Abstract:
Method and apparatus are described for improving the impact angle between a missile and stationary or moving target in the presence of stationary or adverse wind conditions. A triggered bias detector provides an enabling signal to a bias resolver when the pitch line of sight rate of a missile exceeds a predetermined level. The bias resolver supplies first and second bias signals to the pitch forward guidance loop and yaw forward guidance loop in response to the detected pitch line of sight rate. The missile is given a lofting trajectory in response to the applied bias signal as it closes on the target. Impact angles of a more nearly vertical condition over a wide range of missile/target acquisition geometric conditions are obtained using this triggered bias technique over conventional constant gravity bias proportional guidance techniques.

Description:
The present invention is related to guidance systems for terminally guided missiles. Specifically, apparatus and method are provided for biasing the trajectory of a guided missile to increase the top attack angle for certain target conditions. 
     Guided missile weapons have been developed which include proportional guidance in two planes of movement. The proportional guidance control for these systems develops control signals for guiding the missile to a target based upon a sensed line of sight rate with respect to the target. The error information for driving the multiaxis proportional guidance system is developed from a seeker which senses infrared energy, laser energy, radio frequency energy or millimeter wave energy emitted or reflected by the target. 
     Studies conducted by the inventors have shown that the effectiveness of a terminally guided missile with a shallow approach angle requires that the angle of impact be greater than 45° for many target conditions. The higher angle provides for a greater target penetration by impacting areas of the target such as the engine compartment or exit hatch of a military tank where armor protection is necessarily weak. For other target conditions, such as closing targets moving towards the missile, a lofting trajectory which results in a high angle of attack and large seeker gimbal angles, is undesirable. In view of the foregoing design requirements, a bias to the missile trajectory is desirable for those target conditions which require a lofting terminal trajectory. The bias, however, should not be constant, but only supplied for target conditions which require an improved angle of impact based on sensed line of sight rate with respect to the target, increasing the probability of total target destruction. The condition for establishing the bias should be based on a sensed in-flight target to missile geometric relationship. 
     SUMMARY OF THE INVENTION 
     An object of the present invention is to provide trajectory shaping resulting in high impact angles for a terminally guided missile system. 
     A more specific object of the invention is to provide a trajectory bias operative in response to a sensed missile-target geometric relationship. 
     These and other objects are provided by apparatus and methods which, in response to a sensed flight condition, biases the missile terminal trajectory to achieve a high impact angle with the target. A guided missile having proportional guidance control developed from a sensed relative target position and velocity is equipped to sense one flight parameter. A bias to the trajectory is established when the sensed flight parameter exceeds a predetermined minimum. Impact angles of a more nearly vertical condition over a wide range of missile/target acquisition geometric conditions are obtained using this triggered bias technique over conventional constant gravity bias proportional guidance technology. 
     In a preferred embodiment of the invention, the sensed parameter is the missile pitch line of sight rate which triggers first and second bias levels into summation with the missile pitch and yaw forward guidance control signals. The missile bias levels are proportional to the sine and cosine of the roll angle which is detected by on-board instruments during the initial target acquisition phase of the trajectory. The bias levels may also be combined with a gravity bias parameter to further modify the trajectory based upon a predetermined gravity influence related to the missile aerodynamic coefficients. 
    
    
     DESCRIPTION OF THE FIGURES 
     FIG. 1 is a block diagram illustrating the proportional guidance of a guided missile of a preferred embodiment of the invention. 
     FIG. 2 is a flow diagram more specifically describing the bias signal generating technique of the preferred embodiment. 
     FIG. 3 is an illustration of the LOS rate of a terminally guided missile as a function of the geometric relationship between the missile and target. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENT 
     Referring now to FIG. 1, there is shown a block diagram or a proportional guidance system incorporating the preferred embodiment of the invention. The guidance system includes three axis rate control of a guided missile in the pitch, yaw and roll planes. Signals which are supplied to the control system from a conventional seeker head indicative of the relative position of the target and missile include: 
     
       
         Pitch {dot over (λ)}=pitch angle line of sight rate (°/sec) 
       
     
     
       
         Yaw {dot over (λ)}=yaw angle line of sight rate (°/sec) 
       
     
     Other conventional signals which participate in control of the missile include pitch rate and yaw rate feedback signals Q and R generated from on-board gyroscopic instruments. 
     Fin actuators  40 ,  41  and  42  are represented by second order Laplace functions and receive drive commands to reduce the angular error between the missile position and target position in the three guidance planes to zero, thus impacting a sensed target. 
     The guidance system of FIG. 1 is represented in conventional control system functional blocks and will be presently explained in terms of these functional blocks. The system may be realized with hardware components having a transfer function corresponding to each functional block, but may also be advantageously implemented by microcomputer techniques which perform in software the functional blocks of FIG.  1 . Sampling of the input seeker signals, performing the required software functions for each functional block, and supplying multiplexed digital signals to each fin control input  40 ,  41  and  42  are techniques known in the control system field in which the present invention may be implemented and will not be further described. 
     The control system of FIG. 2 includes a rate bias signal generator  10  which provides a pitch rate bias signal and yaw rate bias signal defined as 
     
       
         BIASP  Δ Trigger Bias Value×COS(φref) 
       
     
     
       
         BIASY  Δ Trigger Bias Value×SIN(φref) 
       
     
     where φ ref is the roll reference angle of the missile at target acquisition, detected during the initial portion of the trajectory. The trigger bias value (B.V.) is typically 3°/sec. A resolver  16  supplies each of these signals to the summing junctions  11 ,  12  when switches  14 ,  15  are closed. Switches  14  and  15  are closed and resolver  16  enabled when the PLOS rate is detected to be greater than or equal to a preselected threshold value, and opened for PLOS rates less than the threshold value. This threshold value is unique for different missile-target geometric and aerodynamic configurations. The threshold value, T.V., may be −1.0°/ sec. The summing junctions  11  and  12  receive seeker pitch line of sight rate signals, PLOS rate, and yaw seeker line of sight signals, YLOS rate, for deriving pitch and yaw proportional error signals Perr and Yerr. 
     The pitch and yaw error signals, Perr and Yerr, are applied to pitch and yaw compensation circuits  19 ,  20 . The compensation networks provide a transfer function of              K   P        1     +         τ   PN        S       1   +       τ   PD        S           ;         K   Y        1     +         τ   YN        S       1   +       τ   YD        S                 where             K   P     =     pitch                 guidance                 gain       ,                  K   Y     =     yaw                 guidance                 gain       ,                  τ   PN     =     pitch                 time                 constant                 numerator       ,                  τ   PD     =     pitch                 time                 constant                 denominator       ,                  τ   YN     =     yaw                 time                 constant                 numerator       ,                  τ   YD     =     yaw                 time                 constant                   denominator   .                                
     The compensated error signals are combined in summing junctions  21 ,  25  with feedback signals from amplifiers  22  and  23  corresponding to the pitch rate and yaw rate, V P2  and V Y2 , detected by on-board gyroscopic sensors. The combined signals from junctions  21 ,  25  are further gain normalized in networks  27 ,  28  to provide control signals δpc and δyc. Limiters  29  and  30  provide limiting to signals δpc and δyc to prevent a control command from being applied to fin deflection controls  40 ,  41  which exceeds the capability of the fin controls to respond. The limited output signal from limiter  30  is combined with a portion of the roll control signal from compensation network  39  in summing junction  44 . The resulting signal is applied to fin deflection control  41  as the yaw control signal. 
     Complementing the pitch and yaw proportional guidance controls is a roll guidance loop. The roll guidance control derives a roll error signal from the pitch and yaw error signals, Perr and Yerr, by the function of block  32 ,          tan     -   1              (   Yerr   )       (   Perr   )                              
     to derive the roll error Δφ c . The roll error is further gain normalized by amplifier  33  having a gain of K φ . The resulting signal is applied to limiter  36  which also limits the magnitude of its respective control signal to a level within the response of fin deflection control  42 . The limited signal is further subtractively combined in junction  37  with a gain normalized feedback roll rate V R3 . The feedback signal is representative of the roll rate of the vehicle gain normalized by the factor K {dot over (φ)}  of network  34 . The drive signal produced by junction  37  is further gain modified by K in amplifier  38 . A compensation network  39  having the following transfer function receives the gain normalized signal:          1   +       τ   RN     ×   S         1   +       τ   RD     ×   S                              
     where τ RN  is the roll forward loop numerator time constant and τ RD  is the roll forward loop denominator time constant. The compensated drive signal δrc is summed negatively with the limited yaw control signal in junction  43  to generate the fin control drive signal. 
     The foregoing three axis proportional guidance system responds to the relative target position and velocity such as to reduce the Perr and Yerr error signals as the missile approaches the target. The trajectory is biased to provide a high impact angle with the target when the pitch line of sight rate exceeds a predetermined rate as determined by trigger bias/logic circuit  18 . There is a fundamental relationship between PLOS rate and the target slant range and target offset. Referring to FIG. 3, there is shown the fundamental relationship between the LOS rate of one type of terminally guided missile such is ejected from a terminally guided warhead, and the geometric position of the missile and target. The missile coordinates are X M  and Z M , and the target coordinates are X T , Z T . The slant range to the target R s  is defined as the straight line distance to the target. The horizontal distance, or ground distance, from missile to target is the referred target offset. X r  and the vertical distance from the missile to the target is the reference target offset Z r . 
     Studies have shown that the larger seeker gimbal angles and angle rates experienced for closing targets, those down range from the missile shown as having a negative target offset, render it unadvisable to bias the trajectory. As FIG. 1 illustrates, this condition occurs for an LOS rate less than −1.0°/sec for this particular missile. Those targets which produce a PLOS rate of less than −1°/second have, by geometry, a range and offset which is not within the missile&#39;s capability to produce a high impact angle. 
     Although FIG. 3 demonstrates the composite LOS rate, further studies indicate that the pitch line of sight rate is similarly indicative of target conditions which should not be biased. 
     Targets that have a PLOS rate greater than −1°/second have sufficient range and altitude to permit the trajectory to be biased to achieve a higher impact angle. Thus, the bias is triggered for these PLOS rates, and switches  14  and  15  are closed. For rates less than −1°/sec, the switches are opened. 
     With the bias signals BIASP and BIASY applied for those PLOS rates in excess of −1°/second, a constant bias signal is applied to the pitch and yaw forward guidance loops by the closure of switches  14  and  15 . Of course, it is possible, within the scope of the present invention, to include non-fixed values of bias selected in response to flight parameters of the missile. 
     The system operates as a smart missile by sensing the need for trajectory biasing at target acquisition and then adding a biasing signal to the pitch and yaw guidance control. The triggered bias is also smart in that during the flight of the missile, if the PLOS rate drops below −1°/sec., the bias is removed, avoiding excess arcing of the trajectory. 
     The foregoing bias signals may also include a signal term for providing a gravity bias signal to the forward loop. Those skilled in missile control systems will recognize that every vehicle will be influenced by gravity differently depending on its trajectory parameters. To offset the effect of gravity on the pitch plane, a signal biasing against the effect of gravity may be added to the pitch forward guidance control signal along with a BIASP signal. This gravity bias rate signal may also be triggered in response to a given PLOS rate. 
     The foregoing control system, whether implemented by computer or hardwired networks, may be implemented with the parameters of Table 1. 
     
       
         
               
             
               
               
               
               
             
               
               
               
               
             
           
               
                 TABLE 1 
               
             
             
               
                   
               
               
                 VEHICULAR AUTOPILOT MODEL 
               
               
                 DATA DESCRIPTION 
               
             
          
           
               
                 Symbol 
                 Description 
                 Units 
                 Default Value 
               
               
                   
               
             
          
           
               
                 K P   
                 Pitch Guidance Gain 
                 Sec 
                 0.8 
               
               
                 K y   
                 Yaw Guidance Gain 
                 Sec 
                 1.6 
               
               
                 K DP   
                 Pitch Rate Feedback Gain 
                 Sec 
                 0.04 
               
               
                 K L   
                 Pitch or Yaw Forward Loop 
               
               
                   
                 Gain 
                 Rad 
                 1.0 
               
               
                 K R   
                 Roll Forward Loop Gain 
                 Rad 
                 1.0 
               
               
                 τ RN   
                 Roll Forward Loop 
                 Sec 
                 1.0 
               
               
                   
                 Numerator Time Constant 
               
               
                 τ RD   
                 Roll Forward Loop 
                 Sec 
                 1.0 
               
               
                   
                 Denominator Time Constant 
               
               
                 G.B. 
                 Gravity Bias 
                 Deg/Sec 
                 −1.0 
               
               
                 K φ   
                 Roll Attitude Gain 
                 Rad/Rad 
                 0.06 
               
               
                 K {dot over (φ)}   
                 Roll Rate Feedback 
                 Rad/Rad/Sec 
                 0.01 
               
               
                   
                 Gain 
               
               
                 K DY   
                 Yaw Rate Feedback Gain 
                 Sec 
                 0.04 
               
               
                 B.V. 
                 Trigger Rate Bias Value 
                 Deg/Sec 
                 3.0 
               
               
                 T.V. 
                 Trigger Rate Bias 
                 Deg/Sec 
                 −1.0 
               
               
                   
                 Threshold 
               
               
                 τ PN   
                 Pitch Command Numera- 
                 Sec 
                 1.0 
               
               
                   
                 tor Time Constant 
               
               
                 τ PD   
                 Pitch Command Denominator 
                 Sec 
                 1.0 
               
               
                   
                 Time Constant 
               
               
                 τ YN   
                 Yaw Command Numerator 
                 Sec 
                 1.0 
               
               
                   
                 Time Constant 
               
               
                 τ YD   
                 Yaw Command Denominator 
                 Sec 
                 1.0 
               
               
                   
                 Time Constant 
               
               
                   
               
             
          
         
       
     
     Referring to FIG. 2, there is shown a block diagram of programming steps which will permit implementation of the rate bias signal generator  10  of FIG.  1 . The pitch line of sight rate, PLOS, from the seeker is detected in step  51  and compared with a triggered value rate, T.V., which is preferably −1.0°/second. If the PLOS rate is greater than −1.0°/second, a bias condition is indicated by step  52 , A=1. If the PLOS rate is less than or equal to −1°/sec, no bias condition is indicated by step  59 , A=0, and no bias is applied to the system. An initial bias value B.V. which is preferably 3.0°/second is selected in step  53 . In step  54 , the roll reference angle determined from on-board instrumentation at target acquisition is detected, and the BIASY and BIASP levels are subsequently determined in steps  55  and  56 . 
     The gravity bias may be added by setting a control flag  57  which will combine a bias value G.B., previously determined from the vehicular characteristics, to the BIASP 0  level determined in step  56 . The resulting bias signals are added to the summing junctions  11 ,  12  of FIG.  1 . 
     Thus, there has been shown a proportional guidance system in terms of classical control system parameters which will bias the missile trajectory under certain target conditions to achieve a high impact angle with a moving or stationary target in the presence of stationary or adverse winds. Those skilled in the art will recognize yet other embodiments defined in terms of the claims which follow.