DE3802551C2 - - Google Patents

Description

The invention relates to a method for correcting the trajectory an arrow-stabilized missile, according to the preamble of Claim 1, as well as a missile for performing this method, according to the preambles of claims 8 to 10.

Such a method and such a missile are from DE-OS 30 24 842 known. This known missile points in the area of its Tip, i.e. with a longitudinal distance to the center of gravity of the missile, a device for applying a correction pulse on the runs perpendicular to the direction of flight. The facility for Application of the correction pulse is controlled by a steering electronics a control moment around the focus on the missile to apply. It is assumed that the steering electronics has a device for determining the direction of flight.

Because such a missile is usually relatively long and is slim, it has a pronounced natural vibration behavior on, with a so-called pendulum frequency, which is essentially by the moment of inertia, a form factor (air torque coefficient derivative) and the air density is determined.

This is therefore the case with the correction impulses of the known missile Natural vibration behavior to take into account that the excitation of the missile to natural vibrations under all circumstances is avoided, as this the steering behavior, if possible only by the force of the correction impulses on the position of the stretched missile should be determined by interference overlap, which deflect the missile from its direction of flight and can rise so that in the borderline case there is even a loss of Flight stability can come.  

But if there is a particularly large change in angle due to a correction pulse is to be achieved, there is always the risk of Natural vibration excitation, so that in practice such angle changes cannot perform.

Based on this prior art, the object of the invention is based on the known method mentioned at the beginning and the Develop device for its implementation in such a way that large changes in the angle of the flight path can be made quickly, and that in addition no elaborate devices for determining the direction of flight are required.

This object is achieved by the method features of claim 1 and as to the implementation of these procedural features through the Missile of claims 8 to 10 solved.

Here, the excitation of a pendulum vibration is not possible avoided, on the contrary targeted and with considerable Amplitude provokes that the correction pulse is applied according to the time of release, duration and intensity, that it causes the missile to oscillate immediately.

By a suitable time on the course of the correction pulse matched second correction pulse can end of the control process the natural vibration is completely canceled will.

In this case, a is preferred for applying the correction pulse Thrust vector control used, which is preferably on such Place of the missile is arranged to a large pendulum vibration to create.

It is fundamentally possible, the opposite, possible second correction pulse at another point on the missile to apply this as to that on which the first  Correction pulse is applied. But is preferred same thrust vector control for excitation and damping the pendulum vibration used. The duration of the process is preferred half a period of the natural vibration of the Match missile.

So it is through the correction impulse a moment that creates the launch of the missile against its flight direction, as well as a counter moment, that for the subsequent neutralization of the initiated Vibration serves. The missile turns thus briefly in the desired direction.

The missile is here after the correction in a similar flight condition as before the correction, whereby the longitudinal axis of the missile, which corresponds to the original Direction of flight matched, now with the new direction of flight matches. Only the direction of flight is changed.

The pendulum frequency is either constant or changes over the course of the trajectory, but can for each missile is assumed to be known and does not need to be measured in flight. With With the help of vibration theory, it is easily possible Define control torque curves and thus correction impulses which, if adapted to this pendulum frequency, at any point in the trajectory both vibrations initiate as well as extinguish.

The method according to the invention has in addition to that mentioned above Advantage that the direction of flight is at least outside the correction processes always with the missile longitudinal axis agrees, also the advantage that the employment  of the missile in relation to its flight direction by excitation its natural frequency and thus the fastest possible Way for starting and resetting is done. From In principle, the minimum duration of a flight direction correction is therefore half a period and corresponds to the Excitation of only half a vibration length (pendulum length).

The pendulum length of common missiles and projectiles is between 300 and 1000 flight calibers and is always small compared to the operating ranges.

Because of the achievable by the inventive method short correction distances, e.g. between 1 and 4 pendulum lengths, it is according to a preferred one Embodiment of the invention enables a standardized correction pulse repeated on the missile to act to solve a steering task. This kind of repeated, standardized and quick Correction is only possible because after carrying out any correction the natural vibration of the missile has been completely canceled.

In this way, the guidance of a missile becomes special relieved, because it does not have a defined steering force each specific, adapted size to the missile be exercised, but by triggering one several times The standard correction pulse is the deflection of the missile operated long until its new direction of flight with the desired matches.

It is possible and according to a preferred embodiment the invention may also be advantageous, not only  one correction pulse each to initiate and cancel the Apply natural vibration to the missile, but several such correction pulses. So it is e.g. possible, by applying two correction pulses of the same type a vibration behavior every quarter of a period to achieve, at which the two excited Add vibrations so that a permanent job entry. By two each also for a quarter period it is offset like correction impulses then possible to cancel this form of natural vibration again. This way you can get bigger in no time Angle of attack can be reached.

In the case of such correction pulses, a predetermined one can thus Angle of attack can be set without overshoot, whereby the procedure an adjustment to the setpoint within a flight distance of only one pendulum length. With a sudden onset of a control torque namely not only causes a quick queue, but at the same time a vibration is excited so that the angle of attack can temporarily take on values that can be twice as large as the setpoint to be set. It is known that at large angles of attack aerodynamic Forces can occur that are transverse to the direction of correction lie and thus distort the steering. There according to the invention, however, the overshoot is avoided, the inventive method allows implementation more precise corrections than all previously known methods.

As a rule, missiles of the type mentioned initially carry a rotation around its longitudinal axis, a so-called rolling movement, by. It is therefore possible that by application  a thrust vector at only one point on the missile this depending on the rollage just taken can be deflected in any direction. Here is according to a preferred embodiment of the invention however, take into account that between triggering the natural vibration and the suppression of the natural vibration the missile makes a turn which is less than 180 °. The occurring aerodynamic Forces caused by an intended lateral deflection also act up and down, compensate each other, so that there is only one resultant, the causes the desired distraction of the projectile.

Because after eliminating all contributing to the correction Vibrations the new flight direction and the missile Longitudinal axis match again, there is no angle of attack more so that the flight with minimal drag he follows.

Should the flight conditions during the mission be too strong vary from each other, these flight conditions are determined by Applying additional forces, e.g. by means of a March engine, by extending wings, etc., if possible kept constant.

As mentioned above, it is preferred that the duration a correction with half the period of the natural vibration of the missile matches.

In this case, a and the same device for applying a thrust vector or another control impulse for initiation and termination the natural vibration can be used. Basically  but it would also be possible to have several, at different Places to use devices of this type and choose a correction period that differs from one half period of the natural vibration differs.

It is basically possible that the correction pulse attacks at the level of the center of gravity; um with correction forces as low as possible, however, one as possible To achieve a great moment shows the invention Missile a point of attack for the correction pulse, who is far from the center of gravity and prefers is in front of the focus.

Conventionally used controls (symmetrical Setting duck wings or oars, thrust deflection or the like) can be used; prefers but the missile according to the invention has a gas generator or a reaction engine, the thrust can be controlled quickly and easily, its thrust effect as far as possible in front of the center of gravity of the missile is used to create a big moment, and whose thrust is directed essentially radially, wherein ensures that thrust and aerodynamic Corrective force add up.

According to a preferred embodiment of the invention This makes the missile's controllability significant improved when this special flight characteristics having. It is known that with trajectory corrections, which are caused by the missile being deployed, missiles with great buoyancy. In which  that has a dynamic process of switching on and off Air tipping moment has a big impact. It works with the Employment against the applied control torque and ensures for the default. So it should be small during the Start-up and large during the reset process be. The known linear dependence of the tilting moment from the angle of attack fulfills this condition. The production however, non-linearity means according to a preferred one Embodiment of the invention a further improvement, the tilting moment is small in the direction of attack. With slow rolling, asymmetrical missiles it is possible, a dependence of the tipping moment on the rollage to create. Now the roll frequency is the hiring process adjusted, then the rollage during the employment be chosen so that the tilting moment is small remains. It is thus achieved that with small control torques large angles of attack can be generated.

According to a further preferred embodiment of the invention the missile according to the invention can be deployed Have wings, a march engine or the like, to the flight conditions of the missile during its To keep the flight duration as constant as possible.

To determine the direction of flight is only determining the position of the longitudinal axis of the missile required.

Here is a step-by-step determination of the amount the angular placement of the missile from the target a simple one Adjustment of the accuracy of the deposit measurement to the correction options  given. E.g. with a uniform Correction pulse or standard correction pulse the trajectory rotated by 1 °, the angle rest only has to be at most 0.5 ° accuracy can be determined to ensure that a correction will only be made if it necessary is. In addition, the angular shelf must be to the target not be measured continuously; it just has to be done if there is no correction, i.e. If you is easy to carry out because the flight direction and missile Longitudinal axis match. Furthermore, the facility only a small angular range for shelf measurement capture because during the correction, at the big one Angular displacements can occur, does not have to be measured.

In the preferred use of the missile according to the invention, if this with a device for generating equipped with a standard correction pulse is when steering with a beacon simple determination of the direction of the projectile flight compared to Guide beam possible. After the correction is correct the longitudinal axis of the missile and its direction of flight again, and the deposit angle is always small versus the correction angle.

In the preferred use of the missile according to the invention is an immediate determination of the storage angle possible.

A determination of the relative angular velocity between Target and missile is also without additional Setup feasible. In the first approximation, the angular offset to the target after correction from the Angle storage before the correction by the known angle  the orbit rotation can be reduced by the correction must, and from the angular storage, by the target movement caused during the correction together. From this can be known by difference between two Sizes a mean relative angular velocity of the Target during the correction period.

The object of the invention is based on the schematic drawing, for example, explained in more detail. In this shows:

Fig. 1, applied in the inventive method for vibration principle release,

Fig. 2, the applied principle in the inventive method for the trouble-free set-point adjustment,

Fig. 3, the setting of a residual vibration correction in free spatial movement of a missile,

Fig. 4 shows the use of a missile according to the invention in proportional navigation, and

Fig. 5 shows the use of the missile according to the invention in a beacon guidance.

In order to present the basics underlying the method and missile in more detail, a transition is first made to FIG. 1.

In Fig. 1, two diagrams are superposed shown, where applied at the top, the force acting on a missile during a control operation in the lower torque and the angle of incidence produced by this moment, in each case over time.

A correction pulse 1 forms a jump function in the upper diagram, which acts on the missile, which then responds in its natural frequency (lower diagram). For example, in FIG. 1 the answer to a step function 1 is a cosine oscillation 3 . The direction of the response depends on the direction of the suggestion.

If two opposing, abrupt, jump-like correction pulses 1 and 2 are applied to the rocket capable of oscillation, offset by time T, a single pulse of duration T is obtained by superimposing the two answers. After the action of step function 2 , the two excited vibrations neutralize 3 and 4 .

Here, the response oscillation 4 (shown in dashed lines) is offset by half a period T with respect to the cosine oscillation 3 .

The missile's answer 5 is drawn in bold.

In the embodiment of FIG. 1, a missile is shown which receives a standard correction pulse 1 or 2 , and this in succession until the desired trajectory correction has been carried out.

The embodiment of FIG. 2, on the other hand, in which the deflection torque over time is shown in the upper part of the figure and the pendulum oscillation over time in the lower part, shows the setting of a desired deflection with a predetermined size:

Two abrupt correction pulses 6 and 7 are offset from each other by half a period T / 2, so that there is a cosine oscillation as a response oscillation on each of the excitations, one of which is drawn out and the second is shown in dashed lines. Since the amplitudes of the two cosine vibrations add up at each point, a resulting angle of attack 10 results, which is set in the shortest possible time and is maintained at least for a certain period of time without any overshoot. Due to the correcting pulses 6 and 7 , respectively offset pulses, the offset value, the constant value 10 can be completely canceled again without any post-oscillation.

In Fig. 3, a missile 11 is shown, with a focus 12 and a longitudinal axis 16 , which carries out a rolling movement drawn by a ring-shaped arrow. A gas generator 15 is arranged in the area of the tip of the flight body 11 laterally and well in front of its center of gravity 12 so that its drive beam emerges radially to the flight direction 16 in the direction of the arrow, so that an opposing impulse is exerted on the projectile 11 .

This also has rear wing 14 and duck wing 13 , which can also be adjusted in place of the gas generator 15 to generate a pulse that excites the natural vibration of the projectile.

The current flight direction after the correction has been made is denoted by 17 ; the flight direction 17 and the longitudinal center axis 16 of the missile 11 together form a correction plane 18th

The actual movement of a point of the missile 11 and thus the angle of attack is shown as a space spiral 20 . For a better overview, the projection 21 of this angle 20 is shown on a plane 19 which is perpendicular to the current flight direction 17 .

The duration of the correction pulse or the interval between two gas surges from the gas generator 15 corresponds to the duration of the excited oscillation. An adaptation to the natural frequency of the missile is thus achieved. The natural frequency of the missile, however, does not have to remain constant during the correction.

The course of the amount of the tax moment is thus shown specifies that the required amount of the angle of attack he is enough and that the angle of attack at the end of the mission of the moment assumes and maintains the value zero. It will thus the effect of the pendulum through the applied moment damping and the rotation of the flight direction during the Correction neutralized (with strong pendulum damping can the impulse should be significantly lower than the impulse that initiates the vibration).

The direction of the correction pulse, which initiates the control torque, must lie in the first approximation in the correction plane 18 , the rolling movement, the correction duration and thus also the natural frequency of the missile 11 having to be adapted. To ensure that the residual vibrations are extinguished, the directions of the correction pulse, ie the roll position, and the angle of attack, ie the pendulum position, must match at the end of the correction. This method also compensates for the moments which are generated by pulses which lie transversely to the correction plane 18 , which is spanned by the flight direction 17 and the projectile axis 16 .

By adapting the missile properties to the gefor conditions can be the size of the required Tax torque can be reduced. With aerodynamically asymmetrical tric missiles, the control torque is preferred in Direction of the lowest air tipping moment. The roll speed is adjusted to the correction so that in the first approximation the direction of the angle of attack always matches the rollage. The control torque acts thus always against the slightest tilting moment.

The maximum permissible roll frequency depends on the duration of the Correction determined. To ensure that the corrections always a component that is in the specified Rich device shows, possesses, may rotate the missile do not exceed 180 ° during the correction.

The correction pulse is used in a roll position, which depends on the required correction direction and the rotation of the missile 11 during a correction.

The design is preferably carried out with a trajectory rake model in six degrees of freedom, which contains the characteristics of the missile 11 to be controlled.

The flight characteristics of the missile remain during the Trajectory is approximately constant, for the solution of the  Steering task a uniform correction pulse constant Duration defined, which corresponds to the storage of the missile pers to the goal is initiated repeatedly.

The flight conditions vary during use strong, known means such as marching engines, Aus drive from wings etc. used to their varia tion within limits.

In FIG. 4 the situation is shown of a missile between the two corrections. Here, the missile is initially (in Fig. 4 bottom left) in an attitude in which the flight direction 22 and the missile longitudinal axis 23 coincide. The angle γ _{1 important} for a proportional navigation between the current target position 24 and the flight direction 22 can therefore be measured directly in the missile using known optronic means. If this angle exceeds a predetermined value, for example 1 °, if the trajectory deflection by a standard control pulse is 2 ° for example, then the roll position corresponding to the storage direction must be waited for before the standard correction pulse is given to the missile. During the correction, an angle of attack α (angle between the direction of flight and the longitudinal axis of the missile) is generated, which can be over 10 °. The correction pulse is carried out in accordance with the program regardless of the current position of the target and the missile. Only after the correction has been completed does the flight direction 27 then complied with the missile longitudinal axis 23 again, and the process can be repeated with the new current target position 29 .

From the measurement of the target placement angle (γ ₁ and γ ₂ ) before and after the correction, the relative target angle speed, which is necessary for the determination of a lead angle, is determined. The relative angular calculation of the target during a correction results from:

Δγ _Z = γ ₂ - (γ ₁ - γ _k ) × (1 + Xk / X)

Xk is the distance (e.g. two pendulum lengths) for one Correction is needed, and X is the distance to the target. As long as the missile did not come too close to the target (e.g. a distance of ten pendulum lengths or more the knowledge of the distance to the target is not required. With the help of constant time for one Correction is required, then the relative win target speed determined.

In FIG. 5, the Leitstrahlverfahren is shown in which the angle α _l between a guide beam 30 and the connecting line 31 between the launching site 32 and the missile 33 in principle is small and is of the order of a milliradian. If the angle α _F between the illumination direction 31 and the projectile axis 34 is measured in the missile between two corrections using known optronic means, this angle can be equated to the angle between the flight direction 35 and the guide beam 30 in a first approximation. While this size is only determined in known methods by filtering and derivation of the placement angle, these operations are no longer required in the illustrated embodiment, for example, since after making the correction the flight direction 35 coincides with the longitudinal axis 34 of the missile 33 .

Claims (10)

1. Method for correcting the trajectory of an arrow-stabilized, preferably fast-flying missile which is capable of a transverse natural vibration with a predetermined fixed or variable frequency, at least one correction pulse being applied to the missile, which generates a moment around its center of gravity,
characterized in that the moment is so matched to the natural vibration that

• - This is first so strongly excited that the missile is turned against the direction of flight, and
• - The correction pulse is matched to the natural vibration so that it is completely canceled after excitation and after completion of its action.

2. The method according to claim 1, characterized in that at least a second, counter-directed correction impulse on the Missile applied in such coordination on its natural vibration will be lifted completely. 3. The method according to any one of claims 1 or 2, characterized in that that the correction process is repeated several times. 4. The method according to any one of claims 1 to 3, characterized in that successively applied several correction forces be that a given bend in the trajectory of the missile or a predetermined angle of attack is reached. 5. The method according to any one of claims 1 to 4, wherein the missile performs a rolling movement, characterized in that the Duration of the natural vibration forced and ended by the correction pulses is shorter than the period within which the missile performs a rolling movement of 180 °. 6. The method according to any one of claims 1 to 5, characterized in that that the flight characteristics of the missile by application  additional forces, such as by means of a marching engine, the Extending wings or the like, kept as constant as possible will. 7. The method according to any one of claims 1 to 6, characterized in that that the duration of the forced and ended natural vibration matches half the period. 8. missile for performing the method according to at least one of claims 1 to 7, with a device for applying a transverse to the direction of the correction pulse, the point of attack is in front of the center of gravity, wherein the missile is preferably set up to perform a rolling movement, and wherein the device for applying the correction pulse is preferably designed as a reaction engine for applying a radial force with respect to the missile longitudinal axis, characterized in that the missile is aerodynamically asymmetrical and that the device ( 15 ) for applying the correction pulse is arranged so that it is in Direction of the smallest air tipping moment. 9. Missile for performing the method according to at least one of claims 1 to 7, with a device for applying a correction pulse directed transversely to the flight direction, its point of attack is in front of the center of gravity, with the missile preferred is set up to carry out a rolling movement, and wherein the device for applying the correction pulse is preferred as a reaction engine for applying a radial force in relation is formed on the missile longitudinal axis, characterized by a marching engine or by extending wings to reduce the variation in flight conditions. 10. Missile for performing the method according to at least one of claims 1 to 7, with a device for applying a correction pulse directed transversely to the direction of flight, the point of attack is in front of the center of gravity, and with a device for determining the direction of flight, the missile preferably for carrying out a rolling movement is set up, and the device for applying the correction pulse is preferably designed as a reaction engine for applying a radial force with respect to the longitudinal axis of the missile, characterized in that the device for determining the direction of flight has an axis which coincides therewith and which is always parallel or runs coaxially to the missile longitudinal axis ( 16; 23; 35 ).