning the structure 32 to the angular orientation it had before experiencing such inertial angular velocities.

In particular, these signals representing the inertial angular velocities about the antenna yaw axis Y' and the antenna pitching axis P' sensed by the antenna angular velocity sensing gyros 44, 46 are transmitted to the yawing drive device 40 via an integrator 54, respectively. and is supplied to the pitching drive device 42.

The rolling direction of the body structure 32 is determined by using the output signal of the antenna angular velocity detection gyro 48 as the rolling autopilot 28.
Posture is stabilized by combining with.

The rolling autopilot 28 is used to control the rolling direction of the missile 10 because the aerodynamic response to the difference in deflection of the pitching control wing of the missile 10 is fast enough to inertially stabilize the structure 32 in the rolling direction. Inertial stabilization is possible.

Therefore, since the missile 10 is provided with a reference element whose angular orientation is known at the remote radar station (FIG. 1), the missile
It is possible to guide with the guidance signal from 3.

In particular, guidance signals transmitted from the remote radar station 13 are received by the downwardly supported antenna 15 and processed by the hedrodyne receiver and controller 52.

This guidance signal includes the following information. (1) the guidance mode in which the missile should operate (command or homing) determined by the distance between the missile 10 and the target 12; (2) the maneuver relative to the known angular orientation of the reference element (i.e., the structure 32); and (3) a command signal for orienting the structure 32 in a known direction to bring the structure 32 within the limits permitted by the physical space available to it within the missile 10.

The induction mode selection information is used to determine whether the gate signal should be applied to line C or line H.

The maneuvering signals are converted to electrical signals on line 56, which are fed to computer 58, where these signals are converted into acceleration commands associated with control wings 20 (ie, the heading of the missile's body).

This conversion of maneuver signals from "reference element" coordinates to "missile vehicle" coordinates is accomplished by arranging within the missile 10 a prior art angular transducer 59,60.

Such transducers are suitably mounted to provide measurements of the orientation of the inner element 38 and outer element 36 relative to the missile vehicle.

The output of this transducer 59,600 is provided to a computer 58 along with a steering signal on line 56.

Correctly resolved signals are provided through gates 62 and 64 to pitching autopilot 26 and yaw autopilot 27.

A command signal to deflect the structure 32 in a known direction is processed by a hetrodyne receiver and controller 52 and sent to the line 6.
6. Appears as an angle command signal at 68 and 70.

This signal is provided to a deviation computer 72.

Taking the line signal as an example, this signal represents the desired "yaw" position of the body 32 in inertial space.

Since the output of the integrator 54 is a signal representing the actual payoing position of the structure 32 in inertial space, if
If the desired position and the actual position are not equal, a deviation signal is generated on line 74.

This deviation signal is supplied to the yawing drive device 40 through a gate 75 and an adder 77.

Instructions for the pitching position of the body structure 32 are processed in the same manner.

However, regarding rolling, the deviation computer 72
The .degree.' deviation signal is provided to the rolling autopilot 28 through the upper gate 76.

The rolling autopilot 28 generates signals in response to this °' deviation signal, which cause the pitching control blades to be deflected by different amounts.

The rolling attitude of the missile changes due to the difference in pitching control wing deflection, and this difference in deflection continues until the "deviation signal" becomes zero.

Now assume that the distance between the missile 10 and the target 12, as determined by the remote tracking station, has decreased to a point where it is desired to guide the missile in homing guidance mode.

The missile is believed to be almost on a collision course with Target 12.

A signal is transmitted from the remote base to missile 10 which causes heterodyne receiver and controller 52 to generate a gating signal to line H and remove the gating signal from line C.

Therefore, deviation computer 72 and computer 58
The signals are not supplied to the yawing drive 40, the pitching drive 42, the rolling autopilot 28, the pitching autopilot 26, and the yawing autopilot 27.

The yaw axis misalignment and pitch axis misalignment signals generated by the target tracking antenna 22 are heterodyned in a heterogain receiver 50 to generate a video frequency signal.

These signals are fed to lines 78, 80 and to pitching autopilot 26 and yawing autopilot 27 through gates 82 and 84, respectively.

The signal on line 80 is also fed to yawing drive 40 via gate 86 and adder 77, and the signal on line 78 is fed to pitching drive 42 via gate 88 and adder 89 for target tracking. antenna 22 is target 12
can continue to be tracked.

The signal from antenna yaw gyro 44 is provided through gate 90 to yaw drive 40 with the signal through gate 86 .

Similarly, antenna pitching--gyro 46
The signal from the oscilloscope is provided through gate 92 to pitching drive 40 with the signal through gate 88 .

The signals from the antenna pitching gyro 44 and the antenna yaw gyro 46 are used to detect the inertial angular velocity of the target tracking antenna 22, and the relative position of the antenna with respect to the missile vehicle for the reasons described above. .

FIG. 4 shows another specific embodiment of the invention.

The missile 10 is shown here in command guidance mode.

However, the tracking signal is generated from the signal received by the target tracking antenna 22 rather than from the signal received from the target at the remote radar station 13.

In this manner, signals received at target tracking antenna 22 are sent to heterodyne receiver 50 and transmitted to transmitter 100.
is sent back to the remote radar station 13 by.

The signal from transmitter 100 is transmitted to remote tracking station 1-3 using circulator 102 and downwardly mounted antenna 15.
sent to.

The returned signals are processed by a computer at the remote radar station 13 to determine a guidance signal for the missile.

This guidance signal is transmitted from a remote radar station 13 and received by a downwardly supported antenna 15.

Signals received at the down-supported antenna 15 are processed and responded to by the missile as detailed in the example of FIG.

It is known that target tracking antenna 22 (and structure 32) serves as a target tracking element as well as a reference element as required during command guidance mode.

Directional control and inertial stabilization of this reference element is achieved by means similar to those detailed in the command guidance mode in FIG.

Although the present invention has been described using a velocity sensing gyro attached to target tracking antenna 22, it will be apparent to those skilled in the art that this gyro and its associated integrator may be replaced by an angular velocity integrating gyro.

[Brief explanation of drawings]

FIG. 1 shows a missile guided towards a selected target by a guidance system according to the invention, and FIG. 2 shows the main parts of the guidance and control elements mounted on the missile and their relationship to the target tracking antenna. A block diagram, FIG. 3 shows a detailed view of the receiver and processing device of FIG. 2, and FIG. 4 shows an alternative embodiment of the invention. 10: Missile, 122 aircraft, 13: Radar base,
15: Antenna, 17: Receiver and processing unit, 19: Missile flight control unit, 22: Target tracking antenna, 24: Target antenna control unit, 26, 27° 28: Autopilot, 30: Actuator unit, 32 Two-standard construction element structure, 40: yawing drive device, 42: pitching drive device, 44° 46, 48: speed sensing gyro, 50:
Hetrodyne receiver, 52: control device.

Claims (1)

[Scope of Claims] 1. A flying object that guides the flying object toward a selected target in a command guidance mode during part of the flight and in a homing guidance mode during the other period of the flight. The guidance device includes: a target tracking antenna and a structure supporting the antenna; first, second, and third antennas attached to the structure and detecting inertial angular velocities with respect to the pitching axis, yawing axis, and rolling axis of the target tracking antenna, respectively. an antenna angular velocity detection gyro; a drive device for rotating the structural pitching axis and the yawing axis; a flight control device for controlling the flight direction of the aircraft; and a processing device for specifying a guidance mode in response to a guidance signal. In the command guidance mode, a control signal is supplied to the flight control device according to the guidance signal, and the drive device is controlled in response to the first, second, and third antenna angular velocity detection gyros; stabilizing the attitude of the target tracking antenna in a known angular direction, and supplying a maneuver signal to the flight control device in accordance with the angular velocities detected by the first and second antenna angular velocity detection gyros in the homing guidance mode; a processing device that controls the drive device in response to the first and second antenna angular velocity detection gyros, and causes the target tracking antenna to be gimballed with respect to the aircraft about a pitching axis and a yaw axis; Aircraft guidance system.