Abstract:
The invention relates to a seeker head for target tracking missiles having an image resolving seeker being gimbal suspended in a seeker gimbal assembly and adapted to be aligned to a target by target deviation signals, and inertial sensors. A virtual inertially stabilized reference coordinate system is adapted to be defined from signals from the image resolving seeker and from the seeker gimbal assembly, said stabilized reference coordinate system having an axis aligned to said target. The stabilized reference coordinate system is adapted to be aligned to predicted target positions in case of deterioration of the tracking function of the seeker to the target in accordance with the line of sight information (e.g. direction, angular rate, angular acceleration) of the reference coordinate system then present. The seeker is adapted to be aligned to the axis of the reference coordinate system when the deterioration ceases, the signals from the seeker taking over the tracking function of the seeker again.

Description:
BACKGROUND OF THE INVENTION 
     This invention relates to a seeker head for target tracking missiles having an image resolving seeker being gimbal suspended in a seeker gimbal assembly and adapted to be aligned to a target by target deviation signals, and inertial sensors, 
     Target tracking missiles are known having an image resolving sensor, e.g. in the form of a detector matrix having a two-dimensional array of detector elements. This seeker is gimbal suspended in a seeker gimbal assembly. Inertial sensors respond to the angular movements of the missile in inertial space. Torquers act on the gimbals of the seeker gimbal assembly and decouple the seeker from the thus determined angular movements of the missile. An image of an object scene is generated on the detector matrix. Target deviation data of a target located in the object scene, e.g. an enemy aircraft to be attacked, are generated by image processing of this image. The target deviation data represent the deviation of the target from an optical axis of the seeker. By means of these target deviation data the seeker tracks the target. From the tracking the angular rate of the line of sight is determined. From the angular rate of the line of sight, in turn, steering signals for the missile are derived. By means of a helmet visor a target recognized by the pilot is designated to the seeker. The missile is guided to this target in the described manner. 
     During air combats with close curves (“close-in-combat”) it is desirable to detect a target even at a large look angle of the seeker. However, the look angle of the seeker is, of course, limited by the design. During air combats with close curves, situations can arise, in which the target occurs under an angle of vision, which is larger than the maximum allowable look angle of the seeker. Then the target cannot be designated to the seeker head. During the further course of the curved flight, the angle of sight can be reduced to a value below the maximum allowable look angle. Then the target can be designated to the seeker head and the missile can be fired. The earlier this is made, the greater are the chances of hitting the target. If, however, the missile is fired, then it first has the tendency to align aerodynamically with the direction of the velocity vector of the missile. Then the angle of vision to the target can again exceed the maximum allowable look angle of the seeker, such that the target gets lost. The target can also be covered temporarily by clouds. 
     SUMMARY OF THE INVENTION 
     One of the objects of the present invention is hence to provide a seeker head for target tracking missiles such that, even when the target tracking is disturbed for a short time, the seeker is re-aligned to the target as soon as the disturbance ceases. 
     This object is achieved in that a virtual inertially stabilized reference coordinate system is defined from signals from the image resolving seeker and from the seeker gimbal assembly, the stabilized reference coordinate system having an axis pointing to said target, the stabilized reference coordinate system is caused to point to predicted target positions, in case of disturbance of the target-tracking function of the seeker, in accordance with the line of sight information (e.g. direction, angular rate, angular acceleration) of said reference coordinate system then present, and the seeker is aligned with the axis of the reference coordinate system, when the disturbance ceases, the signals from the seeker resuming the tracking function of the seeker again. 
     Thus, according to the invention, a reference coordinate system is permanently defined, the axis of which points to the target. This is a type of “virtual” seeker. Normally, this reference coordinate system follows the target in the same manner as the seeker tracks the target from the deviation data. If the tracking movement of the seeker to the target is deteriorated, e.g. when the seeker attains its maximum allowable look angle or when the seeker temporarily cannot “see” the target anymore due to clouds, the reference coordinate system tracks a predicted target position. The predicted target position is determined by a kind of extrapolation from the line of sight information determined immediately before the deterioration occurs. When the deterioration then ceases, that means, for example, that the target occurs under an angle of vision falling below the maximum allowable look angle again, the seeker is aligned with the reference coordinate system. Then the seeker again detects the target, which target has been lost for a short time in its field of view. Then the seeker again tracks the target exactly by means of the deviation data supplied by the image processing. 
     Further objects and features of the invention will be apparent to a person skilled in the art from the following specification of a preferred embodiment when read in conjunction with the appended claims. 
    
    
     BRIEF DESCRIPTION OF THE DRAWING 
     The invention and its mode of operation will be more clearly understood from the following detailed description when read with the appended drawing in which: 
     FIG. 1 shows an example of a situation, in which, during air combats with close curves, the tracking function of the seeker to the target and the target designation of a target tracking missile can be deteriorated by limitation of the look angle of the seeker to a maximum allowable value; 
     FIG. 2 shows an example of another situation, in which, during air combats with close curves, the tracking function of the seeker to the target and the target designation of a target tracking missile can be deteriorated by limitation of the look angle of the seeker to a maximum allowable value; 
     FIG. 3 shows the geometry when a missile is fired by an aircraft; 
     FIG:  4  is a schematic illustration of an infrared-sensitive seeker in a target tracking missile; 
     FIG. 5 schematically shows the tip of a missile having a seeker head and illustrates the limitation of the look angle; 
     FIG. 6 is a simplified block diagram and shows the generation of increments of the angular rate of the line of sight for the tracking function of the reference coordinate system; and 
     FIG. 7 is a simplified block diagram and shows the illustration of a missile-fixed system (s) relative to an inertial system and a reference coordinate system (r) relative to the missile system. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring now to FIG. 1, there is shown an air combat situation, in which a combat aircraft  10  moves along a narrow circular trajectory  12 , which is curved about a point  14 . An enemy combat aircraft  16  (target) moves along a likewise narrow circular trajectory  18 , which is curved about a point  20  located relatively far away from the point  14 . Both of the combat aircrafts  10  and  16  follow the circular trajectories clockwise. On a narrow circular trajectory  12  or  18 , the combat aircrafts  10  and  16 , respectively, fly with large load factor and, thus, as illustrated, with large angle of attack. This means that the longitudinal axis  30  (aircraft datum line) of the combat aircraft  10  forms an angle with the velocity vector. 
     Numeral  22 ,  24 ,  26  and  28  designate lines of sight from the combat aircraft  10  to the target  16 , which lines of sight exist at different moments. It can be seen that the enemy combat aircraft (target)  16  occurs, as seen from the combat aircraft  10 , at first at an angle of vision &gt;90°. This results in the line of sight  22 . The line of sight  24  extends at an angle of vision of 90° with respect to the longitudinal axis  30  of the combat aircraft  10 . With regard to the lines of sight  26  and  28 , the angle of vision, at which the enemy combat aircraft  16  occurs to the pilot and to the seeker of a missile provided on the combat aircraft  10 , is getting smaller and smaller during the further course of the trajectories  12  and  18 . There is a maximum angle of vision, under which the target, namely the enemy aircraft  16 , can be designated to the missile by the pilot by means of a helmet visor. This maximum angle of vision for the target designation is, for example, near by 90° and, thus, corresponds to the line of sight  24 . 
     With reference to FIG. 4, there is shown a seeker  32  of a target tracking missile  34  (FIG.  5 ). The seeker  32  comprises an image resolving detector  36  responding to infrared radiation and an imaging optical system  38 . As illustrated in FIG. 5, the seeker  32  is pivotable by a seeker gimbal assembly  40  about a pitch axis  42  relative to the longitudinal axis  44  of the missile  34 . Furthermore, a rotation of the seeker  32  about this longitudinal axis  44  (roll axis) is possible. The seeker  32  has an optical axis  46 . The angle between the optical axis  46  of the seeker  32  and the longitudinal axis  44  of the missile  34  is called “look angle”. Due to the construction the look angle is limited to a “maximum allowable look angle”, as can be seen in FIG.  5 . The seeker  32  is located behind a transparent dome-shaped window, the “dome”  48 , in the tip of the missile  34 . The maximum allowable look angle is, for example, determined by the fact that the imaging path of rays of the imaging optical system  38  has to at least partly pass through the dome  48 . 
     The pilot now has to try to catch the enemy combat aircraft  16  as soon as possible, that is at a large angle of vision in the example of FIG. 1, and to designate the target to the target tracking missile  34 . The earlier the missile  34  is fired, the larger is the probability of success of shooting down the enemy combat aircraft  16 . The limitation of the look angle acts as deterioration. 
     FIG. 2 shows a similar air combat situation as in FIG.  1 . Corresponding elements are designated by the same reference numerals in FIG. 2 as in FIG.  1 . In this air combat situation the points  14 A and  20 A, about which the two trajectories  14 A and  18 A are curved, are located close together. 
     A further problem arises because the missile  34  after the firing and release of the steering system has the tendency to at first be oriented with its longitudinal axis  44  in the direction of the velocity vector  50  of the combat aircraft  10 . Thereby, the angle of vision to the target can be increased to an angle, which is larger than the maximum allowable look angle, even if this angle of vision is smaller than the maximum allowable look angle and the seeker  32  of the missile  34  can detect the enemy combat aircraft  16  when the missile  34  is fired. 
     This is illustrated in FIG:  3 . In FIG:  3  the longitudinal axis (“aircraft datum line”) of the combat aircraft  10  is designated by  30 . A straight line  44 A designates the longitudinal axis of the missile  34  (missile boresight”) in the launcher, that means before firing. The straight line  44 A generally forms a small angle with the longitudinal axis  30 . Numeral  54  designates the line of sight from the center of mass of the combat aircraft  10  to the target. This line of sight  54  forms an angle α (“lag angle”) with the velocity vector  50 . Numeral  58  designates the line of sight from the seeker  32  of the missile  34  to the target. This line of sight  58  is parallel with the line of sight  54  and forms an angle β (“missile off-boresight angle at launch”) with the longitudinal axis  44 A of the missile  34 . Numeral  60  designates the line of sight from the helmet visor of the pilot to the target. This line of sight  60  is almost parallel to the lines of sight  54  and  58 . The line of sight  60  forms an angle γ (“destinator off-boresight angle at launch”) with the longitudinal axis  30  of the combat aircraft  10 . Numeral  60  designates the line of sight from the seeker  32  of the missile  34  to the target at the time when the control surfaces are unlocked after firing. Also this line of sight  62  is parallel to the lines of sight  54 ,  58  and  60 . The line of sight  62  forms an angle δ (“off-boresight angle at control unlock”) with the longitudinal axis  44  of the missile  34 . 
     Before firing the missile  34 , the angle β is smaller than the maximum allowable look angle. Therefore, the seeker  32  detects the target and can track the target resulting in a measured angular rate of the line of sight. As can be seen from FIG. 3, the missile  34  is oriented, after the firing, at first with its longitudinal axis  44  substantially in the direction of the velocity vector  50 . At the time when the steering is unlocked, the line of sight angle δ temporarily becomes &gt;90° again and larger than the maximum allowable look angle of the seeker  32  (FIG.  5 ). The seeker  32  cannot “see” the target anymore. Again, a “deterioration” of the tracking function occurs. 
     As can be seen from FIG. 5, three coordinate systems are defined, which are represented by their respective x-axes in FIG. 5. A missile coordinate system having the axis x 5  is missile-fixed. The x s -axis corresponds to the longitudinal axis  44  of the missile. A seeker coordinate system having the axis x h  is seeker-fixed. The x h -axis corresponds to the optical axis of the seeker  32 . A third coordinate system having the axis x r  is a virtual reference coordinate system, which is determined by calculation. Furthermore, there is an inertial system, that means a coordinate system which, with respect to its orientation, is stationary in inertial space. 
     In FIG. 6 the seeker, that is an image resolving electro-optical unit, is mounted in the missile  34  through a seeker gimbal assembly  40 . Numeral  62  designates a missile-fixed inertial sensor unit. The inertial sensor unit  62  can be constructed with gyros, laser gyros or other inertial sensors responding to angular rates. The inertial sensor unit  62  supplies angular rates p, q and r about three missile-fixed axes. 
     The seeker  32  supplies image data at an output  64 . The image data are applied to an image processing system  66 . The image processing system  66  supplies deviation data corresponding to a target deviation in the seeker-fixed coordinate system, which deviation data can be represented by a vector ε h . These deviation data ε h  are applied to means  68  for coordinate transformation. The means  68  for coordinate transformation receive, on one hand, gimbal angles from the seeker gimbal assembly, as illustrated by the connection  70 . On the other hand, the means  68  for coordinate transformation also receive direction cosine data corresponding to a direction cosine matrix C r   s . The direction cosine matrix C s   r  represents the rotation from the reference coordinate system to the seeker coordinate system, as will be described later. The means  68  for coordinate transformation then supply deviation data with respect to the reference coordinate system. These deviation data ε r  are applied to an estimator filter  72 . The estimator filter  72  supplies increments Δσ y  and Δσ z  of the angular rate of the line of sight. 
     The increments Δσ y  and Δσ z  of the angular rate of the line of sight are applied to means  74  for defining a reference coordinate system. Initial look angles λ y0  and λ z0  are applied to means  76  for defining an initial position of the reference coordinate system. In this initial position of the reference coordinate system the look angles λ are still smaller than the maximum allowable look angle. The seeker  32  still detects the target. The data of the initial position of the reference system are likewise applied to the means  74  for defining the reference coordinate system. 
     In the illustrated preferred embodiment, the reference coordinate system is represented by a quaternion having the elements I r0 , I r1 , I r2  and I r3 . Correspondingly, also the initial position of the reference coordinate system is represented by a quaternion q r0 . The means  74  for defining the reference coordinate system, at the same time, achieve scaling. 
     The inertial sensor unit  40  supply the three angular rates p, q and r about three missile-fixed axes. The scanning of the angular rates p, q and r in a fixed clock cycle supplies angle increments ΔΦ x , ΔΦ y  and ΔΦ z . The scanning with a fixed clock cycle is symbolized in FIG. 7 by a three-pole switch  78 . The angle increments ΔΦ x , ΔΦ y  and ΔΦ z  are applied to means  80  for representing a missile coordinate system. The position of the missile coordinate system is related to an inertial system. The missile coordinate system is likewise defined by a quaternion. This quaternion has the elements I i0 , I i1 , I i2  and I i3 . 
     The quaternion from the means  74  representing the reference coordinate system and the quaternion from the means  80  representing the missile coordinate system, that means the elements I i0 , I i1 , I i2  and I i3  are “multiplied” by multiplication means  82 . The multiplication of the quaternions supply the relative position of the missile coordinate system and the reference coordinate system. This is represented by a quaternion q r   s . 
     The quaternion q r   s  representing the relative position between the missile coordinate system and the reference coordinate system is likewise applied to means  86  for forming the associated direction cosine matrix C r   s . 
     The direction cosine matrix C r   s  provides the position of the reference coordinate system relative to the missile. As illustrated in FIG. 6, this direction cosine matrix C r   s  is applied to means  68  for coordinate transformation. Thus, these means  68  for coordinate transformation provide the deviation data with respect to the reference coordinate system. From the elements of the direction cosine matrix C r control signals for the seeker gimbal assembly  40  are obtained, such that this movement of the missile  34  is compensated for at the seeker  32  and the seeker  32  is decoupled from the movements of the missile  34 . 
     The described seeker head operates as follows: 
     In the normal operation, when the seeker  32  detects the target and follows it with a look angle smaller than the maximum allowable look angle, the seeker coordinate system with axis x h  and the reference coordinate system with the axis x r  approximately coincide. When the seeker  32  has reached the maximum allowable look angle, then the seeker  32  is stopped in its position. The reference coordinate system, however, moves further relative to the missile  34 . This movement is determined by the angular rate of the line of sight, which was valid when the maximum allowable look angle had been attained. This angular rate of the line of sight supplies further increments ΔΦ y  and ΔΦ z  to the means  74  for defining the reference coordinate system in inertial space. By this, the reference coordinate system is tracked to a predicted position of the target. It is assumed that the angular rate of the line of sight in inertial space substantially remains constant for a short period of time. The predicted positions are obtained by a kind of extrapolation. By the multiplication of the quaternions by means of the multiplication means  82 , the position of the reference coordinate system relative to the missile is obtained. When the thus calculated look angle of the reference coordinate system becomes smaller than the maximum allowable look angle again, then the real seeker  32  is aligned according to this reference coordinate system. Thus, the seeker  32  is directed to the predicted positions of the target. It can be assumed that these predicted positions are located in the proximity of the real target and, thus, the target is detected in the field of view of the seeker  32  again. 
     In the situation illustrated in FIG. 3, the seeker  32  at first loses the target after the firing of the missile  34 , because the angle of vision δ to the target is increased beyond the maximum allowable look angle of the seeker  32  due to the alignment of the seeker  34  with the velocity vector  50 . The axis x r  of the reference system is, as described, aligned to the predicted position of the target. However, after the control surfaces has been unlocked, the missile  34 , taking the last angular rate of the line of sight measured by the seeker  32  as a basis, is guided such that it tracks the target. Thus, the missile  34  is rotated to the direction to the target. Thereby, the “angle of vision” of the “virtual seeker” represented by the reference coordinate system is reduced again. The angle of vision falls below the maximum allowable look angle. Due to this, as described, the seeker  32  can be aligned according to the reference coordinate system again and can detect the target. 
     The use of quaternions for representing the coordinate systems avoids singularities, which would appear at a took angle of 90° when using other representations.