DE2210561B2 - Guiding flying object to target - flying object linked with moving target using location and control instruments - Google Patents

Abstract

The arrangement comprises at least 1 pitch and 1-jaw range for the return guide signals from the flying object. In the link up arrangement interference parts of flying objects are built up from flight path measurements and are superimposed on the two guide signals and are fed further into the command receiver of the flying body to make corrections for the interference signals. The path of the flying object, and the target is followed by the guide arrangement and the various angles obtained by instruments are sent to command calculator where they are combined with the element and interference signals and sent to a command receiver to actuate the controls.

Description

The invention is a method for steering a missile with at least one pitch and roll a yaw control loop, for the feedback signals from the measured values from one outside the center of gravity the missile arranged transverse accelerometer can be obtained, with the inner control loops for the connection of in a steering system generated steering signals are meshed with corresponding external control loops.

With a missile of this type, it is well known, see German patent specification 14 81 508, for extraction the activation signals for the control loops to provide three accelerometers, which are in a defined Distance in front of the 'center of gravity of the missile provided in the plane the attitude gyro used for roll position control is also located, and its output signals in input values that represent a combination of rotational and lateral acceleration.

The use of only one leveling knife per axis in the missile allows the jiect measurement a signal that is made up of the lateral acceleration and the product of the angular acceleration with the Distance of the E'.nbauortes the accelerometer from the center of gravity of the missile composed. Since the transverse acceleration is selected as the reference variable, the problem arises, this from the Isolate composite signals. It can be gecigne'c by choosing the location of the accelerometer can be achieved that the signal amplitude of the transverse acceleration when turning the missile predominates. Furthermore, it can be assumed that the angular acceleration occurs when the flight is almost straight the sign constantly changes, so that the signal of the angular acceleration is given when the lateral acceleration is small has an effect with the mean value = 0. There is also the feedback signal for missile control is required, however, it is necessary to integrate the aforementioned composite signal in order to derive from the angular acceleration component to gain the angular velocity.

When integrating the measured value to form the feedback signals, however, not only the angular acceleration component, but also the portion of the transverse acceleration is also recorded, so that a mixed signal that cannot be separated is obtained. The lateral acceleration after the iiucgiation results in a disturbing one Quergeschwir.digkeitssignal, which is the feedback signal for the command generation of the rudder deflection falsified.

The invention is based on the problem of this annoying Quergeschwindigkeilsanieile when steering to eliminate missiles of the aforementioned type, in particular by forming a largely trouble-free commands for the rudder deflection.

Based on a steering method of the type mentioned at the outset, this object is according to the invention solved in that disturbing transverse speed components of the missile in the Lenkanlagc Trajectory measurements are simulated, connected to the generated steering signals and sent to a command receiver of the missile are passed on and that the steering signals affected by transverse speed switched on the inner control loop in the sense of eliminating the disruptive transverse speed components will.

According to a further embodiment of the invention, the tracks are in the steering system with a locating device of the missile and an approaching target pursued, wipe the respective WinkdüilTcrcnzen / its two straight lines in a communication computer entered, there with the distance of the missile / u obtained through ongoing time measurements do the replication of the disturbing cross-velocity ratio of the missile containing steering

ignalen processed and the steering signals to a command transmitter for transmission to the Kcmmandoenipiar.gcr of the missile forwarded. According to further development of the invention

Missile the readings of the lateral accelerometer integrated in integrators, m flight controllers the speeds resulting from the integration processed by addition with the measured values to form the feedback signals and the feedback signals with the steering signals coming from the command receiver for the purpose of eliminating the -enden transverse speed component subtracted to provide undisturbed control signals for the actuation of servomotors to generate rudder deflections. .

By eliminating the disruptive parts of the transverse speeds in the return signals are even with constant turning flight, in which the share of the lateral speed is proportional to time increases rapidly, the flight controller always determines the correct rudder deflections. It can therefore the advantages of equipping a lost device missile with only three accelerometers be fully exploited. It is needed in missiles besides those for the formation of the control command provided facilities no further facilities, e.g. B. Filter, to eliminate the interference are provided. The three accelerometers require associated networks and other facilities only a single room outside the center of gravity of the missile, thus the space around the center of gravity, for example for the accommodation of the fuel tank can remain free. The ones caused by the shifts in the center of gravity in the missile - that essentially occur by fuel consumption - errors caused, di; due to the changing distances the mounting planes of the accelerometer from the center of gravity and thus the change in the factors to calculate the pure signal components of the accelerometer are smaller.

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

Fig. 1 is a schematic representation of a steering system for an object approaching a target,

Fi g. 2 is a schematic representation of in the missile arranged accelerometers for performing the method according to the invention. Fig. 3 is a block diagram of the missile control

lung and

4 is a block diagram of the missile facilities and steering system for transverse acceleration control of the missile around its pitch axis.

As FIG. 1 shows, a steering system 1 for command control of a missile 10 which is approaching a target 15 has a locating device OR, a command computer KR and a command transmitter KS . The locating device consists of a double radar device to determine the angle f enclosed by the degrees of connection to the missile 10 and to the target 15. The Konimando computer is used to generate the information to be described below, and the command transmitter transmits steering signals to the missile.

The one in F i μ. 2 iiahcr shown missile has Cartesian coordinate axes -Y going through the center of gravity S as the zero point. Y and / on. The arrows each point in the positive direction. In the K-direction are arranged Stabilisafioniiflächen-la and 26, and tail Za and 36 P.uderP.ächen 5a and 56, and in the Z direction with Aa tail control surfaces 6a. Opposite the latter is a tail unit 46 (not shown) with a rudder 66 offset by 180 degrees. The deflections of the oars 5 a and 56 are marked with η, those of the oars 6 a and 6 6 with ξ . At a distance from a to the center of gravity V, in alignment with the tip 7 of the missile 1 in a transverse plane 3, there are three accelerometers M ₁ , M, and Μ. _Λ arranged. The accelerometer 'Mj, which is designed as an angular accelerometer i >> i lüiü largely corresponds to a rate gyro in structure, measures the acceleration of the missile 1 around the roll axis X. The accelerometers M., and M _{; j} , each at a distance b from the Roll axis X are arranged, 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 other devices to be explained in connection with FIG. 3 for steering the missile are installed.

The control commands are transmitted from the command transmitter KS of the floor position shown in FIG. 1 to a command receiver KE located in the missile (cf. FIG. 3). The MelJwerte the accelerometer Λί., And Λί., Are integrated in integrators IY for the pitch axis and the yaw axis for IZ and transmitted together with the command received by the receiver KE Steuerkommandcs to flight controllers RY and RZ. In the latter there is a conversion into signals for the servomotors SY and SZ, of which the servomotors SF cause deflections) j of the rudders 5 α and 5 6 and the servomotors SZ cause deflections ί of the rudders 6 a and 66. For the sake of completeness, it should be mentioned that the formation of the signals for the rolling movement takes place exclusively in the missile 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 the steering control of the missile I is shown in the yaw plane, the devices in the missile 10 and those in the steering system are each summarized by a large block shown with dashed lines. The command computer KR _y, 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 - / (/) is always available. The Ortungscinrichtunf OR continuously measures the angular deviation t occurring m in the y-plane. between the straight lines connecting the missile and the target 15 and forwards these values to the Korninando calculator KR. There the latter forms the product of the respective transverse deviation with the respective distance s of the missile. I ,.

so .I ₁ . -

The transverse deviation A _y , transformed into a sleuerkommando, must be connected to the disturbance component of the transverse velocity which results from the integration of the measured values of the accelerometer Ml (M 3) in the missile.

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

W _y = / (4th) (2)

forms in the following way.

To understand the simulation of the interference component of the transverse speed, first the formation of this component from the measured values z. B. the accelerometer Ml described, which is known to be the measured value

ml - y -f- 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

Focus 5,
toz is the angular acceleration around the z-axis.

The measured values em2 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 _ζ. The component of the transverse acceleration y is of course also recorded, as a result of which the feedback signal Ry is obtained as a mixed signal of transverse velocity y + angular velocity w 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 2. 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:

multiplying constants Cl and C2, so that: the two portions of the Riickführsignals with one of the individual members can of the feedback signal to be signal for the rudder angle η in the right size to the servomotors SY forwarded After multiplication Ry with the constants C 1 and Cl yields fell the corrected feedback signal

= C \ ■ y + a- Cl-w _z + C2 ■ y - \ - a-C2-

Since the transverse acceleration y is the reference variable, it must be isolated from the mixture of signal values. This is done by simulating the constant C2 · y on the ground and adding it to the steering command generated there with the opposite sign to the return signal of the missile.

The computer in the command KR the ideal control command for the Quergeschwindigkiet W _3, to add noise component of the transverse velocity C2 ■ \ is formed by a time derivative formed from the lateral deviation Ay and multiplied by C 2, whereby the disturbing member

German

arises.

The corrected control command is created by adding to W _y

W _x =

+ J (y + a ■ i _{) z} ) df, (4)

from what

This command is forwarded to the command transmitter KS and from there to the command receiver KE of the missile, which forwards it to the controller RY. In the controller, by subtracting Wy-Ή ^, the interference component C 2 ·,. out

separated and from the remaining acceleration and speed components an undisturbed control signal for the servomotors SY is formed, which controls the rudders 5a and 5 ft up to the required deflection?; adjust.

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

Ry - y + a ■ M _z - \ - y + a · ω _ζ

(5) m3 = ζ + a

<Oy

results.

The accelerometer Ml supplies the measuring

It is necessary to take into account werf ml = ώζ, from which the signal for the servomotors 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)

Patent claims: 1. A method for steering a missile with at least one pitch and yaw control loop, for which feedback signals are obtained from the measured values of one lateral accelerometer each arranged outside the center of gravity of the missile, the inner control loops for switching on steering signals generated in a steering system with corresponding outer ones Control loops are meshed, characterized in that interfering transverse speed components (C2-J, Cl-z) of the missile (10) are simulated from flight path measurements in the steering system (1), connected to the generated steering signals (W _y , W _z ) and sent to a command receiver ( KE) of the missile are passed on, and that the steering signals (W _y , W _z ) affected by transverse velocity are looked up to the inner control loops in the sense of eliminating the disruptive transverse velocity components (C2-y, C2-z) . 2. The method according to claim 1, characterized in that the trajectories of the missile (10) and a target to be approached (15) followed in the Lenkanlagc with a locating device (OR) _{, the respective Winkeidiffercnzen (f v} ) between their two lines of sight in a command computer (KR) entered, there with the distance (s) of the missile obtained through ongoing time measurements to the steering signals (TT _1. , W _z ) containing the simulation of the disruptive Quergeschwindigkcitsteile (C 2 ·> ·, C 2 · z) of the missile and the steering signals are passed on to a command transmitter (KS) for transmission to the command receiver (KE) of the missile. 3. The method according to claims 1 and 2, characterized in that in the missile (10) the measured values (m 2, m 3) of the transverse accelerometer (M 2, M 3) integrated in integrators (/ 2, / 3), in flight controllers (RY, RZ), the velocities resulting from the integration (1) + a · W ₂ , i + α- W _y ) processed by addition with the measured values to the feedback signals (7? Y, 7?:.) And the Feedback signals from the steering signals (Wy, W, ) coming from the command receiver (KE) are combined for the purpose of eliminating the disruptive Qucrgeschwindigkcitsanieilc (Cl- i'i, C2-z) in order to generate undisturbed control signals (»7, J) for the Actuation of servomotors (SY, SZ) to generate rudder deflections (j /, ί).