EP3882564A1 - Guided missile and method for ejecting a guided missile from a support platform - Google Patents

Abstract

The invention relates to a method for dropping a guided missile (2) which can be steered about a roll, pitch and yaw axis from a carrier platform (1), the guided missile (2) having two pairs of rudder elements (3a, 3b, 4a, 4b) for generating Roll, pitch and yaw moments, the rudder elements (3a, 3b, 4a, 4b) of each pair being arranged opposite one another on the circumference of the fuselage (5) of the guided missile (2) in relation to the longitudinal axis of the guided missile (2), and one Flight control device (6) for controlling the alignment of the two pairs of rudder members (3a, 3b, 4a, 4b). Before the launch from the carrier platform (1) or immediately after the launch from the carrier platform (1) when the guided missile enters a flow field (7) of the carrier platform (1), the rudder elements (3a, 3b, 4a, 4b) deflect in opposite directions ) each pair by means of the flight control device (6) in such a way that the flow resistance that the guided missile (2) experiences after being dropped in the flow field (7) of the carrier platform (1) is increased compared to a flow resistance that the guided missile (2) would experience if the rudder elements (3a, 3b, 4a, 4b) of each pair were in a zero position in which the flight direction-influencing surfaces of the rudder elements (3a, 3b, 4a, 4b) of the pairs are parallel to the longitudinal axis (L) of the guided missile ( 2) extend. After leaving the flow field (7) of the carrier platform (1), the rudder elements (3a, 3b, 4a, 4b) of each pair are deflected by means of the flight control device (6) in such a way that the guided missile (2) is steered towards a target also relates to a guided missile (2) set up to carry out the method.

Description

The invention relates to a method for dropping a guided missile that can be steered about a roll, pitch and yaw axis, comprising two pairs of rudder elements for generating roll, pitch and yaw moments from a carrier platform and such a guided missile.

Missiles are often controlled by rudders in relation to the direction of flight. The rudders are controlled by a flight control device in such a way that the missile adopts the requested flight direction and can fly quickly to a specified destination. To avoid damage to a carrier platform while a guided missile is being dropped from the carrier platform, it is known to mechanically lock the steering gear or the actuators for the rudder so that the rudder is held in a neutral position that does not generate any torque.

From the U.S. 3,355,130 A a guided missile with four rudders is known. The guided missile is constructed by the actuators in such a way that a so-called "squeeze mode" can be avoided, in which two pairs of oars arranged opposite one another are controlled so that one pair has a roll moment in one direction and the other pair a roll moment in generated the other direction. In other words, the guided missile is designed in such a way that rudder positions that lead to opposing forces are prevented.

the KR 10 2 136 266 B1 deals with a method for calculating and executing a non-maneuvering, braking wing offset command to reduce the flight speed of a guided missile.

In the case of light-weight guided missiles, when they are dropped or uncoupled from a carrier platform, e.g. from its weapon bay or from a launch rail, when entering the flow field generated by the carrier platform or the flow field prevailing below the carrier platform, the latter is deflected by the flow field and collides with the carrier platform. This can result not only in damage to the carrier platform, but in particular also in damage to the guided missile.

It is therefore an object of the invention to specify a method for ejecting a guided missile from a carrier platform, which method allows the guided missile to leave the carrier platform safely.

• zwei Paare Ruderorgane zur Erzeugung von Roll-, Nick- und Giermomenten aufweist, wobei die Ruderorgane jedes Paares bezogen auf die Längsachse des Lenkflugkörpers einander gegenüberliegend am Umfang des Rumpfes des Lenkflugkörpers angeordnet sind, und
• eine Flugregelungseinrichtung zur Ansteuerung der Ausrichtung der zwei Paare Ruderorgane aufweist,

umfasst die Verfahrensschritte:

1. a) gegensinniges Ausschlagen der Ruderorgane jedes Paares mittels der Flugregelungseinrichtung vor dem Abwurf von der Trägerplattform oder unmittelbar nach dem Abwurf von der Trägerplattform mit dem Einritt des Lenkflugkörpers in ein Strömungsfeld der Trägerplattform derart, dass ein Strömungswiderstand, den der Lenkflugkörper nach dem Abwurf im Strömungsfeld der Trägerplattform erfährt, erhöht ist gegenüber einem Strömungswiderstand, den der Lenkflugkörper erfahren würde, wenn sich die Ruderorgane jedes Paares in einer Nullstellung befinden würden, in welcher sich die flugrichtungsbeeinflussenden Flächen der Ruderorgane jedes Paares parallel zur Längsachse des Lenkflugkörpers erstrecken, und
2. b) nach Verlassen des Strömungsfelds der Trägerplattform Ausschlagen der Ruderorgane jedes Paares mittels der Flugregelungseinrichtung derart, dass der Lenkflugkörper in Richtung eines Ziels gelenkt wird.

This object is achieved by a method according to the features of patent claim 1. The method for dropping a guided missile which can be steered about a roll, pitch and yaw axis from a carrier platform, which

• has two pairs of rudder elements for generating roll, pitch and yaw moments, the rudder elements of each pair being arranged opposite one another on the circumference of the fuselage of the guided missile in relation to the longitudinal axis of the guided missile, and
• has a flight control device for controlling the alignment of the two pairs of rudder organs,

comprises the process steps:

1. a) opposing deflection of the rudder organs of each pair by means of the flight control device before the launch from the carrier platform or immediately after the launch from the carrier platform with the entry of the guided missile into a flow field of the carrier platform in such a way that a flow resistance that the guided missile has in the flow field after the launch Carrier platform experiences, is increased compared to a flow resistance that the guided missile would experience if the rudder elements of each pair were in a zero position in which the flight direction-influencing surfaces of the rudder elements of each pair extend parallel to the longitudinal axis of the guided missile, and
2. b) after leaving the flow field of the carrier platform, deflecting the control elements of each pair by means of the flight control device in such a way that the guided missile is steered in the direction of a target.

The invention is based on the knowledge that although unwanted deflections of rudder organs can be avoided with a mechanical lock, a mechanical lock also means an additional weight load for a guided missile, which can have a detrimental effect on its maximum range, especially if the guided missile is designed without a drive is.

Furthermore, the invention is based on the knowledge that the probability of unwanted roll, pitch or yaw moments causing damage to the carrier platform and / or guided missile shortly after the missile is dropped from the carrier platform, in the case of guided missiles with a low weight, such as less than 50 kg , is increased compared to heavier (steering) missiles, since they are weight-related and exposed to the flow field below the carrier platform for a longer period of time after being dropped.

The invention is further based on the consideration that the risk of damage to the carrier platform and / or guided missile can be reduced if the guided missile can leave the flow field below the carrier platform as quickly as possible, the time span or time phase in which the guided missile follows its release is located in the flow field caused by the carrier platform, so it is as short as possible.

These disadvantages or dangers can be avoided by the method according to the invention. Due to the opposing deflection of the rudder organs of the two pairs by means of the flight control device before being released from the carrier platform or immediately after being released from the carrier platform with the entry of the guided missile into the flow field of the carrier platform in accordance with process step a), the guided missile is braked in the flow field, whereas the carrier platform flies on. As a result, the guided missile occurs earlier than a "non-braking" position of the two pairs of rudder organs, ie at a zero position in which the flight direction-influencing surfaces of the rudder organs of the two pairs would extend parallel to the longitudinal axis of the guided missile or, in other words, in the flow direction the flow field below the carrier platform, whereby the time span in which collisions can occur is shortened and thus reduced.

The rudder organs of the two pairs are preferably deflected in opposite directions in such a way that no aerodynamic moments are generated or, in other words, no forces changing the direction of flight become effective except for an increased drag force. This ensures that a guided missile can be moved safely and without the risk of collision out of the flow field of the carrier platform by appropriate control of the two pairs of rudder elements by means of its flight control device.

An advantageous embodiment provides that the rudder elements of the two pairs in method step a) are deflected in opposite directions by means of the flight control device in such a way that no change in flight direction is brought about.

In order to generate roll-pitch-yaw moments, the rudder elements can be pivoted at least partially about a radial axis, also known as a pivot axis, in relation to the fuselage of the guided missile. In the present case, deflection in opposite directions is understood to mean that one of the two rudder elements of a pair is deflected clockwise and the other of the two rudder elements counterclockwise, based on the zero position, in which the surfaces of the control elements of each pair that influence the direction of flight extend parallel to the longitudinal axis of the guided missile. To avoid changes in flight direction, it is provided that the angles by which the rudder elements of the two pairs are deflected in opposite directions in relation to the zero position are equal in terms of amount, i.e. each rudder element of the two pairs is deflected by the same angle in terms of amount.

In the present case, a carrier platform is understood in particular to be an aircraft, such as an airplane or a drone. The guided missile can be dropped by ejecting it from a weapon bay on the carrier platform or by uncoupling it from a launch device attached to the carrier platform, such as a launch rail. It is assumed that the carrier platform and guided missile move in the same direction of flight immediately after they have been dropped from the carrier platform.

The flight control device is designed so that the rudder elements of the two pairs can be controlled in such a way that a roll-pitch-yaw movement of the guided missile is possible, whereby the guided missile can be guided along a desired trajectory, in particular to a target. For this purpose, data from a data memory of the flight control device, which contains information on a desired mission of the guided missile, can be used. Receipt of external data such as from the or a carrier platform via a corresponding receiving device, e.g. B. an antenna provided in the guided missile and associated electronics is possible. However, the guided missile is expediently equipped with a radar seeker or an optical viewfinder, with data or signals for actuation of the flight control device being generated from the information collected via a sensor system.

In one embodiment of the invention it is provided that at the latest after a period of one second, in particular at the latest after a period of half a second, particularly preferably at the latest after a period of 250 milliseconds, the flight control device activates the alignment of the two pairs of control elements in accordance with method step a ) ends and continues with the control according to process step b). For this purpose, the guided missile is equipped in one embodiment with a timer that is connected to the flight control device by circuitry. The timer is set up to sense the launch of the guided missile from the carrier platform and to forward a signal to the flight control device after a predetermined period of time after the launch has been sensed, which enables control in accordance with method step b). If the carrier platform is, for example, a combat aircraft in which the guided missile is arranged under a wing of the combat aircraft during a transport phase, it can be assumed that after a period of maximum 250 milliseconds after being dropped, the guided missile will be at a distance of 10 m, measured from the wing, and based on the flight direction of the carrier platform is located behind the fighter aircraft and has thus left the flow field prevailing under the carrier platform.

The invention further relates to a guided missile with two pairs of rudder elements for generating roll, pitch and yaw moments, the rudder elements of each pair being arranged opposite one another on the circumference of the fuselage of the guided missile with respect to the longitudinal axis of the guided missile, and with a flight control device for controlling the Alignment of the two pairs of rudder organs, which is set up to carry out one or more of the aforementioned method steps.

The method according to the invention is particularly suitable for guided missiles with two pairs of rudder organs, which are distributed in "X" shape or in "+" shape on the circumference of the fuselage of the guided missile and are arranged in particular in the rear area of the guided missile. The rudder members of the two pairs are preferably arranged at the same axial height in relation to the longitudinal axis of the guided missile on the fuselage. In the "+" form, in contrast to the "X" form, the four rudder elements of the two pairs are evenly distributed around the circumference same angular range of 90 °. In the case of an "X" shape, a radial axis closes with each of the two adjacent radial axes different angular range, each radial axis of the four rudder organs always including the same different angular ranges with their respective adjacent radial axes. With a guided missile of this type, an otherwise undesirable "squeeze mode" can be brought about in a targeted manner by deflecting the rudder elements in opposite directions in order to generate a braking effect after being dropped from a carrier platform. Compared to a guided missile with only one pair of rudder elements, a higher braking effect can be achieved by deflecting in the opposite direction, since here four - instead of only two - flight direction-influencing surfaces contribute to an increase in the flow resistance that the guided missile experiences in the flow field of the carrier platform. The in and for itself undesirable "squeeze mode" while guiding a guided missile towards a target, which leads to impairment of the maneuverability or steering ability of a guided missile due to the generation of opposing forces, is in the launch phase, i.e. in the period of time in which of the guided missile is exposed to the flow field below the carrier platform, deliberately used to prevent unwanted maneuvers that could steer the guided missile towards the carrier platform and to quickly bring the guided missile outside of the flow field prevailing below the flying carrier platform.

The method according to the invention can furthermore preferably be used for guided missiles which have a wing arranged on the upper side of their fuselage and in particular centrally arranged in relation to the total length of the fuselage. The top of a fuselage can be viewed as the side on which the wing is at the top in a horizontal flight direction of the guided missile. The wing extends laterally symmetrically over the fuselage in relation to the longitudinal axis of the fuselage. Such a wing, the longitudinal axis of which extends at least during a flight phase of the guided missile in the direction of a target transversely, i.e. perpendicular to the longitudinal axis of the guided missile, gives the guided missile high lift forces, whereby the maximum range or flight distance can be increased. A disadvantage of such a wing, however, is that, immediately after the missile has been dropped, it can lead to the missile being pressed - quasi passively - in the direction of the carrier platform and thereby being damaged. By means of the method according to the invention, the opposing deflection of the two pairs of rudder elements and the resulting braking effect can ensure that the guided missile is not so long in the flow field below the carrier platform and thus the time span for any collisions is reduced or the risk of collision is reduced.

In the case of guided missiles in which the wing is pivotable between a position in which it extends along the longitudinal axis of the fuselage and a position in which it extends transversely to the longitudinal axis of the fuselage, it can be provided that the flight control device is set up in this way that the wing is only pivoted from its "longitudinal position" to its "transverse position" after the flow field prevailing below the carrier platform has been sensed. As a result, during a launch phase, i.e. the period in which the guided missile is still in the flow field below the carrier platform, lift forces caused by the wing and having a negative effect with regard to any collisions with the carrier platform can be avoided. In other words: In its "longitudinal position" the wing has no wingspan compared to its "transverse position", so it does not protrude laterally beyond the fuselage, or has only a reduced wingspan, i.e. it protrudes laterally beyond the fuselage, but not that far as compared to his "transverse position". If the guided missile is equipped with a timer described above, it can be provided that the flight control direction is set up to swivel the wing from its "longitudinal position" to its "transverse position after receiving the signal forwarded by the timer in addition to controlling the rudder elements in accordance with process step b) "to effect. For this purpose, the flight control device can be functionally connected to a pivot drive of the wing, e.g. connected by circuitry.

The method according to the invention can also be used in a particularly practical manner in the case of guided missiles weighing less than 50 kg, in particular less than 30 kg. Especially light missiles, also called ultra-light missiles, benefit from the increase in the flow resistance caused by the opposing deflection of the two pairs of rudder elements, as this at least partially compensates for the disadvantage of lower weight compared to higher weight missiles with regard to the time it takes to exit the flow field of the carrier platform.

The method according to the invention can also be used for guided missiles with a length of less than 1.5 m. Especially with small-sized and therefore often light-weight guided missiles or guided missiles with a relatively large volume with a relatively low weight and low density, a flow field prevailing beneath a carrier platform has a stronger effect on unintentional deflection compared to large and heavy missiles. This disadvantageous effect can at least be reduced by deflecting the two pairs of rudder elements in opposite directions.

The method according to the invention is preferably also suitable for non-powered guided missiles which, due to lack of propulsion, require a longer period of time to get out of the flow field under the carrier platform after being dropped.

The flight control device of the guided missile can have a control device, the control device preferably being designed as an analog and / or digital data processing device. For example, the control device is designed as a microcontroller, FPGA, DSP, etc. The control device is designed in particular in terms of program technology and / or circuit technology to adjust the control elements. The setting can be understood as a control, in particular as a control or regulation.

The control device is designed in particular in terms of data technology to set the rudder elements in a first operating state so that the guided missile is steered in a flight direction, in particular in a target flight direction, the guided missile having or experiencing a flight flow resistance.

The control device is designed in particular in terms of data technology to set the rudder elements in a second operating state so that the guided missile is steered in the same flight direction, in particular in the same target flight direction, and thereby has or experiences a braking flow resistance.

It is provided that - especially if the same flight speed is assumed - the braking flow resistance is designed to be greater than the flight flow resistance. Thus, the difference between the first and the second operating state is that the rudder elements are set so that the change is direction-neutral, but the flow resistance experienced by the guided missile is different, the flow resistance being greater than the braking flow resistance in the second operating state than the flow resistance is as flight flow resistance in the first operating state. The flow resistance relates in particular to the same airspeed.

It is proposed that the first operating state be designed as a flight operating state and the second operating state as a launching operating state. The release operating state is in particular programmatically and / or in terms of data technology activated. The release operating state is present when the guided missile is in a release phase, in particular when the guided missile is in the flow field prevailing below the carrier platform. The flight operating state occurs in a flight phase after the launch phase of the guided missile, in particular when it has left the flow field prevailing below the carrier platform.

One consideration here is that the second operating state stabilizes the guided missile on departure / take-off from the carrier platform, thereby improving the separation behavior from the carrier platform. A safe departure of the guided missile is at least supported or made possible in the first place by braking the guided missile by increasing the flow resistance in the launching operating state compared to the flight operating state. The launch mode has a similar effect to a braking parachute, so the aircraft flies past the guided missile, so to speak. The guided missile can thus quickly leave the flow field of the aircraft. In particular, small guided missiles with a relatively large volume with a relatively low weight and low density can thus be braked from the near field of the aircraft and in this way avoid a collision with the aircraft.

The flight direction is preferably designed as a neutral flight direction and / or as a straight flight direction. In particular, a total pitching moment and a total yawing moment of the guided missile are set to zero by the two pairs of rudder members, at least in the launching operating state.

The guided missile has a first pair of rudder members. The control members of the first pair are preferably arranged opposite one another on the circumference of the guided missile. It is provided that the rudder elements of the first pair are set in opposite directions in the release operating state. In this way, a first rolling moment and, at the same time, an increased flow resistance as braking flow resistance are generated.

The guided missile also has a second pair of rudder members. The control elements of the second pair are arranged opposite one another on the circumference of the guided missile. The rudder members of the first and the second pair are preferably distributed regularly in the direction of rotation around the guided missile. It is provided that the rudder elements of the second pair are set in opposite directions in the release operating state. In this way there is a second rolling moment and, at the same time, an increased one Flow resistance generated as a braking flow resistance. The total rolling moment, which is formed from the first rolling moment and the second rolling moment, is preferably zero. As an alternative or in addition, the first and the second roll moment are in opposite directions so that no aerodynamic roll moment is generated. As a further alternative, the total roll moment in the release operating state can be non-zero.

In a preferred embodiment, the guided missile has an explosive device for military use. It is particularly preferred that the guided missile has a mass of less than 50 kg, preferably less than 40 kg and in particular less than 30 kg. In particular, it is a light test missile.

In one embodiment, the guided missile is designed as a non-propelled missile. In particular, the guided missile is implemented as a glide bomb. However, it can also be provided that the guided missile has a drive.

Another object of the invention relates to a carrier platform, the carrier platform having at least one guided missile as described above or according to one of the preceding claims. In particular, the carrier platform can assume a launching state, the guided missile being arranged directly adjacent to the carrier platform and / or in its flow field and assuming the launching operating state.

In one embodiment of the method, the rudder elements can be placed in a "squeeze mode" in a phase of the launch of the guided missile from the carrier platform. The "squeeze mode" is preferably characterized in that, apart from an increase in flow resistance, no aerodynamic total yaw moments and total pitching moments are generated, i.e. the guided missile adopts or maintains the flight direction, in particular the straight flight direction.

In an alternative or further development of the method, the rudder elements are set in such a way that the guided missile is steered in a straight flight direction as the flight direction and has an increased braking flow resistance.

In a further alternative or development of the method, the guided missile is designed according to one of the preceding claims or as described above, the release operating state being activated in the release phase and remaining set, for example, over a period of 250 milliseconds. In particular, the release operating state is activated at the latest when it is released and remains continuously activated at least until the guided missile is behind the carrier platform.

It is particularly preferred that the launch phase occurs when the guided missile is in the flow field of the carrier platform and / or when the guided missile is at a distance from the carrier platform of less than 20 m, preferably less than 10 m.

In one possible embodiment, the “squeeze mode” and / or the launch operating state is already set when the guided missile leaves the carrier platform. As an alternative to this, the guided missile first leaves the carrier platform and then the "squeeze mode" and / or the launch mode is set if the guided missile is still in the launch phase.

The description given so far of advantageous embodiments of the invention contains numerous features, some of which are summarized in several dependent claims. The features can, however, expediently also be viewed individually and combined into meaningful further combinations, in particular when claims are referred back, so that an individual feature of a dependent claim can be combined with an individual, several or all features of another dependent claim. In addition, these features can be combined individually and in any suitable combination both with the method according to the invention and with the guided missile according to the invention. Process features are thus also to be seen objectively formulated as properties of the corresponding guided missile, and functional features of the guided missile are also to be seen as corresponding process features.

The properties, features and advantages of this invention described above and the manner in which they are achieved will become clearer and more clearly understood in connection with the following description of the exemplary embodiments, which are explained in more detail in connection with the drawings. The exemplary embodiments serve to explain the invention and do not restrict the invention to the combination of features specified therein, not even with regard to functional features. In addition, suitable features of each exemplary embodiment can also be considered explicitly in isolation, removed from one exemplary embodiment, introduced into another exemplary embodiment to supplement it, and / or combined with any of the claims.

Show it:

Figure 1

a highly schematic representation of a carrier platform when a guided missile is dropped as an embodiment of the invention;

Figures 2a, 2b

a highly schematic, three-dimensional representation of a sub-area of the guided missile in an exemplary flight operating state and in an exemplary launching operating state;

Figures 3a, 3b

a highly schematic representation of a top view of a partial area of an exemplary guided missile with a wing on its upper side of the fuselage during a launch phase and a flight phase.

the Figure 1 shows a highly schematic representation of a carrier platform 1 in the form of an aircraft with a guided missile 2, the guided missile 2 in three different phases, namely transport phase TP, i.e. before launch, launch phase AP, in which the guided missile 2 is in the flow field below the carrier platform 1 is located, and flight phase FP, in which the guided missile 2 is guided onto a target, is shown.

The guided missile 2 is designed as a lightweight test missile and / or with a mass of less than 50 kg. The guided missile 2 carries an explosive device, in particular a military explosive device for destroying particularly hard military targets. The guided missile 2 is designed as a passive and / or unpowered missile.

As long as the guided missile 2 is in or on the carrier platform, it is in a transport phase TP or a transport phase is present. For example, the guided missile 2 is arranged in a loading area and / or ammunition area, such as a weapons bay, of the carrier platform 1. In the transport phase TP, the guided missile 2 is flown to a target area by means of the carrier platform 1.

In the launch phase AP, the guided missile 2 is brought out of the carrier platform 1 into the free flight space. For example, the guided missile 2 is dropped. The launch phase AP thus extends from the launch of the guided missile 2 up to the point in time at which the guided missile 2 is under the carrier platform 1 Has left flow field 7. Alternatively or additionally, the launch phase AP extends up to the point in time at which the guided missile 2 is located behind the carrier platform 1 in the direction of flight F and / or at which the guided missile 2 has taken a distance of at least 10 m from the carrier platform 1. In the launch phase AP, there is the risk that the guided missile 2 will be deflected by the flow field 7 prevailing below the carrier platform 1 and, in the worst case (passively), will be directed towards the carrier platform 1. In the release phase AP it must therefore be ensured that no collision occurs.

After the end of the launch phase and at the beginning of the subsequent flight phase FP, the guided missile 2 is located behind the carrier platform 1 and / or is more than 10 m or more than 20 m away from the carrier platform 1, so that the guided missile 2 is independent of the carrier platform 1 and / or flies, in particular glides, in the free flight space.

the Figures 2a, 2b show a part, in this case the tail part, of the guided missile 2 in a possible flight mode ( Figure 2a ), in which the guided missile 2 is guided on a desired trajectory in the direction of a target, and in a launch mode ( Figure 2b ).

The guided missile 2 has a plurality of rudder members 3a, 3b, 4a, 4b, more precisely, two pairs of rudder members, namely 3a, 3b and 4a, 4b, which are controllably arranged at the rear of a fuselage 5 of the guided missile 2. The body 5 has a rotationally symmetrical, in particular cylindrical or cigar-shaped, shape. It is also conceivable that the fuselage has an angular cross section, for example in the form of a rectangle or more generally in the form of a polygon.

The guided missile 2 also has a flight control device 6 for controlling the alignment of the rudder elements 3a, 3b and 4a, 4b. The flight control device 6 is designed in such a way that the rudder elements 3a, 3b, 4a, 4b can be controlled in such a way that a roll-pitch-yaw movement of the guided missile is possible, as a result of which the guided missile can be guided along a desired trajectory, in particular to a target can. The flight control device 6 can comprise a control device designed as a digital or analog data processing device, which can implement the control of the rudder elements 3a, 3b and 4a, 4b via a control, regulation or another control strategy.

The rudder members 3a, 3b and 4a, 4b are arranged in pairs, the rudder members 3a, 3b forming a first pair and the rudder members 4a, 4b forming a second pair. The rudder members 3a, 3b are arranged at the same axial height in relation to the longitudinal axis L of the guided missile 2 on the fuselage 5 and are offset from one another by 180 ° in the direction of rotation. The rudder elements 4a, 4b of the second pair are arranged at the same axial height as the rudder elements 3a, 3b of the first pair and are also offset by 180 ° to one another in the direction of rotation. In the direction of rotation, the rudder elements 3a, 3b and 4a, 4b are arranged in such a way that they form an "X" shape. The rudder elements 3a, 3b, 4a, 4b can be pivoted at least partially around a radial axis A relative to the fuselage 5 of the missile 2 by means of the flight control device 6 to generate roll-pitch-yaw moments, whereby the flight direction of the guided missile 2 can be influenced.

In the transport phase TP, the setting of the rudder elements 3a, 3b and 4a, 4b is not relevant and can therefore in principle be freely selected. In the launch phase AP, the guided missile 2 is in a launch operating state, such as this one in FIG Figure 2b is shown. In the flight phase FP, the missile is in a flight operating state, such as this, for example, in FIG Figure 2a is shown.

In a possible flight status, like this one in the Figure 2a is shown, the rudder members 3a, 3b and 4a, 4b are aligned so that the guided missile 2 is in a straight flight direction as the flight direction. The rudder elements 3a, 3b and 4a, 4b are adjusted by the flight control device 6 in such a way that a maximum flight speed and / or a minimum flight flow resistance is implemented. The rudder elements 3a, 3b, 4a, 4b of the two pairs are in a zero position in which the flight direction-influencing surfaces of the rudder elements 3a, 3b, 4a, 4b of the pairs extend parallel to the longitudinal axis L of the guided missile 2. In order to steer the guided missile 2 - in particular on a desired trajectory - towards a target, the control elements 3a, 3b, 4a, 4b can be controlled by means of the flight control device 6 in such a way that a roll-pitch-yaw movement of the guided missile 2 is possible - This is brought about by corresponding pivoting of the rudder members 3a, 3b, 4a, 4b about their respective radial axis A. In the present case, the guided missile 2 is equipped with a viewfinder, for example a radar viewfinder or an optical viewfinder, with data or signals for actuation of the flight control device 6 being generated from the information collected via a sensor system.

In the drop mode, like this one in FIG Figure 2b is shown, the two pairs of rudder members 3a, 3b, 4a, 4b, are in a "squeeze mode". In the "Squeeze Mode "the respective rudder organs 3a, 3b and / or 4a, 4b are brought into a direction-neutral combination of rudder deflections for the guided missile 2. The rudder deflections are deflected in opposite directions so that no external forces are effective, except for a drag force caused by the braking flow resistance .

A safe departure of the guided missile 2 is at least supported or made possible in the first place by the braking of the guided missile 2, which can thus quickly leave the flow field 7 of the carrier platform 1. The "Squeeze Mode" works like a braking parachute. The carrier platform 1 thus flies past the guided missile 2, so to speak. In particular, small and light guided missiles 2 with low weight, but at the same time relatively high lift forces - as described here - can thus be braked from the near field of the carrier platform 1 and thus avoid a collision.

The pairs of rudder elements 3a, 3b and 4a, 4b are deflected in opposite directions before or immediately after the launch of the missile 2 from the carrier platform 1 in such a way that apart from an increase in resistance, no aerodynamic moments are generated, so the missile 2 continues to fly straight ahead . In other words: opposite a zero position of the control elements 3a, 3b, 4a, 4b, in which the flight direction-influencing surfaces of the control elements 3a, 3b, 4a, 4b extend parallel to the longitudinal axis L of the guided missile 2, as shown in FIG Figure 2a is shown and in which the guided missile 2 is not executing a steering maneuver, are in the in Figure 2b Shown release operating state, the flight direction influencing surfaces of the rudder members 3a, 3b, 4a, 4b pivoted about their radial axes in such a way that their flight direction influencing surfaces no longer extend parallel to the longitudinal axis L of the missile 2, but still no steering maneuver or no flight direction change of the missile 2 causes will.

In the Figure 2b the rudder element 3a is pivoted counterclockwise about its radial axis A and the rudder element 3b pivoted clockwise about its radial axis A. The rudder element 4a is pivoted counterclockwise about its radial axis A and the rudder element 4b is pivoted clockwise about its radial axis A. To avoid changes in flight direction, it is provided that the angles by which the rudder elements 3a, 3b, 4a, 4b are deflected in opposite directions with respect to the zero position are equal in terms of amount.

It can be provided that the pivoting of the rudder elements 3a, 3b, 4a, 4b in the launching operating state compared to a flight operating state and / or their zero position is selected so that no additional total yaw moments and total pitching moments, but different total rolling moments occur. The release phase AP can be designed to be particularly reliable if, in addition, no total roll torques are generated, but only the flow resistance as braking flow resistance is increased.

In the in the Figures 3a and 3b The embodiment shown is a central portion of a guided missile 2 is shown in a highly schematic plan view. The guided missile 2 is equipped with a pivotable wing 9 which is arranged centrally on the upper side of the fuselage 5 and based on the total length of the fuselage 5. The wing 9 is between a position in which it extends along the longitudinal axis L of the fuselage 5 ( Fig. 3a ), and a position in which it extends transversely to the longitudinal axis L of the fuselage 5 ( Fig.3b ), pivotable. It is provided that the two pairs of rudder elements 3a, 3b, 4a, 4b arranged in the tail area of the guided missile 2 in an "X" shape are deflected in opposite directions before they are released from the carrier platform 1 in such a way that the flow resistance that the guided missile 2 after the launch in the flow field of the carrier platform 1 is increased compared to a flow resistance that the guided missile 2 would experience if the rudder organs 3a, 3b, 4a, 4b of the pair were in a zero position in which the flight direction influencing surfaces of the rudder organs are 3a, 3b, 4a, 4b of the pair extend parallel to the longitudinal axis L of the guided missile 2. After leaving the flow field 7 of the carrier platform 1, the rudder elements 3a, 3b, 4a, 4b are deflected by means of the flight control device 6 in such a way that the guided missile 2 is steered in the direction of a target. When the flow field 7 prevailing below the carrier platform is left, the wing 9 is moved from its "longitudinal position" by means of the flight control device 6 ( Fig. 3a ) in its "transverse position" ( Figure 3b ) pivoted, the rudder elements 3a, 3b, 4a, 4b are aligned by the flight control device in such a way that the guided missile 2 is steered in the direction of a target. The guided missile 2 is for this purpose with a timer 8, as in FIG Figures 2a, 2b shown, equipped, which is connected to the flight control device 6 in terms of circuitry. The timer 8 is set up to sense the launch of the guided missile 2 from the carrier platform 1 and to forward a signal to the flight control device 6 after a predetermined period of time, in the present case 250 milliseconds, after the launch has been sensed. After the period of time has elapsed, the flow field prevailing under the carrier platform 1 is deemed to have been abandoned and the flight control device 6 begins with the reception of the signal from the timer 8 the control of the rudder elements 3a, 3b, 4a, 4b in such a way that the guided missile 2 is steered on a desired trajectory in the direction of a target. The signal has the effect that the launching operating state is ended by the flight control device 6 and the flight operating state is started. At the same time, the flight control device is set up so that when the signal from the timer 8 is received, the wing 9 is pivoted from its “longitudinal position” to its “transverse position”. For this purpose, the flight control device 6 is operatively connected to a pivot drive of the wing 9.

List of reference symbols

1

Carrier platform

2

Guided missile

3a, 3b

Rudder organs of the first pair

4a, 4b

Rudder organs of the second pair

5

hull

6th

Flight control device

7th

Flow field

8th

timer

9

wing

TP

Transport phase

AP

Release phase

FP

Flight phase

L.

Longitudinal axis

A.

radial axis

Claims (12)

Method for dropping a guided missile (2) which can be steered about a roll, pitch and yaw axis from a carrier platform (1), the guided missile (2) having - Two pairs of rudder members (3a, 3b, 4a, 4b) for generating roll, pitch and yaw moments, the two rudder members (3a, 3b, 4a, 4b) of each pair in relation to the longitudinal axis (L) of the guided missile (2 ) are arranged opposite one another on the circumference of the fuselage (5) of the guided missile (2), and - a flight control device (6) for controlling the alignment of the two pairs of rudder organs (3a, 3b, 4a, 4b), comprehensive the procedural steps a) opposing deflection of the rudder elements (3a, 3b, 4a, 4b) of each pair by means of the flight control device (6) before being released from the carrier platform (1) or immediately after being released from the carrier platform (1) with the entry of the guided missile into a Flow field of the carrier platform (1) in such a way that the flow resistance that the guided missile (2) experiences after being dropped in the flow field of the carrier platform (1) is increased compared to a flow resistance that the guided missile (2) would experience if the rudder elements ( 3a, 3b, 4a, 4b) of each pair would be in a zero position, in which the flight direction-influencing surfaces of the control elements (3a, 3b, 4a, 4b) of the pairs extend parallel to the longitudinal axis of the guided missile (2), and b) after leaving the flow field of the carrier platform (1) deflecting the rudder elements (3a, 3b, 4a, 4b) of each pair by means of the flight control device (6) in such a way that the guided missile (2) is steered towards a target. Method according to claim 1,
characterized,
that the rudder elements (3a, 3b, 4a, 4b) of each pair in method step a) are deflected in opposite directions by means of the flight control device (6) in such a way that no change in flight direction is caused. Method according to claim 1 or 2,
characterized,
that after a period of one second at the latest, in particular after a period of half a second at the latest, the flight control device (6) terminates the control of the alignment of the two pairs of rudder elements (3a, 3b, 4a, 4b) in accordance with method step a) and continues with the control continues according to process step b). Guided missile (2) with - Two pairs of rudder members (3a, 3b, 4a, 4b) for generating roll, pitch and yaw moments, the rudder members (3a, 3b, 4a, 4b) of each pair lying opposite one another in relation to the longitudinal axis of the guided missile (2) Circumference of the fuselage (5) of the guided missile (2) are arranged, and - a flight control device (6) for controlling the alignment of the two pairs of control elements (3a, 3b, 4a, 4b), set up to carry out a method according to one or more of the preceding claims. Guided missile (2) according to claim 4,
marked by
two pairs of rudder organs (3a, 3b, 4a, 4b), which are arranged in "X" shape or in "+" shape on the circumference of the hull (5). Guided missile (2) according to one of claims 4 or 5,
marked by
one on the top of the fuselage (5) and in relation to the total length of the fuselage (5) in particular centrally arranged wing (9). Guided missile (2) according to claim 6,
characterized,
that the wing (9) is pivotable between a position in which it extends along the longitudinal axis (L) of the fuselage (5) and a position in which it extends transversely to the longitudinal axis (L) of the fuselage (5) . Guided missile (2) according to one of claims 4 - 7,
marked by
a weight of less than 50 kg. Guided missile (2) according to one of claims 4 - 8,
marked by
a length of less than 1.5 m. Guided missile (2) according to one of claims 4 - 9,
characterized,
that it is designed as a non-propulsion guided missile (2). Guided missile (2) according to one of claims 4 to 10,
marked by
a timer which is connected to the flight control device (6) in terms of circuitry. Guided missile (2) according to claim 11,
characterized,
that the timer is set up to sense a launch of the guided missile (2) from a carrier platform (1) and to forward a signal to the flight control device (6) after a predetermined period of time after the launch has been sensed, which enables control in accordance with method step b) .

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