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.

The invention provides an operational capability estimation method for an air-defense anti-missile system, and the method comprises the steps: decomposing an operational activity based on an operational mission, and extracting corresponding evaluation indexes; establishing an operational capability evaluation model based on the evaluation indexes, obtaining quantitative index data of the to-be-evaluated anti-aircraft anti-missile system based on the operational capability evaluation model, and constructing a corresponding capability requirement satisfaction degree evaluation matrix; estimatingthe comprehensive operational capability of the to-be-evaluated anti-aircraft anti-missile system through employing weighted aggregation operation based on the comprehensive weight of each of the evaluation indexes and the capability requirement satisfaction degree evaluation matrix corresponding to the quantitative index data. The method can effectively reduce the subjectivity of an evaluation process through building the complete, reliable and feasible air-defense anti-missile system while making the better use of the experience of an expert group, and improves the effectiveness and the reliability of an evaluation result.

In a first aspect, the invention uses the thrust and axial aerodynamic forces of and on the missile system as the primary mechanism for altering the missile system's flight path as opposed to the missile system's aerodynamic normal force. In accordance with this first aspect, a first embodiment is a method for changing the direction of travel of a vehicle moving through a fluid medium to a desired direction of travel measured in the vehicle's inertial frame of reference, the vehicle comprising a body including a nose, a moment generating control device, and an axial propulsive device. The method comprises actuating the moment generating control device to align the nose substantially with the desired direction of travel, and applying the axial propulsive device as the primary mechanism to redirect the vehicle's velocity to the desired direction of travel. In a second aspect, the invention provides a missile control system implementing the first aspect of the invention. In a third aspect, the invention simulates in software a missile control system implementing the first aspect of the invention.

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.

A specially design missile system use for hindering the centrifugal forces inside vortexes commonly known as tornados. The missile system composed of a typical modified vehicle with a design platform. The platform is design of the insides of a platform that contains liquid and compressed gas. A command and control module and a projectile missile system complement above the platform. When a super cell appears and probable starting of tornado activity commences the Thermal Energy Radiance Expander system is send over to the targeted area. As the initial start of the tornado commences the vehicle and whole system is sent to the target area. The missile is then loaded with the gaseous element turn live and launch on to the target. The missile control through the command module explodes intentionally at the side of the vortex. The aim of the system is to break and hinder further activity of the vortex hence terminating tornado.

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 invention relates to the field of aviation experiments, in particular to a flutter model of a plane missile system. The flutter model solves the problems that at present, pitch and yaw mode coupling of a flutter model is not prone to being identified, and rigidity adjustment is inconvenient. According to the flutter model of the plane missile system, rigidity of a missile model in the three directions is decoupled, and identification of a side flat mode, a yaw mode and a pitch mode in an experiment is facilitated; in addition, three replaceable leaf springs which are correspondingly arranged simulate frequencies of the side flat mode, the yaw mode and the pitch mode of a missile respectively; the leaf springs can be replaced so as to adjust the frequency of a certain mode of the missile model, the frequency adjustment on the missile model in the experiment can be facilitated, experimental expenses are saved, and the experimental time is shortened.

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 relates to a missile comprehensive test simulation system, which can carry out pre-installation test, post-installation test and flight simulation for semi-active laser air-to-ground missiles and missile-borne equipment, so as to check the coordination and correctness of the work of the whole missile and missile systems. It is characterized in that it includes two parts: a computer part and an electromechanical part, and the computer part includes four parts: a test machine, a simulator, a board card and a power supply unit; the electromechanical part includes a missile supporting mechanism, a target moving mechanism, a turntable and a laser target simulator and other auxiliary equipment. The system has the function of automatic batch test, the test items are optional, the format of the test data is unified and standardized, and the curve can be dynamically displayed along with the test process. The system is an independent box structure, and the missile support mechanism and target movement mechanism can be disassembled, which is convenient for transportation and field operation.

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 provides a manufacturing method of a ground-interception missile defense system and application of the missile defense system. The manufacturing method of the ground-interception missile defense system is characterized in that the base and the inner layer of the ground-interception missile defense system are cast by use of an alloy steel, a plurality of layers are formed for mounting an outer layer, a second layer is made of a composite bullet resistant material composed of a soft bullet resistant material and a shock-absorptive material in addition to the alloy steel, a third layer is knitted by use of carbon fibers and a c/c composite material, an outer layer is coiled by use of a high-strength alloy steel, welded by use of steel hoops at intervals and hooped section by section, the joint of each layer is hooped by use of one steel hoop and fastened by use of a screw button, each layer is spaced with a fire-resistant antiknock glue, a groove is formed in the inner layer of a device head, sleeves a connection and is tightened by use of a screw, an outer opening is in a flaring shape, and an additional interception protective layer is mounted in the device; the ground-interception missile defense system is mounted on a heavy traction vehicle; a vehicle-mounted radar is combined with optical detection to realize tracking of a defense penetration missile and to transmit inducing signals; a vehicle-mounted navigation system is used for realizing tracking the defense penetration missile along with a phased array radar, and correcting and commanding the traction vehicle to move by use of alignment and calibration so that the missile is capable of falling into the device.

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.

The invention discloses an air-ground missile hit precision evaluation method based on an error spectrum, and the method comprises the following steps: S1, carrying out the preprocessing of the air-ground missile hit precision data, so as to expand the sample size, and enabling the data to be converted into a relative distance error; S2, determining a key measurement index which can be used for evaluating the hitting precision of the air-ground missile system; S3, constructing an air-ground missile hitting precision attribute matrix; s4, establishing an air-ground missile hitting precision attribute competition matrix according to the air-ground missile hitting precision attribute matrix; and S5, solving the feature vector of the air-ground missile hit precision attribute competition matrix, and giving an air-ground missile hit precision evaluation result corresponding to the element value according to the element value in the feature vector. According to the method, the evaluation index is steady, and in addition, the hitting precision of the air-ground missiles of the same type can be evaluated, and the hitting precision of the air-ground missiles of different types can also be evaluated.

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.

The invention discloses an ablation-resistant composite material launch canister structure and a forming method thereof and belongs to the technical field of missile system parts. The ablation-resistant composite material launch canister structure comprises a mould, an inner surface abrasion-resistant layer and an outer surface structure layer. The inner surface ablation-resistant layer wraps theouter surface of the mould, and the outer surface structure layer wraps the outer surface of the inner surface ablation-resistant layer. According to the structure and the forming method thereof, thetechnical problems that in the prior art, the canister structure adopts a mould and outer surface structure layer mode, when the canister structure is used, a canister is prone to damage after being subjected to high-temperature ablation, the number of replacement times is large, and the canister use cost is high are solved; and the inner surface ablation-resistant layer is wrapped between the mould and the outer surface structure layer, the ablation property of high-temperature gas to the canister structure is improved, the number of repeated use times of the canister is increased, and the canister launch cost is lowered.