DE2210561C3 - Method of guiding a missile - Google Patents

Description

will.

60 According to a further development of the invention, in the steering system with a locating device, the path

The invention relates to a method for steering the missile and an approaching target of a missile with at least one pitch and tracked the respective angular differences between a yaw control loop for the feedback signals from their two straight lines of sight into a command line Measured values entered by one outside the heavy 65, there with the current time measuring point the missile arranged transverse thrusts to the distance of the missile obtained inclinometer can be obtained, the inner ones being the simulation of the disruptive lateral velocity control loops for the connection of steering components contained in a steering component of the missile

Signals processed and the steering signals to a command transmitter for transmission to the command receiver of the missile forwarded.

According to a further development of the invention, the measured values of the (accelerometer integrated in integrators, in flight controllers the speeds resulting from the integration by adding the measured values to the Processed feedback signals and the feedback * signals with the Lenkisgnaien coming from the command receiver for the purpose of eliminating the disruptive ones Transverse speed components are subtracted to provide undisturbed control signals for the actuation of servomotors to generate rudder deflections.

By eliminating intrusive shares of the cross-rates in the feedback signals are even with constant turning flight, in which the proportion of Quergeschwindig '-. Ince time disproportionately increases, always determined by the flight controller so right rudder deflections. Thus, the advantages of equipping a missile device with only three accelerometers can be fully utilized. In missiles, apart from the facilities provided for the formation of the control command, no further systems, e.g. B. Filters, can be provided to eliminate the sturgeon infuses. The three accelerometers, with the associated networks and other facilities, only require a single space outside the center of gravity of the missile, which means that the space around the center of gravity can remain free, for example for accommodating the fuel tank. The errors caused by the shift in the center of gravity in the missile - which occur mainly due to the fuel consumption - which occur due to the changing distances between the installation levels of the accelerometer and the center of gravity and thus the change in the factors for calculating the pure signal components of the accelerometer, become smaller.

The invention is described below using an exemplary embodiment shown in the drawing explained in more detail. It shows

F i g. I a schematic representation of a steering system for a missile approaching a target,

F i g. 2 shows a schematic representation of accelerometers arranged in the missile Carrying out the method according to the invention,

3 shows a block diagram of the missile control and

4 is a block diagram of the missile device t'ingen and steering system for lateral acceleration control of the missile around its pitch axis.

Like F i g. 1 shows a steering system 1 for command control of a missile tO which is approaching a target 15, a locating device OR, a command computer KR and a command transmitter KS . The locating device consists of a Doppler radar device for determining the angle f enclosed by the straight line connecting the missile 10 and the target 15. The command computer is used to generate the steering signal to be described, and link signals are transmitted to the missile via the command transmitter.

The in F i g. 2 has Cartesian coordinate axes A ', Y and Z going through the center of gravity S as the zero point. The arrows each generally show a positive direction. Stabilization surfaces la and Ib and tail units 3 a and 3 b with rudder surfaces Sa and Sb and in Z-direction tail units 4 a with rudder surfaces 6a are arranged in the ^ direction. The latter is a not shown tail unit 4 b with a rudder 6 b offset by 180 degrees opposite. The deflections of the rudders 5 a and Sb are marked with »;, those of the rudders 6 a and 6 b with ξ . At a distance from a to the center of gravity S _{, three accelerometers M 1} , M, and M ₃ are arranged in the direction of the tip 7 of the missile 1 in a transverse plane 8. The accelerometer Mj, which is designed as an angular accelerometer and largely corresponds in structure to a rate gyro, measures the acceleration of the missile 1 around the roll axis X. The accelerometers M, and M ₃ , which are each arranged at a distance b from the roll axis X , measure the Angular and transverse accelerations around the pitch axis y and the yaw axis Z. Furthermore, a box 9 is shown in which the others in connection with FIG. 3 to be explained in more detail devices for the steering of the missile are installed.

From the command transmitter KS of the in F i g. 1, the control commands are transmitted to a command receiver KE located in the missile (see FIG. 3). The measured values of the accelerometers M ₂ and M ₃ are integrated in integrators IY for the pitch axis and IZ for the yaw axis and forwarded to flight controllers RY and RZ together with the control commands received by the command receiver KE . The latter is converted into signals for the servomotors SY and SZ, of which the servomotors SY cause deflections ?; the rudders 5 α and 5 b and the servomotors SZ cause deflections ξ of the rudders 6a and 6ft. For the sake of completeness, it should be mentioned that the formation of the signals for the rolling movement without influencing external control commands The measured values of the rate gyro M ₁ are evaluated directly in the flight controller RZ for the yaw axis Z and converted into control signals for the servomotors SZ .

In FIG. 4 shows the steering control of the missile 1 in the yaw plane, the devices in the missile 10 and those in the steering system 1 being combined in each case by a large block shown with dashed lines. The command computer KR, the functions of which are outlined by a dashed line, is connected to a launching system, not shown, from which an impulse is sent to the command computer when the missile is launched. This triggers a predetermined program with which the respective distance s is obtained from the speed of the missile and the time elapsed after leaving the launching system, so that s = / (i) is always available. The locating device OR continuously measures the angular deviation s _y occurring in the y-plane between the straight lines connecting the missile 10 and the target 15 and forwards these values to the command computer KR. There the latter forms the product of the respective transverse deviation A _y with the respective distance s of the missile

so A ₃₁ -

The transverse deviation Λ _v transformed into a control command must be added to the interference component of the transverse speed which results from integrating the measured values of the accelerometer Ml (M 3) in the missile.

The replication of the disruptive transverse speed components of the accelerometers Ml and M 3 of the missile 10 takes place in the steering system 1, the command computer KR of which the "ideal control command" z. B. for the angular velocity around the y-axis

multiply the two components of the feedback signal by a constant C 1 and C 2 so that the signal for the rudder deflection η can be passed on to the servomotors SY in the correct size. After multiplying the individual elements of the feedback signal Ry by the constants C 1 and Cl , the corrected feedback signal is obtained

_H ,

7fy = C 1 ■ y - \ a ■ Cl · ω _ζ 4 Cl · ν + a ■ Cl

^{v y}

forms in the following way.

In order to understand the simulation of the disturbance component of the lateral speed, the formation of this is first Share from the measured values z. B. of the accelerometer Ai 2 described, which is known to the Measured value

_m 2 - ·· 1 _a

(3)

supplies. Mean therein

y is the lateral acceleration in the y direction,
α the lever arm of the measured value level 8 to

Heavy jump S
ωζ is the angular acceleration around the ζ axis.

The measured values m 2 are integrated in the integrator IY to obtain a feedback signal Ry required for the regulation, in order to obtain the angular velocity ωζ from the angular acceleration ω _2. Of course, the component of the transverse acceleration y is also recorded, as a result of which the feedback signal Ry is obtained as a mixed signal of transverse velocity keily + angular velocity ω times the distance from the center of gravity α .

In order to obtain the angular velocity alone, the angular acceleration ώ _ζ must therefore make up an evaluable portion of the original measured value m 1. The feedback signal Ry is obtained by adding the measured value m 2 and its integral obtained from the integrator IY in the controller RY according to the following equation:

Since the transverse acceleration y is the reference variable, it must be isolated from the mixture of signal values. This takes place in that the constant Cl · _{y is} simulated on the ground and added to the steering command generated there with the opposite sign to the return signal of the missile.

^ao ^ ^{cr 'm} command calculator KR the ideal control command for the Quergeschwindigkiet W _y to add noise component of the transverse velocity Cl ■ j> ^We d formed in that formed from the lateral deviation ^Λ> * f ^me _v derivative with respect to time and with C 2

^{a5 mu} 't'P "is ^adorned , making the Storghed

= egg

arises.

The corrected control command is created by adding to W _y

TP = W + C

^yy "

= y + ο ωζ + J (y + a ώ _ζ ) dt, (4)

from what

This command is forwarded to the command transmitter KS and from there it is transmitted to the command receiver KE of the missile, which forwards it to the controller RY. In the controller, the interference component Cl- _{y is} eliminated by forming the difference between W _{y ~ Έ y} and an undisturbed control signal for the servomotors SY is formed from the remaining acceleration and speed components, which adjust the rudders 5a and Sb up to the required deflection.

Corresponding to the above-described steering of the missile in the yaw plane, steering in the pitch plane takes place in the same way, with the accelerometer M 3 delivering the following measured value:

= y + α · ω _ζ + y -f a ■ ω _ζ

(5)

m3 = ζ + α ω _ν

results. The accelerometer Ml supplies the measuring

It is necessary to take into account the value ml = ό «, from which the signal for the servomotor electrical values of the networks and devices 55 SZ is formed without influencing the installation location of the accelerometer and the floor in the controller RZ.

For this purpose 2 sheets of drawings

Claims (3)

'S system generated steering signals with corresponding claims: external control loops are meshed. In a missile of this type it is known ! Getting lost to steer a missile vgi. German patent specification 14 81. 508, for extraction with at least one pitch and yaw 5 of the switching signals for the control loops three control loops, to provide for the feedback signal from the measuring scMcunigungsmessers in an im of a distance outside the center of gravity in front of the center of gravity of the flight of the missile arranged transverse propellants are arranged in the plane provided inclinometer can be obtained, with the bearings also serving to regulate the roll position Control loops for the connection of an io gyro located, and their output signals in the on-steering system generated steering signals are converted with change variables that speak a combination represent outer control loops meshed from the rotational and lateral acceleration, are characterized in that in The use of only one accelerometer the steering system (1) disruptive transverse velocity per axis in the missile allows the direct measurement of the missile (C 2-y, Cl-z) of a signal that is reproduced from the lateral acceleration from trajectory measurements, the he and the The product of the angular acceleration and the generated steering signals (W _y , W ₂ ) are added and the distance between the installation location of the accelerometer and a command receiver (KE) of the missile is composed of the center of gravity of the missile, and that the transverse acceleration is the reference variable speed-affected steering signals (W _y , W ₁ ) 20 is selected, the problem arises of isolating them from the inner control loops in the sense of an Elim'i signal mixture. It can be switched on by suitable modification of the disruptive lateral speed selection of the installation location of the accelerometer components (C2-'i /, C2-i). be achieved that when turning the missile 2. The method according to claim 1, characterized GE pers the signal amplitude of the transverse acceleration indicates that in the steering system with a 25 predominates. Furthermore, it can be assumed that in the case of a locating device (OR) the orbits of the missile almost straight-line flight, the angular acceleration (10) and a target (15) to be approached constantly changes the sign that if it is smaller, the respective angular differences (ε _ν ) between - Lateral acceleration the signal of the angular sluice of their two lines of sight in a grain inclination has an effect with the mean value = 0. Since input to the mandate computer (KR) , there with the missile control also the feedback signal obtained through ongoing time measurements is needed, it is necessary to integrate the mentioned distance (5) of the missile to the post-signal mixture in order to get out the angular formation of the disruptive transverse velocity acceleration component contains the angular velocity components (C2 ■ j ·, Cl ■ z) of the missile. tend steering signals (W _y , W _z ) are further processed 35 When integrating the measured value to form the and the steering signals to a command transmitter, feedback signals are not only the angle (KS) for transmission to the command acceleration portion, but also the portion of the receiver (KE) of the missile forwarded transverse acceleration is also recorded, with which a not to be. separating mixed signal is obtained. The transverse 3. The method according to claims 1 and 2, 40 acceleration results after the integration of a stod characterized in that in the missile (10) rendes transverse speed signal, which the measured values (m 2, m 3) of the transverse acceleration feedback signal for the command formation of the rudder knife (M 2, M 3) in integrators (/ 2, / 3) of interest falsified. grated, in flight controllers (RY, RZ), which result from the invention is based on the object, this the integration resulting speeds 45 disturbing transverse speed components during steering (y + a- W ₁ , i + a ■ W _y ) by addition with the measured values of missiles of the aforementioned type to eliminate the feedback signals (Ky, R ~ z) ren, in particular largely processed by forming a and the feedback signals from the interference-free commands for the rudder deflection. Command receiver (KE) coming steering starting from a steering method of the initially Signals (W _y , W _z ) for the purpose of the Eliminie- 5 ° mentioned type, this object is achieved according to the invention tion of the disruptive transverse speed components in that disruptive (C2-y, C2-z) are brought together in the steering system to un- Transverse speed components of the missile from disturbed control signals (η, J) simulated for the actuation of flight path measurements, the generated by servomotors (SY, SZ) connected to generate steering signals and to gain command rudder deflections (η, ξ). 55 recipients of the missile are forwarded and that the steering signals subject to transverse speed control the inner control loop in the sense of eliminating the interfering transverse speed components switched on