GB1601829A - Vehicle guidance apparatus - Google Patents

Description

(54) VEHICLE GUIDANCE APPARATUS (71) We, RAYTHEON COMPANY, a corporation organized under the laws of the State of Delaware, United States of America, of Lexington, Massachusetts, United States of America, 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 generally to vehicle guidance apparatus wherein a vehicle is guided for a portion of its flight in a command guidance mode and for a further portion of its flight in a homing guidance mode. More particularly, the invention pertains to such vehicle guidance apparatus wherein a reference element having a known attitude orientation is required to be contained within such vehicle.

As is known in the art, a vehicle, such as a missile, may be guided towards a target by guidance signals developed from tracking data obtained either at a remote radar station or by radar means contained within the missile. The former system is commonly called a command guidance system and the latter a homing guidance system. For example, in a command guidance missile system wherein a missile is used to intercept an airborne target, a large, remotely located high resolution radar system and high speed digital computer may be provided for selecting one of a plurality of targets, tracking both the missile and the selected target, calculating proper guidance signals for the missile from generated tracking data, and transmitting such calculated guidance signals to the missile. As is known, a reference element such as an attitude stabilized platform, having an angular orientation which is known at the remote station, is generally required to be contained within the missile for enabling transformation of the transmitted guidance signals into missile control signals.

Further. in a homing guidance missile system a smaller, light weight, low power tracking radar system may be provided for generation of both target tracking data and guidance signals. Such low power tracking radar system (or at least the relative portion thereof as in a semi-active application) may, because of its relative lighter weight, be contained within the missile. Generally such homing guidance system includes a target tracking antenna. The target tracking antenna is generally gimballed to substantially eliminate the effect of missile body rate on the tracking data.

In one type of known missile systems, the features of command guidance and homing guidance techniques are combined. During the early portion of the missile's flight, guidance signals are developed by a digital computer operated in response to signals obtained by tracking both the missile and a selected target with a high resolution radar system. During the latter portion of the flight guidance signals are obtained by tracking the target with the radar receiver portion of a radar system fed by a gimballed tracking element carried by the missile.

For reasons discussed above, such a missile would require an attitude reference element during at least the early portion of the missile's flight. One arrangement considered for providing such an attitude stabilized reference is to use an attitude stabilized platform. Such an arrangement requires, additionally, a gimballed radar receiving antenna having at least two degrees of freedom with respect to the missile body for providing tracking data during the latter portion of the missile's flight. The attitude stabilized platform generally includes three rate sensing gyros disposed to measure angular rates about respective ones of three mutually orthogonal axes and three corresponding drive means controlled, respectively, by each one of the rate sensing gyros to rotate, relative to the missile's body, the platform in a manner so as to compensate for any angular rotation experienced thereby. A common mechanism for providing such drive means is a mechanical servo. Such servos are relatively costly and are generally relatively heavy.

With a view to avoiding this disadvantage, the present invention provides a vehicle guidance apparatus, whereby a vehicle is guided during a portion of flight towards a target in a command guidance mode and during another portion of the flight in a homing guidance mode the apparatus comprising target tracking means, including a target tracking element carried by the vehicle. for generating guidance signals for the vehicle during the homing guidance mode; first means, fixed relative to the target tracking element, for sensing the inertial angular rates of this element about a pitch axis and a yaw axis; means, responsive to the first means, for gimballing the target tracking element with respect to the vehicle body about the pitch axis and the yaw axis during the homing guidance mode; second means, fixed relative to the target tracking element, for sensing the inertial angular rate of the target tracking element about an axis orthogonal to both the pitch axis and the yaw axis; and means, responsive to the first means and the second means, for stabilizing the attitude of the target tracking element at a known angular orientation during the command guidance mode.

In the preferred embodiment of the invention, the gimballed tracking element contained within a vehicle functions additionally as an attitude reference element which is stabilized by mechanical servo drives in pitch and yaw and stabilized by aerodynamical means in roll.

A gimballed target tracking antenna on a missile having two degrees of freedom with respect to the vehicle body has mounted thereon, in addition to its conventional pair of rate sensing gyros, a third rate sensing gyro. The three gyros are disposed to sense any angular change in the inertial orientation of the antenna. The output signals from the conventional pair of rate sensing gyros are used as input signals for the mechanical servo drives coupled to the antenna. The output signals from the third rate sensing gyro are used as input signals to the missile's role autopilot. The two servo drives and the roll autopilot thus function to attitude stabilize the antenna, the attitude stabilized antenna thus being adapted to serve as a reference element for enabling proper transformation of guidance signals transmitted between the missile and a remote station during a portion of the missile's fight and as a target tracking antenna. The two servo drives and the roll autopilot are also adapted to respond to signals transmitted to the missile from the remote station to position the reference element in any desired attitude orientation.

The invention will be described in more detail, by way of example, with reference to the accompanying drawings, in which: FIG. 1 shows a missile being guided towards a selected target by guidance apparatus embodying the invention.

FIG. 2 shows a block diagram of the main portions of guidance control elements carried by the missile and their relationship to a target tracking antenna; FIG. 3 shows in some detail the receiver/ processor of FIG. 2; and FIG. 4 shows an alternative embodiment of the invention.

Referring now to FIG. 1, a missile 10 is shown in flight. being guided to intercept a target 12 by responding either to guidance signals transmitted to the missile 10 from a remote radar station 13 (in a command guidance mode) or to guidance signals developed by a radar contained within the missile 10 (in a homing guidance mode).

When operating in the command guidance mode, the remote radar system 13 tracks both the missile 10 and the target 12. The tracking data is processed in a digital computer (not shown but located at the remote radar station) to convert such tracking data into guidance signals. The guidance signals are then transmitted to the missile 10. The transmitted guidance signals are received by a downlink antenna 15 and are then fed to receiver/processor 17 where they are converted into control signals for the missile's flight control section 19. It is readily apparent that the guidance signals computed at the remote radar station 13 are signals calling for missile 10 to manoeuvre in desired directions. The guidance signals are converted by the flight control section 19 to manoeuvre the missile 10. In particular, missile 10 being cruciform in tail cross-section may be manoeuvred in any lateral direction by controlling pitch and yaw control surfaces 20 about two orthogonal axes (commonly called the pitch and yaw axes) respectively. It follows that the guidance signals calling for a manoeuvre of the missile in the desired direction must be resolved properly between the pitch and yaw control surfaces 20 to bring about the called-for manoeuvre. Such resolution may be made, provided the angular orientation of the surfaces 20 is known. Knowledge of the angular orientation of the wing surfaces 20 is possible by carrying within the missile 10 a reference element (to be described). Suffice it to say here that the angular orientation (i.e. attitude) of such reference element is known at the remote radar station 13.

When operating in a homing guidance mode the missile 10 is aligned so that its heading at the initiation of such mode is approximately along a collision course with target 12. The remote radar station 13 transmits radar frequency energy towards target 12. A portion of the radar frequency energy is reflected by the target 12 and is received by the missile's tracking antenna 22.

Signals representative of angular deviation of the target 12 from the boresight axis of the target tracking antenna 22 are developed in a conventional manner and passed through the receiver/processor 17 to both the target tracking antenna control section 24 and the flight control section 19. These signals are used to drive the radar tracking antenna 22 so as to maintain track of target 12 and also to manoeuvre the missile 10 so as to maintain the missile on a collision course with the target 12. To put in aother way, let it be assumed that the missile 10 is aligned along a collision course with the target 12 and the boresight axis of the target tracking antenna 22. is pointing at the target 12, then if the missile 10 stays on the collision course the boresight axis of the target tracking antenna 22 will remain pointing at the target 12.

However, if the missile 10 deviates from such collision course (for example, if the target 12 manoeuvres) the boresight axis of the target tracking antenna 22 will no longer be pointed at target 12. A boresight error will then be developed, the boresight error being proportional to the change in light of sight angle between the missile 10 and target 12. It is immediately apparent that the target tracking antenna 22 must be driven to null out the boresight error to prevent losing track of target 12. Further, as is known in proportional navigation guidance, it is desirable to have the missile 10 turn from its present flight path at a rate proportional to the rate of change in the line of sight angle. The missile turn rate is produced by having the missile accelerate laterally (i.e. normal to the line of sight) in a direction to null the change in the line of sight angle. That is, the missile 10 must manoeuvre to get back on a collision course with target 12. The lateral acceleration of the the missile 10 is produced by deflecting selected control surfaces 20. In particular, the tracking of the target 12 by the radar tracking antenna 22 and the manoeuvring of the missile 10 to maintain the missile on a collision course is accomplished by developing a signal representative of the boresight error and feeding the signal to both flight control section 19 and target tracking antenna control section 24.

It is noted that the target tracking antenna 22 is gimballed with two degrees of freedom with respect to the missile body. The gimballing is accomplished in a conventional way by mounting two rate sensing gyros on the antenna to sense the inertial rates experienced thereby and by providing drive means to move the target tracking antenna 22 relative to the body of the missile 10. The target tracking antenna 22 is gimballed to prevent the missile's body motion from developing an erroneous boresight error signal.

Referring now to FIG. 2, the flight control section 19 is shown to include pitch, yaw and roll autopilots 26, 27, 28, each one of such autopilots being controlled by receiver/processor 17. The output signals from each one of such autopilots is fed to an actuator section 30. The autopilots 26, 27 and 28 and actuator section may be of any conventional design. The actuator section is mechanically coupled to the control surfaces 20. The control surfaces 20 are pivotably mounted on the missile body. Deflection of one pair of opposing control surfaces (called pitch control surfaces, i.e. those driven by the pitch autopilot) produce manoeuvres of the missile about a pitch axis P and deflection of the other pair of control surfaces (called the yaw control surfaces, i.e. those driven by the yaw autopilot) produce manoeuvres of the missile about a yaw axis Y. The differential deflection of one pair of control surfaces, say the pitch control surfaces, produce rolling of the missile 10 about its longitudinal axis. Such differential deflection is produced when the actuator section 30 responds to command signals from the roll autopilot 28.

The target tracking antenna control section 24 includes a structure 32 on which the target tracking antenna 22 is mounted. The structure 32 is designed to enable the target tracking antenna 22 to move with two degrees of freedom within the missile. In particular, the structure 32 includes a base gimbal 34 fixed to the missile body. A gimbal 36 is pivotly mounted on the base gimbal 34 so as to rotate about an antenna pitch axis P', parallel to the pitch axis P. An inner member 38 is pivotly mounted on the gimbal 36 so as to rotate about an antenna yaw axis Y', perpendicular to the antenna pitch axis P'.

The inner member 38 supports the target tracking antenna 22. The boresight axis of the antenna is orthogonal to both the antenna pitch axis P' and the antenna yaw axis Y'. The target tracking antenna 22, which may be any conventional target tracking antenna, is a monopulse antenna in this embodiment. Therefore, a pair of signals is produced by the antenna, one signal representing the component of the boresight error about the antenna pitch axis P' and the other signals representing the component of the boresight error about the antenna yaw axis Y'. The gimbal 36 and the inner member 38 are coupled to the output of a pitch drive unit 42 and a yaw drive unit 40 respectively, in any convenient way (here shown by dotted lines). That is, the yaw drive unit 40 pivots the inner member 38 about the antenna yaw axis Y'. Likewise, the pitch drive unit 42 pivots the gimbal 36 about the antenna pitch axis P'. The yaw and pitch drives units 40, 42 may be electrical or mechanical motors responsive to electrical signals supplied by the receiver/processor 17. Three antenna rate sensing gyros 44, 46, 48 are mounted on the inner member 38. The input axes of the three antenna rate sensing gyros are disposed along mutually orthogonal axes. In particu lar, the antenna rate sensing gyros 44, 46 and 48 are so oriented that the output signals from the gyros represent the angular rates of the target tracking antenna 22 about the antenna yaw axis Y', the antenna pitch axis P' and its boresight axis respectively. It is therefore apparent that the output signals also provide a measure of the inertial angular rate of the member 32.

Referring now to FIG. 3, the receiver/processor 17 is seen to include a heterodyne receiver 50 fed by the target tracking antenna 22 and a heterodyne receiver and control 52 fed by the downlink antenna 15. The heterodyne receiver and control 52 produces gating signals on either line H or line C in accordance with the guidance mode required for the missile 10. During the homing guidance mode, line H has a gating signal aplied to it, whereas during the command guidance mode line C has a gating signal applied to it.

Let it be assumed that missile 10 is launched in the command guidance mode. It will be first noted that signals from the heterodyne receiver 50 (i.e. from the target tracking antenna 22) are inhibited from passing to either the flight control section 19 or the target tracking antenna control section 24. It will be further noted that the member 32 is attitude stabilized and therefore may be considered as a reference element. That is, during the command guidance mode, inertial rates experienced by the member 32 are sensed by the antenna rate gyros 44, 46, 48 and signals representative of such inertial rates are fed to means (to be described) for returning the member 32 to the angular orientation it has prior to experiencing such inertial rates. In particular, the signals representative of inertial rates about the antenna yaw axis Y' and antenna pitch axis P' sensed by antenna rate sensing gyros 44 and 46 are fed, respectively, through integrators 54 to the yaw drive unit 40 and pitch drive unit 42.

The roll orientation of structure 32 is attitude stabilized by coupling the output signals from the antenna rate gyro 48 to the roll autopilot 28. The roll inertial stabilization of the missile 10 through use of the roll autopilot 28 is possible because the aerodynamic response of the missile 10 to differential deflection of the pitch control surfaces is fast enough to inertially stabilize the structure 32 in roll. Therefore, because a reference element having an angular orientation which is known at the remote radar station 13 (FIG.

1) is contained within the missile 10, the missile may be guided with guidance signals transmitted from remote radar station 13. In particular, guidance signals transmitted from the remote radar station 13 are received by the downlink antenna 15 and then processed by heterodyne receiver and control 52. Such guidance signals contain the following information: (1) the guidance mode in which the missile should operate (i.e. command or homing) as determined by the range of the missile 10 from the target 12; (2) the manoeuvre signals referenced to the known angular orientation of the reference element (i.e. member 38); and (3) command signals for positioning the member 38 in a known orientation to constrain the member within limits allowable by the physical space provided for it within the missile 10. Guidance mode selection information is used to determined whether a gating signal should be applied to line C or to line H. Manoeuvre signals are converted into electrical signals on line 56, such signals being fed to a computer 58 where they are converted into acceleration commands referenced to the orientation of the control surfaces, 20, (i.e.

the orientation of the missile's body). The conversion of the manoeuvre signals from "reference element" coordinates to "missile body" coordinates is made by providing in the missile 10 conventional angle transducers 59, 60. These transducers are suitably mounted to provide a measure of the orientation of the inner member 38 and gimbal 36 with respect to the missile body. The outputs of the transducers 59, 60 are fed to the computer 58 together with the manoeuvre signals on line 56. The properly resolved signals pass through gates 62 and 64 to the pitch and yaw autopilots 26, 27. Command signals for positioning the structure 32 in a known orientation are processed by the heterodyne receiver and control 52 and appear as angle command signals on lines 66, 68, 70. Such signals are fed to error computers 72. Let us consider the signal on line 66.

This signal represents the desired "yaw" position of the structure 32 in inertial space.

If the "desired" position and the "actual" position are not equal an error signal appears on line 74. This error signal passes through gate 75 and a summer 77 to the yaw drive unit 40. Command of the "pitch" position of the structure 32 is accomplished in a similar way as shown. For roll, however, the error signal out of the error computer 72 is fed through gate 76 to the roll autopilot 28. The roll autopilot 28 develops signals in response to this error signal, whereby the pitch control surfaces are differentially deflected. The missile's roll attitude is thereby changed because of such differential deflection of the pitch control surfaces, such differential deflection continuing until the error signal is nulled.

Let us now assume that the range between the missile 10 and target 12, as determined by the remote tracking station 13, has been reduced to where it is desirable to have the missile fly in the homing guidance mode.

The missile may be assumed to be aligned along an approximate collision course with target 12. A signal is transmitted to the missile 10 from the remote station whereby the heterocyne receiver and control 52 produces a gating signal on line H and removes the gating signal from line C. Therefore, signals from the error computers 72 and computer 58 are inhibited from passing to the yaw drive unit 40, the pitch drive unit 42, the roll autopilot 28, the pitch autopilot 26 and the yaw autopilot 27. The yaw boresight error signals and pitch boresight error signals developed by the target tracking antenna 22 are heterodyned in the heterodyne receiver 50 to produce video frequency signals. The video frequency signals corresponding to the pitch boresight error and yaw boresight error appear on lines 78, 80, respectively. These signals are fed to the pitch autopilot 26 and yaw autopilot 27 through gates 82 and 84.

The signals on line 80 are also fed through a gate 86, via summer 77, to the yaw drive unit 40 and the signals on line 78 are fed through a gate 88, via summer 89, to the pitch drive unit 42 so that target tracking antenna 22 may maintain track of the target 12. Signals from the antenna yaw gyro 44 are passed through a gate 90 to the yaw drive unit 40 additionally to the signals passing through gate 86. Likewise signals from the antenna pitch gyro 46 are fed through gate 92 to the pitch drive unit 42 additionally to the signals passing through the gate 88. The signals from the antenna pitch and yaw gyros 44, 46 are used therefore to sense the inertial rates of the target tracking antenna 22 and thereby of the antenna with respect to the missile body for reasons previously discussed.

Referring now to FIG. 4, an alternative embodiment is shown. In this embodiment the missile 10 is shown as operating in the command guidance mode. However, the tracking signals are generated from the signals received by the target tracking antenna 22 rather than signals received from the target at the remote radar station 13.

Thus signals received by the target tracking antenna 22 are fed to the heterodyne receiver 50 and then are retransmitted by a transmitter 100 back to the remote radar station 13.

The signals from the transmitter 100 pass through a circulator 102 and the downlink antenna 15 to the remote tracking station 13.

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

The guidance signals. are transmitted from the remote radar station 13 and are received by the downlink antenna 15. The signals received by downlink antenna 15 are processed and responded to by the missile in the manner described with reference to FIG. 3. It is here noted that the target tracking antenna 22 (and structure 32) serve, in addition to a target tracking element, as the reference element required during the command guidance mode. The orientation control and inertial stabilization of the reference element is achieved by means equivalent to those previously discussed in the command guidance mode of FIG. 3.

While the invention has been described using rate sensing gyros affixed to the target tracking antenna 22, it will be apparent that such gyros and and the integrators coupled thereto may be replaced by rate integrating gyros.

WHAT WE CLAIM IS: 1. A vehicle guidance apparatus, whereby a vehicle is guided during a portion of flight towards a target in a command guidance mode and during another portion of the flight in a homing guidance mode, the apparatus comprising target tracking means, including a target tracking element carried by the vehicle, for generating guidance signals for the vehicle during the homing guidance mode; first means, fixed relative to the target tracking element, for sensing the inertial angular rates of this element about a pitch axis and yaw axis; means, responsive to the first means, for gimballing the target tracking element with respect to the vehicle body about the pitch axis and the yaw axis during the homing guidance mode; second means, fixed relative to the target tracking element, for sensing the inertial angular rate of the target tracking element about an axis orthogonal to both the pitch axis and the yaw axis; and means, responsive to the first means and the second means, for stabilizing the attitude of the target tracking element at a known angular orientation during the command guidance mode.

2. Apparatus according to claim 1, wherein the stabilizing means includes means, responsive to the second means, for controlling the attitude of the vehicle aerodynamically in accordance with the inertial angular rate sensed by the second means.

3. Apparatus according to claim 2, wherein the controlling means is a roll autopilot carried within the vehicle.

4. Apparatus according to claim 1, 2 or 3, including further stabilizing means, coupled to the gimballing means and the first said stabilizing means, for positioning the target tracking element at a predetermined angular orientation in the command guidance mode.

5. Apparatus according to any of claims 1 to 4, wherein the first said stabilizing means includes means coupled to the first means and the second means, for integrating the inertial angular rates about the pitch axis, the yaw axis and the axis orthogonal thereto.

6. A vehicle guidance apparatus substantially as hereinbefore described with reference to and as illustrated in Figs. 1 to 3 or Figs. 1 to 3 as modified by Fig. 4 of the accompanying drawings.

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

Claims (6)

**WARNING** start of CLMS field may overlap end of DESC **. the heterocyne receiver and control 52 produces a gating signal on line H and removes the gating signal from line C. Therefore, signals from the error computers 72 and computer 58 are inhibited from passing to the yaw drive unit 40, the pitch drive unit 42, the roll autopilot 28, the pitch autopilot 26 and the yaw autopilot 27. The yaw boresight error signals and pitch boresight error signals developed by the target tracking antenna 22 are heterodyned in the heterodyne receiver 50 to produce video frequency signals. The video frequency signals corresponding to the pitch boresight error and yaw boresight error appear on lines 78, 80, respectively. These signals are fed to the pitch autopilot 26 and yaw autopilot 27 through gates 82 and 84. The signals on line 80 are also fed through a gate 86, via summer 77, to the yaw drive unit 40 and the signals on line 78 are fed through a gate 88, via summer 89, to the pitch drive unit 42 so that target tracking antenna 22 may maintain track of the target 12. Signals from the antenna yaw gyro 44 are passed through a gate 90 to the yaw drive unit 40 additionally to the signals passing through gate 86. Likewise signals from the antenna pitch gyro 46 are fed through gate 92 to the pitch drive unit 42 additionally to the signals passing through the gate 88. The signals from the antenna pitch and yaw gyros 44, 46 are used therefore to sense the inertial rates of the target tracking antenna 22 and thereby of the antenna with respect to the missile body for reasons previously discussed. Referring now to FIG. 4, an alternative embodiment is shown. In this embodiment the missile 10 is shown as operating in the command guidance mode. However, the tracking signals are generated from the signals received by the target tracking antenna 22 rather than signals received from the target at the remote radar station 13. Thus signals received by the target tracking antenna 22 are fed to the heterodyne receiver 50 and then are retransmitted by a transmitter 100 back to the remote radar station 13. The signals from the transmitter 100 pass through a circulator 102 and the downlink antenna 15 to the remote tracking station 13. The retransmitted signals are processed by the computer at the remote radar station 13 to determine guidance signals for the missile. The guidance signals. are transmitted from the remote radar station 13 and are received by the downlink antenna 15. The signals received by downlink antenna 15 are processed and responded to by the missile in the manner described with reference to FIG. 3. It is here noted that the target tracking antenna 22 (and structure 32) serve, in addition to a target tracking element, as the reference element required during the command guidance mode. The orientation control and inertial stabilization of the reference element is achieved by means equivalent to those previously discussed in the command guidance mode of FIG. 3. While the invention has been described using rate sensing gyros affixed to the target tracking antenna 22, it will be apparent that such gyros and and the integrators coupled thereto may be replaced by rate integrating gyros. WHAT WE CLAIM IS:

1. A vehicle guidance apparatus, whereby a vehicle is guided during a portion of flight towards a target in a command guidance mode and during another portion of the flight in a homing guidance mode, the apparatus comprising target tracking means, including a target tracking element carried by the vehicle, for generating guidance signals for the vehicle during the homing guidance mode; first means, fixed relative to the target tracking element, for sensing the inertial angular rates of this element about a pitch axis and yaw axis; means, responsive to the first means, for gimballing the target tracking element with respect to the vehicle body about the pitch axis and the yaw axis during the homing guidance mode; second means, fixed relative to the target tracking element, for sensing the inertial angular rate of the target tracking element about an axis orthogonal to both the pitch axis and the yaw axis; and means, responsive to the first means and the second means, for stabilizing the attitude of the target tracking element at a known angular orientation during the command guidance mode.

2. Apparatus according to claim 1, wherein the stabilizing means includes means, responsive to the second means, for controlling the attitude of the vehicle aerodynamically in accordance with the inertial angular rate sensed by the second means.

3. Apparatus according to claim 2, wherein the controlling means is a roll autopilot carried within the vehicle.

4. Apparatus according to claim 1, 2 or 3, including further stabilizing means, coupled to the gimballing means and the first said stabilizing means, for positioning the target tracking element at a predetermined angular orientation in the command guidance mode.

5. Apparatus according to any of claims 1 to 4, wherein the first said stabilizing means includes means coupled to the first means and the second means, for integrating the inertial angular rates about the pitch axis, the yaw axis and the axis orthogonal thereto.

6. A vehicle guidance apparatus substantially as hereinbefore described with reference to and as illustrated in Figs. 1 to 3 or Figs. 1 to 3 as modified by Fig. 4 of the accompanying drawings.