EP2594891A2 - Procédé destiné à repousser une fusée balistique se rapprochant en volant et système d'interception - Google Patents

Procédé destiné à repousser une fusée balistique se rapprochant en volant et système d'interception Download PDF

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Publication number
EP2594891A2
EP2594891A2 EP12007727.6A EP12007727A EP2594891A2 EP 2594891 A2 EP2594891 A2 EP 2594891A2 EP 12007727 A EP12007727 A EP 12007727A EP 2594891 A2 EP2594891 A2 EP 2594891A2
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EP
European Patent Office
Prior art keywords
rocket
ballistic missile
interceptor
missile
ballistic
Prior art date
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Granted
Application number
EP12007727.6A
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German (de)
English (en)
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EP2594891A3 (fr
EP2594891B1 (fr
Inventor
Thomas Dr. Kuhn
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Diehl Defence GmbH and Co KG
Original Assignee
Diehl BGT Defence GmbH and Co KG
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Publication of EP2594891A2 publication Critical patent/EP2594891A2/fr
Publication of EP2594891A3 publication Critical patent/EP2594891A3/fr
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Publication of EP2594891B1 publication Critical patent/EP2594891B1/fr
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F41—WEAPONS
    • F41H—ARMOUR; ARMOURED TURRETS; ARMOURED OR ARMED VEHICLES; MEANS OF ATTACK OR DEFENCE, e.g. CAMOUFLAGE, IN GENERAL
    • F41H11/00—Defence installations; Defence devices
    • F41H11/02—Anti-aircraft or anti-guided missile or anti-torpedo defence installations or systems

Definitions

  • the invention relates to a method for averting an approaching ballistic missile, in which a probable trajectory of the ballistic missile is determined, an interceptor rocket is launched, it flies to the probable trajectory and destroys the ballistic missile in flight by means of its active charge.
  • the focus of this invention is interception systems for defense or methods for the defense of tactical ballistic missiles (so-called “tactical ballistic missiles", short “” TBM ”) and ballistic medium-range missiles (so-called” intermediate range ballistic missiles ", short” IRBM ").
  • Tactical ballistic missiles usually have a range of up to 500 km, while with medium range ballistic missiles ranges up to 3,000 km can be achieved.
  • Your flight is divided into three characteristic mission phases.
  • the rocket engine operates and accelerates the ballistic missile up to the burnout speed dependent on the desired range.
  • the low-friction flight phase outside the atmosphere, most of the flight route is traveled at speeds of between about 1000 m / s and 5000 m / s.
  • the third flight phase also known as the final phase, begins with the re-entry into the atmosphere and ends with the hitting the finish.
  • a high performance interceptor missile is required.
  • One implementation could be on airborne platforms, e.g. As unmanned aerial vehicles, which revolve in a crisis situation over a critical area, register a rocket launch and then start a guided missile. Even such an interception system is very complex and expensive and possible only for limited areas.
  • the object directed to the method is achieved by a method of the type mentioned, in which the interceptor rocket according to the invention flies before the ignition of the active charge in the direction of flight of the ballistic missile before this.
  • the combat takes place expediently in the final phase, ie after reentry of the ballistic missile into the atmosphere.
  • the interceptor missile also flies in front of or next to the ballistic missile at the moment of the ignition of the active charge.
  • the invention is based on the consideration that the difficulty in an endoatmosphärischen combat during the final flight phase of the ballistic missile is that the relative velocity between the ballistic missile and the counter-launching missile flying this is very high.
  • the active charge is ignited only in the vicinity of the ballistic missile, the mission success is due to the very high velocity of encounter (also referred to in the specialist terminology as "closing velocity") between 2000 m / s for simple tactical ballistic missiles and over 5000 m / s not secured in ballistic medium-range missiles, as the effect of the charge can not be placed with sufficient precision. It should therefore be sought a direct hit.
  • a significant reduction in the relative speed makes it possible to dispense with a direct coincidence of the interceptor with the ballistic missile, and a detonation of the active charge in the vicinity of the ballistic missile may be sufficient for the success of the mission.
  • the differential speed between the two missiles is reduced to the overtaking speed, that is reduced compared to a counter-attack by twice the flight speed of the interceptor. This simplifies placing a destructive effect of the active charge, so that can be dispensed with a direct hit.
  • the cost of sensors and actuators of the interceptor rocket can be significantly reduced, so that a much cheaper interceptor rocket is made possible.
  • the control of the interceptor rocket can be considerably simplified, since it can be dispensed with the consideration of the tumbling motion of the approaching ballistic missile.
  • the interceptor rocket is conveniently controlled to a lane within the anticipated wobble path of the ballistic missile.
  • the tumble track is a spiral track that traverses a substantially cylindrical interior.
  • the central orbit within a spiral wobble path can be calculated by means of an averaging of measurement data of the wobble path.
  • the interceptor rocket is controlled in the center of the spiral wobble path.
  • a flyby of the ballistic missile at a distance of the same size is to be expected.
  • an ambient ignition of the active charge can destroy the ballistic missile or its active set, without a direct hit being necessary.
  • the attack of the interceptor missile on the ballistic missile is advantageously carried out in the rectified flight.
  • the interceptor missile does not leave the atmosphere and advances the ballistic missile after re-entering the atmosphere.
  • the order of the steps of determining the probable trajectory and starting the interceptor rocket is arbitrary.
  • the interceptor missile can fly in total to the anticipated trajectory of the ballistic missile or drop the rocket motor after a boost phase and fly only as a missile head for the expected trajectory of the ballistic missile.
  • the attack on the ballistic missile can be done by the interceptor missile overall or only part of it, such as the rocket head.
  • a rocket head is referred to in this context for the sake of simplicity as an interceptor rocket.
  • the interceptor rocket flies according to the invention before the approaching ballistic missile, previously a counter-directional movement of the two missiles can be done.
  • the interception system with the interceptor rocket is stationed in the vicinity of an object to be protected, for example a city. After a launch, the interceptor rocket first flies against the ballistic missile, then turns to fly ahead of the ballistic missile to reduce the differential speed. A direct turn near the expected trajectory is just as possible as the wide swinging on the probable trajectory with a more energy efficient large flight radius.
  • the corresponding approach maneuver can be made dependent on the stationing of the interception system and the approach direction of the approaching ballistic missile on the object to be protected.
  • the ignition of the active charge takes place depending on the overtaking of the ballistic missile on the interceptor rocket. This makes it possible to dispense with a pre-calculation of overtaking, whereby the sensor technology of the interceptor rocket can be further simplified. Ignition may occur when overtaking or depending on time take place at least one overtaking parameter, for example, the differential speed. Appropriately, the ignition of the active charge is triggered by the overtaking process directly.
  • a further advantageous embodiment of the invention provides that the ballistic missile is measured by at least one further sensor carrier and a rough estimate of the probable trajectory of the ballistic missile is first determined from measured data of the further sensor carrier.
  • a sensor carrier is understood to be a platform which is provided with at least one sensor which is capable of measuring a ballistic missile.
  • the further sensor carrier may be an unmanned missile, for example a reconnaissance drone, whereby, of course, the use of manned aircraft is possible.
  • ground-based sensor carriers such as radar stations are conceivable.
  • the ballistic missile is measured by a plurality of sensor carriers, so that triangulation is enabled and performed to calculate the expected trajectory.
  • the interceptor rockets after determining the rough estimate of the expected trajectory of the ballistic missile on this in the direction of flight and thus flies in front of the ballistic missile.
  • the interceptor misses the ballistic missile in advance of the ballistic missile and determines a more accurate estimate of the anticipated trajectory of the ballistic missile.
  • the interceptor suitably has its own sensor.
  • the interceptor then flies ahead of this more precisely estimated prospective trajectory of the ballistic missile.
  • the measurement is expediently carried out already during the swivel-in on the probable trajectory, which is iteratively specified by a further sensor carrier, in particular starting from a rough preamble on the basis of a rough estimate of the probable trajectory.
  • a determination of the probable trajectory of the ballistic missile from the measured data can be done on the ground or by the interceptor missile itself. In the case of a determination on the ground, corresponding control commands can be transmitted via uplink to the interceptor rocket.
  • the simplest, however, is the determination of probable trajectory through the interceptor itself, so that it can swivel by own data on the trajectory.
  • a further advantageous embodiment of the invention provides that the ballistic missile is measured by at least one further sensor carrier and measurement data are transmitted from the sensor carrier to the interceptor rocket.
  • a control computer of the interceptor rocket can then be determined from the measured data of the sensor carrier a first rough estimate of the probable trajectory of the ballistic missile. This makes it possible for the interceptor rocket to swerve on this roughly estimated expected trajectory in the direction of flight of the ballistic missile.
  • the measurement data from a sensor carrier is raw data, such as the time stamped positions and bearing angle of the sensor carrier with respect to the ballistic missile.
  • Raw data is understood herein to be unfiltered data.
  • Raw data has the advantage that it can be easily merged with its own measurement data of the interceptor rocket, without having to merge data from different processing levels.
  • a control computer of the interceptor missile measurement data of the interceptor rocket merges with measurement data of one or more further sensor carrier, compared to an only on measurement data of the interceptor or Messdatein a sensor carrier estimated estimate of an anticipated trajectory even more accurate Estimate of an expected trajectory of the ballistic missile on which the interceptor can then swing in the direction of flight of the ballistic missile.
  • a time of the overtaking process is determined, expediently by the interception rocket itself, and an ignition of the active charge as a function of this time takes place.
  • the determination of the azimuthal direction in which the ballistic missile overtakes the interceptor missile can be dispensed with.
  • a particularly suitable ignition could be a polar ignition at two opposite ends of the active charge, whereby some of the splinters move in a plane perpendicular to the axis of symmetry of the interceptor rocket with increased speed and thus can achieve an increased end ballistic performance. This would create a splinter curtain around the interceptor rocket that will strike the ballistic missile at its warhead level, assuming the firing point is correctly set, regardless of the azimuthal angle the ballistic missile passes the interceptor missile.
  • a more favorable orientation of the explosive force of the active charge is to align the explosive force by a multi-point ignition in expediently a largely one-dimensional direction.
  • One of the ways is to provide the interceptor rocket with an additional engine. This is expediently designed to give the interceptor rocket a lateral acceleration with which the interceptor missile is brought directly into the trajectory of the passing ballistic missile or at least in its direction. An ignition of the additional engine is expediently carried out shortly before a probable coincidence of the two missiles. For this purpose, advantageously, the time of the overtaking process and in particular an azimuthal direction in which the ballistic missile passes the rocket head, determined in advance.
  • the ignition of the additional engine expediently causes at least the active charge to be moved towards the passing ballistic missile.
  • the active charge is moved towards the passing ballistic missile.
  • Prerequisite for this is a measurement of distance and speed difference between the rockets.
  • an active sensor is necessary.
  • the tumbling can be detected passively, for example by an infrared sensor.
  • An active sensor can be dispensed with in a second possibility.
  • several interceptor missiles which are at least near the expected trajectory, fly ahead of the ballistic missile.
  • several missile heads ejected from a launcher are referred to in this context as multiple interceptor missiles.
  • the interceptor missiles can be arranged here around an expected trajectory of the ballistic missile.
  • the ballistic missile is measured in each case by means of at least one sensor of an interceptor rocket. Measurement data from all interceptor missiles may be relayed to at least one of the interceptor rockets and a distance and a differential velocity of the ballistic missile to at least one of the interceptor missiles may be determined from the measurement data.
  • the active sensor is dispensable in this embodiment, since the differential speed and distance can be detected, for example, triangulation.
  • the wobble helix can also be detected as an anticipated trajectory.
  • One or more of the interceptor missiles will ultimately fight the ballistic missile, and this interceptor missile (s) will conveniently calculate the expected trajectory.
  • the invention is also directed to an interceptor system comprising at least one interceptor rocket having a rocket motor, a control computer, a steering system, and an active charge.
  • control computer be prepared to determine an anticipated trajectory of the ballistic missile, to control the steering system in such a way that at least the rocket head flies in front of the latter in the direction of flight of the ballistic missile, and in particular to ignite the active charge in response to a ballistic missile overtaking the interceptor rocket. Due to the greatly reduced differential speed compared to the opposite flight, a direct coincidence of the two missiles can be dispensed with and the interception system can be kept simple and therefore cost-effective.
  • the rocket motor (booster) is dropped after a boost phase, so that only the rocket head - also referred to here as an interceptor rocket - pivots in the expected trajectory of the approaching ballistic missile.
  • the rocket motor it is possible for the rocket motor to remain on the rocket head and launch the interceptor missile as a whole in front of and destroy the ballistic missile.
  • the control computer is advantageously prepared to control the execution of any one, several arbitrary or all of the mentioned method steps. Such a preparation may be provided by a corresponding control program of the control computer, the sequence of which - such as in conjunction with suitable input signals, such as sensor signals - causes such a control.
  • the control computer expediently comprises electronic elements, such as a processor and data memory, which are necessary for running the control program.
  • the interceptor missile comprises a data receiver for receiving measurement data from other sensor carriers for calculating the anticipated trajectory of the ballistic missile.
  • the interceptor missile comprises a seeker head directed to the rear, the field of view of which is directed at the trajectory already passed through by the interceptor rocket or its rocket head.
  • the rearwardly directed seeker head makes it possible to passively recognize the ballistic missile approaching from behind, for example by heat radiation emitted by the latter when using an IR seeker head.
  • the seeker head comprises active sensors.
  • the interceptor rocket has a separation point for separating the rocket motor from a rocket head, and a rearward seeker, which is arranged on the rocket head so that it receives a free field of view by the separation of the rocket motor backwards.
  • the seeker head is protected by the separation point during storage of the interceptor rocket, and especially during the boost phase. Only by separating the rocket motor, the seeker reaches the surface or gets a clear view to the rear.
  • control computer is prepared to measure the ballistic missile approaching from behind with the aid of the seeker head and to determine the probable trajectory of the ballistic missile from measured data.
  • This trajectory may be a more precise estimated probable trajectory than, for example, previously calculated from measurement data of at least one further sensor carrier.
  • control computer is expediently designed to hide, ie ignore, a tumbling motion of the ballistic missile approaching from behind, and to control the interceptor missile on an expected central orbit within the real wobble path of the ballistic missile.
  • a further advantageous embodiment of the invention proposes that the interceptor rocket has a proximity sensor that is prepared to detect overtaking of the ballistic missile. Expediently, a sensory recognition of the overtaking process is performed as a currently occurring event. On a pre-calculation of the overtaking time and in particular the overtaking direction can then be dispensed with. Both can be detected sensory in real time.
  • a simple embodiment of the proximity sensor includes having it emitters that are prepared to place an active beam curtain in the environment.
  • the radiation curtain can be placed in the vicinity of the interceptor rocket and expediently surrounds the interceptor rocket by 360 °, ie completely.
  • the proximity sensor and / or the control computer are prepared to detect a breakthrough of the beam curtain by the ballistic missile. This can be done for example by a re-radiation of the rays and the corresponding sensory recognition of the reverberation. Appropriately, several sensors are present, so that a 360 ° monitoring of the beam curtain is easily possible.
  • an additional engine in particular a transverse thrust engine, present and the control computer prepared to ignite the engine in such a way that at least the effective charge undergoes a lateral acceleration on the passing ballistic missile, the effect can be increased again.
  • Fig. 1 shows an interceptor rocket 2 with a rocket head 4 and a rocket motor (booster) 6.
  • the rocket head 4 is equipped with a steering system 8 having 10 fixed rudder 12 on fixed wings.
  • the rocket motor 6 is equipped with wings 14, which are not necessarily movable.
  • Fig. 2 shows the interceptor rocket 2 from Fig. 1 in a somewhat more detailed schematic representation.
  • the rocket motor 6 is from the rocket head 4 at a separation point 16th severable. It contains propellant 18 and is equipped with a parachute 20 to which it floats after burnout of the propellant 18 to the ground.
  • the rocket head 4 is equipped at its front tip with a warhead, which contains an active charge 22.
  • Behind a control computer 24 is arranged, which includes a guide unit for calculating trajectories and control signals for controlling the rocket head or the interceptor rocket.
  • the control computer is equipped with an inertial navigation system, a GPS receiver and a radio interface for exchanging data with a ground station and / or other sensor carriers such as missiles.
  • the control computer includes an actuator control unit prepared to control the moving rudders 12.
  • a seeker head 26 is arranged, the field of view is directed to the rear.
  • the seeker 26 is an IR seeker for detecting infrared radiation. It contains a movable optic for panning the visual field.
  • FIG. 12 shows a flowchart of a method for repelling an approaching ballistic missile 28.
  • the launch of the ballistic missile 28 and the launch phase during which the ballistic missile 28 is ascending are detected by a plurality of sensor carriers, according to this embodiment missiles 30, which fly far away from the starting location of the ballistic missile 28 at high altitude.
  • the launch and / or flight phase of the ballistic missile 28 may be detected off-atmosphere by one or more satellites, also referred to herein as a sensor carrier or missile.
  • the missile or missiles 30 determine their position and their angle of view to the flying ballistic missile 28 at known times. From this raw data, they or a ground station calculate the prospective flight path 36 and airspeed and the respective location of the ballistic missile 28. On the basis of this flight data, the Interceptor rocket 2 started from a launcher 32. The launch of the interceptor rocket 2 can be ground-based, ship-based or airborne. After completion of the boost phase of the rocket motor 6 is separated from the rocket head 4 and the rocket motor 6 floats on the parachute 20 to the ground. Before or after the separation of the rocket motor 6 from the rocket head 4, the interceptor rocket 2 or the rocket head 4 from the pure ascent direction already becomes expected trajectory 36 steered. The subsequent steering into the trajectory 36 can be done in different ways.
  • the flight data of the ballistic missile 28 can be transmitted to the rocket head 4, which then steers into the flight path 36 on the basis of this data.
  • a particularly advantageous method is that the raw data are transmitted to the rocket head 4 and the control computer 24 calculates the expected trajectory 36 of the ballistic missile 28 itself and controls the self-calculated trajectory 36.
  • This method has the advantage that own measurement data can be merged with these raw data in a particularly simple and error-free manner, so that later on a more precise prospective trajectory 36 can be estimated more accurately.
  • the corresponding data of the external sensor system are supplied by uplink to the control computer 24.
  • the rocket head 4 tries to detect the ballistic missile 28 optically using the seeker head 26.
  • the rocket head 4 is at this time, for example, at an altitude of about 20 km. It is there to go out of an infrared range of well over 25 km, so that the approaching ballistic missile 28 can be tracked by the seeker head long enough and the prediction of the expected trajectory 36 can be corrected using the data of the seeker head 26.
  • the rocket head 4 pivots on this trajectory 36 and continues to fly in front of the ballistic missile 28.
  • the ballistic missile 28 is permanently tracked and the probable trajectory 36 is corrected at regular intervals.
  • the reference numeral 34 in FIG Fig. 3 In this case, the line of sight from the seeker head 26 to the ballistic missile 28.
  • the rocket head 4 flies at a speed of, for example, 700 m / s and turns at a high speed, z. B.> 1500 m / s, approaching target in the way.
  • the differential speed between the two rockets 4, 28 is about 800 m / s.
  • a re-entry speed of the rocket into the atmosphere of about 10 Mach is to reckon with a re-entry speed of the rocket into the atmosphere of about 10 Mach.
  • the approaching rocket is slowed down during the passage through the atmosphere to about 3 Mach.
  • the flight speeds of the interceptor rocket 2 is thus a suitable area of coincidence of the two rockets 4, 28 to calculate, on the one hand at high altitude and as far away from the protected target, such as the city, and on the other hand, the medium-range missile already a relatively low speed has to be destroyed with relatively little effort can.
  • the location of this area is of course also dependent on the achievable airspeed of the interceptor rocket 2.
  • the actual trajectory of the ballistic missile 28 is on a spiral with a diameter of, for example, five meters and a swept frequency of z. B. about one hertz per revolution.
  • the differential speed between the two rockets 4, 28 and the distance between them is not calculated.
  • the rocket head 4 flies alone on sight.
  • a time t 1 at which the ballistic missile 28 passes the rocket head 4 is not known in advance.
  • Fig. 4 is shown schematically that this overtaking process can still be easily registered by the rocket head 4 and the control computer.
  • the rocket head 4 with a plurality of radiation sensor units 38 (see Fig. 1 ) fitted.
  • the radiation radiated by them 40 surrounds the rocket head like a radiate in all directions, so that extends around 360 ° 4. If the tip of the ballistic missile 28 reaches this radiate wreath 40 and reflects the radiation back to the rocket head 4, the reflected radiation is registered by the radiation sensor unit 38. Because of this, it is assumed that the overtaking process has occurred.
  • This time t 1 is registered by the control computer 24 and it is a time until the time t 2 wait until the ignition of the active charge 22 is initiated.
  • the azimuth angle must be known, at which the rocket 28 overtakes the rocket head 4, ie the direction in which the beam curtain 40 is broken. If the emission sensor unit 38 comprises a plurality of emission means or sensors, then this azimuth angle can be detected sufficiently fine. Depending on which of the annularly arranged around the rocket head 4 sensors detects the backscattered radiation most, the azimuth angle is determined accordingly and controlled the blasting of the active charge 22 accordingly.
  • the elevation angle 44 is 90 °. However, it is also conceivable that this angle 44 is less than 90 °, the beam curtain 40 is thus oriented backwards. Accordingly, an overtaking operation is detected earlier. It should be noted, however, that the farther the distance between the rockets 4, 28, the sooner the overtaking process is noticed. Does the rocket head 4 z. B. via a device for distance measurement, such as a laser rangefinder, about the distance between the missiles 4, 28 can be determined. The elevation angle 44 can then be adjusted so that the explosive effect of the active charge 22 is always optimal.
  • the splinters of the active charge 22 need the longer to to reach the rocket 28, the farther it is away from the rocket head 4. It is thus advantageous to ignite the active charge 22 the sooner, the farther the distance between the two rockets 4, 28 is.
  • the differential velocity between the missiles 4, 28 is known, for example, from the continued estimation of the prospective trajectory 36, as related to Fig. 3 previously described.
  • the elevation angle 44 can be adapted to this difference. The greater the differential speed, the smaller the elevation angle 44 and the farther back the beam 40 is thus directed.
  • the rocket head 4 can be controlled by an additional engine 46 in the form of a transverse thrust engine in the direction of the passing ballistic missile 28. This is in Fig. 5 indicated schematically. Shortly before the overtaking process, the engine 46 is ignited and the rocket head 4 is guided into the actual trajectory 48 of the rocket 28. For this purpose, however, it is necessary to include the wobbling motion in the trajectory calculation in order to know in advance in which azimuth angle the rocket 28 passes. Accordingly, the rocket head 4 must be rolled before igniting the shear thrust engine.
  • the differential speed and the distance between the two rockets 4, 28 are determined by means of active sensors. From this an estimated impact point (PIP) can be calculated, in which the rocket head 4 is now controlled by means of the engine 46.
  • PIP estimated impact point
  • a plurality of interceptor rockets 2 are started instead of a single interceptor rocket 2. This can be done from a single launcher 32 or from multiple launcher 32. Accordingly, several rocket heads 4 are available, which are initially grouped around the expected trajectory 36. For this purpose, the rocket heads 4 remain so far apart that the distance between them and the approaching ballistic missile 28 can be determined by triangulation, and from this also the differential speed. In this way, can be dispensed with active sensors and the rocket heads 4 can selectively put the missile 28 in the way.

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  • Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Radar, Positioning & Navigation (AREA)
  • Remote Sensing (AREA)
  • General Engineering & Computer Science (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
EP12007727.6A 2011-11-21 2012-11-15 Procédé destiné à repousser une fusée balistique se rapprochant en volant et système d'interception Not-in-force EP2594891B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102011118927 2011-11-21
DE102012000709A DE102012000709A1 (de) 2011-11-21 2012-01-16 Verfahren zum Abwehren einer anfliegenden ballistischen Rakete und Abfangsystem

Publications (3)

Publication Number Publication Date
EP2594891A2 true EP2594891A2 (fr) 2013-05-22
EP2594891A3 EP2594891A3 (fr) 2015-07-08
EP2594891B1 EP2594891B1 (fr) 2016-07-27

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EP12007727.6A Not-in-force EP2594891B1 (fr) 2011-11-21 2012-11-15 Procédé destiné à repousser une fusée balistique se rapprochant en volant et système d'interception

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EP (1) EP2594891B1 (fr)
DE (1) DE102012000709A1 (fr)
IL (1) IL222655A (fr)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN110906795A (zh) * 2019-12-13 2020-03-24 陈忠响 一种中段拦截导弹、系统以及方法
CN111216890A (zh) * 2014-07-28 2020-06-02 英西图公司 用于反击无人机的系统和方法
CN112179213A (zh) * 2018-08-06 2021-01-05 蓝箭航天空间科技股份有限公司 导弹拦截方法、存储器及服务器
CN114727165A (zh) * 2022-06-09 2022-07-08 北京航天驭星科技有限公司 基于遥测数据确定火箭起飞时刻的方法、装置

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IL125455A (en) * 1998-07-22 2003-12-10 Rafael Armament Dev Authority System for destroying enemy ballistic missiles
US6584879B2 (en) * 2001-11-14 2003-07-01 Northrop Grumman Corporation System and method for disabling time critical targets
US8242422B2 (en) * 2009-02-23 2012-08-14 Raytheon Company Modular divert and attitude control system

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
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Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN111216890A (zh) * 2014-07-28 2020-06-02 英西图公司 用于反击无人机的系统和方法
CN111216890B (zh) * 2014-07-28 2024-02-02 英西图公司 用于反击无人机的系统和方法
CN112179213A (zh) * 2018-08-06 2021-01-05 蓝箭航天空间科技股份有限公司 导弹拦截方法、存储器及服务器
CN112179213B (zh) * 2018-08-06 2022-05-27 蓝箭航天空间科技股份有限公司 导弹拦截方法、存储器及服务器
CN110906795A (zh) * 2019-12-13 2020-03-24 陈忠响 一种中段拦截导弹、系统以及方法
CN110906795B (zh) * 2019-12-13 2022-07-05 谢勋 一种中段拦截导弹、系统以及方法
CN114727165A (zh) * 2022-06-09 2022-07-08 北京航天驭星科技有限公司 基于遥测数据确定火箭起飞时刻的方法、装置

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IL222655A (en) 2017-02-28
DE102012000709A1 (de) 2013-05-23
EP2594891A3 (fr) 2015-07-08
EP2594891B1 (fr) 2016-07-27

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