83 results about "Missile system" patented technology

The disclosed system, device and method for an anti-missile system generally includes a ground-based sensor array generating tracking data of a guided missile tracking a target. A control node in communication with the ground-based sensor array generates targeting data from the tracking data. A phased array directed-energy unit in communication with the control node radiates the guided missile with microwave radiation based on the targeting data received from the control node, where the microwave radiation disrupts an electronic component of the guided missile such that the guided missile discontinues tracking the target.

A system and method for locating and suppressing a forest fire in the beginning stage comprising a fire location means such as one or more unmanned drones, a processing and control means configured to control the drones, to process location and other information about the forest fire, to develop a fire suppression proposal and to control a delivery means, a delivery means such as a guided missile system, and a fire suppression means such as a missile containing fire retarding chemicals.

A missile, such as a cruise missile, has a nose payload portion having a frangible nose cover and a relatively hard target penetration nose cone. The nose cone may have a liquid fuel tank within, and a chemical energy explosive charge, such as a shaped charge, aft of the liquid fuel tank. The target penetration nose cone enables perforation of certain types of targets prior to detonation of the chemical energy explosive and the liquid fuel. The frangible nose cover is configured to be easily perforated or otherwise removed by the explosive force of the chemical energy explosive charge when the missile system is utilized for the attack of hard targets. The nose payload portion may have a fragmentation case, with one or more features designed to enhance fragmentation during detonation of the explosive and/or the liquid fuel.

The invention discloses a high-density floated gyro strap-down inertial measurement unit, which comprises a shell bracket (1). The shell bracket (1) is provided with X, Y and Z gyroscope mounting planes and three accelerometer mounting planes respectively; the gyroscope mounting planes are provided with floated gyroscopes; the accelerometer mounting planes are provided with quartz flexible accelerometers; the shell bracket (1) is also provided with a secondary combination power supply (2) and a circuit board; the secondary combination power supply (2), the circuit board, the floated gyroscopes and the quartz flexible accelerometers are connected through internal cables to form a measurement circuit; and the shell bracket (1) is in sealing connection with a cover board (3) through a bolt. The invention provides the high-density floated gyro strap-down inertial measurement unit to meet the requirements of a tactical missile system on small volume and good anti-vibration property of the strap-down inertial measurement device; and the unit also has the characteristics of light weight, simple structure and convenient processing.

A system (10) for determining the range between a missile (14) and a target (18) adapted for use with a semi-active missile system. The system (10) includes a first circuit (12) for generating a periodic signal (24) that is periodically frequency modulated. A second circuit (16) determines a closing rate at which the missile 14 is approaching the target 18 via the periodic signal 24. A third circuit (16) determines a value containing information corresponding to the range and the closing rate via the periodic signal. A fourth circuit (16) determines the range from the closing rate and the value. In a specific embodiment, the first circuit (12) includes an illumination system (12). The illumination system (12) includes a periodically modulated carrier signal generator (32) that generates the periodic signal (24). The periodically modulated carrier signal generator (32) includes a frequency source, a frequency modulator, and an illumination system computer. The illumination system computer runs software for adjusting the modulation parameters of the frequency modulator. The second, third and fourth circuits (16) are included in a receiver system (16) onboard the missile (14). The receiver system (16) includes a front receiver located near the front of the missile and a rear receiver located near the rear of the missile (14). A receiver system computer runs software that implements the fourth circuit (16). The receiver system (16) includes a local oscillator for providing a reference frequency for the receiver system. The local oscillator derives the reference frequency from the periodic signal provided by the first circuit. In illustrative embodiment, the fourth circuit (16) runs computer software that subtracts the closing rate from the combination of closing rate and range and provides the range in response thereto.

A fire retarding missile and method of use having a unitary construction and containing a propulsion system as well as fire treatment materials. Dispersal of the materials can be initiated both actively and passively, with passive dispersal allowing for a fail safe mode of operation. The fire treatment materials are of the oxygen reduction type, and the method of the invention is to target the hot spot of a fire in order to reduce the spread and intensity thereof. An active guidance system may be of the heat seeking type, or may alternatively be remote controlled video. Stabilizer and guidance fins are controlled in response to signals from the guidance system in order to precisely position the device. The device can be modified for use in fires of specific heat ranges by using nose cones designed to melt in a respective heat range.

Disclosed is a separation device of an ejector motor for a portable missile, capable of being separated from a missile without any additional separation devices after ejecting the missile. An ejection rocket motor and a separation device are integrally formed. Ejection is performed by a thrust generated while the ejection rocket motor is combusted, and separation is performed by cutting off shearing bolts with using a force generated from a separation cylinder when the combustion of the ejection rocket motor has been completed. This operation and configuration may minimize a weight and a space occupied by a general ejection and separation device in a missile system.

A missile system has target detection and target destruction facilities interlinked by digital communication lines, the target detection facilities including a sonar and a digital receiver for space intelligence radio signals, wherein the sonar and the digital receiver for space intelligence radio signals are mounted on a self-elevating buoy placed inside a sealed radio-transparent case.

The invention discloses a guidance law for simultaneously attacking a static target by multiple missiles based on a proportional guidance law. The method comprises the steps that when n missiles simultaneously attack the same static target, r, V and phi of the i missile are obtained; then the consistency deviation xi<t,i> of the n missiles in the remaining flight time is obtained; whether theconsistency deviation xi<t,i>=0 is true or not is judged, and if yes, the current flight state of each missile is maintained; if not, the current consistency deviation xi<t,i> is input into the guidance law of fixed time convergence, and tangential control input a<n,i> and normal control input a<t,i> of a multi-missile system are obtained; and a missile kinetic model is established, r, V and phi at the next moment are obtained through the missile kinetic model, and the initial step is returned to. According to the method, by adding additional items, the situation that the convergence completion time of the system tends to be infinite due to the fact that the included angle between the missile flight direction and the missile visual line is 0 before the multi-missile system completes stateconsistency is avoided.

The present invention relates to electro-optic guided missile systems and, in particular, it concerns systems and methods providing enhanced navigation capabilities based on ego-motion processing of seeker images. The missile system comprising: a missile; a seeker located at a nose portion of said missile, said seeker comprising an electro-optic imaging sensor; and a control arrangement for steering the missile along a flight path to a target, characterized in that the missile system further comprises: a navigation subsystem receiving images from said imaging sensor, said navigation subsystem being configured to: co-process a plurality of said images from said imaging sensor to derive ego-motion of said missile relative to a region viewed by said imaging sensor; derive from said ego-motion a calculated target direction from said missile to a target.

The invention belongs to the technical field of missile weaponry launching control, and particularly relates to a safely and credibly launching control method for a missile. The safely and credibly launching control method for the missile particularly relates to a starting operation control method achieved for missile launching preparation software based on safety control methods of identity authentication, real-time authorization and the like. A launching control computer is revised, a safety control module is designed and added, and a set of brand new authentication, identification and starting control process is designed through an external interface, so that reliable and effective missile launching control is achieved. Above all, the strict safety control process for starting and software operation of an embedded type system is designed, so that launching of the missile can be safely and effectively controlled. A severe result possibly caused by misoperations or deliberate sabotage is avoided while reliably launching of the core missile system is ensured, and therefore safety and reliability of the core missile system are ensured.

The invention discloses a remote test system applied to missile ground joint debugging and a test method thereof. According to the remote test system and the test method, different test subsystems areadopted, or the same integrated test system is adopted; the problem of joint debugging of scientific research units in different places is solved, remote semi-physical simulation is supported, and additionally, remote missile tests including an integrated system power supply test, a fuse interface test and a communication interface reliability test are synchronously completed; the current clock information of a remote test terminal is corrected according to satellite timing information, so that a global synchronous clock is obtained, the clock synchronization of the information is controlled,and the missile development test efficiency is improved; and the power consumption of to-be-tested equipment is sampled and then is fed back to a to-be-tested thermal battery, and the real-time powerinformation of a direct-current load simulator of the remote test terminal is adjusted to be converted into load of the thermal battery on a to-be-tested missile, so that the thermal battery on the missile can be closed to the missile-borne system test, and the complete closed-loop test of the missile system can be realized.

The invention discloses an evaluation method for efficiency indexes of a combat mission finished by an antitank guided missible system. The evaluation method comprises the following steps: 1, building an availability model of the antitank guided missile system; 2, building a credibility model of the antitank guided missile system; 3, building a fighting ability model of the antitank guided missile system; 4, building an evaluation model for the efficiency indexes of the combat mission finished by the antitank guided missile system by means of the three models built by the three steps; 5, solving the availability model and the credibility model, calculating the fighting ability model, and finally bringing into the evaluation model to obtain the efficiency indexes of the combat mission finished by the antitank guided missile system. According to the evaluation method disclosed by the invention, the evaluation model with fusion of multiple algorithms is obtained by combining the characteristics of an antitank guided missile and fusing the multiple algorithms, the defects that a single algorithm is too large in calculation amount, unfavorable for comparison and incapable of evaluating the special fighting ability and the like are overcome, and the evaluation science is improved.

A missile system wherein a missile carrier delivers a plurality of submissiles over a target area. A passive or active scanner on the missile carrier repeatedly scans the target area as the missile carrier descends by drag brake over the target area. Submissile assignment to identified targets is performed at a selected altitude. A data processor controls the launching of the submissiles and actively guides the submissiles for substantial portions of their flights via optical fiber or wire links separately connecting each submissile to the missile carrier. Each submissile includes either passive or active identification devices enabling the data processor to track the submissiles individually. A missile assignment table stores a correspondence between identified targets and submissiles assigned thereto. Submissile assignment is effected on the basis of the strength of target signature. Alternatively, the most closely spaced group of targets equal in number to the number of submissiles will be assigned submissiles.

A leveling system for a missile system includes a holder defining an xz-plane. Four flexure plate accelerometers are coupled to the holder at angles such that they are sensitive to movement in an xy-plane. The accelerometers all generate signals in response to movements of the flexure plates. An actuator activates an object control device in response to a platform control signal. A processor receives the accelerometer signals and generates the platform control signal in response thereto.