74 results about "Ballistic missile" patented technology

The invention relates to an autonomous integrated navigation system which belongs to the technical field of navigation systems. The SINS (Strapdown Inertial Navigation System) / SAR (Synthetic Aperture Radar) / CNS (Celestial Navigation System) integrated navigation system takes SINS as a main navigation system and SAR and CNS as aided navigation systems and is established by the following steps: firstly, designing SINS / SAR and SINS / CNS navigation sub-filters, calculating to obtain two groups of local optimal estimation values and local optimal error covariance matrixes of the integrated navigation system state, then transmitting the two groups of local optimal estimation values into a main filter by a federal filter technology for fusion to obtain an overall optimal estimation value and an overall optimal error covariance matrix, and finally, performing real-time correction on the error according to the overall optimal estimation value so as to obtain an optimal estimation fusion algorithm of the SINS / SAR / CNS integrated navigation system. The autonomous integrated navigation system, disclosed by the invention, is less in calculation amount and high in reliability, is applicable to aircrafts in near space, aircrafts flying back and forth in the aerospace, aircrafts for carrying ballistic missiles, orbit spacecrafts and the like, and has wide application prospect.

The light-weight handheld less-lethal ballistic projectile launcher is configured as an “over / under” double barrel handheld device with a “break open” loading action. The launcher includes a barrel section pivotally connected to receiver section. The barrel section pivots between an open load / unload position and a closed firing position. The barrel section includes two metal barrel sleeves that are press fit into axial bores formed in the body of the barrel section. The barrel section also includes an ejector, which locks the barrel section in the closed firing position and partially expels spent rounds from the barrel sleeves when the barrel section is opened. The receiver section houses the launcher's fire control mechanism. The fire control mechanism uses a traditional single action operation and ensures that rounds are alternatively discharged from each barrel and prevents rounds from both barrels from being discharged simultaneously.

The invention discloses a trajectory target launch point estimate apparatus based on model updating and a method, and the method comprises the steps of reverse tracking and filtering the point data measured by a radar, calculating the flight path and the launch point position of a ballistic missile through Runge-Kutta method of extrapolation, and updating the estimated launch point position through model deviation. By adopting an insensitive filtering algorithm based on the ballistic missile motion equation, non-linear filtering in high precision is realized; the computational complexity is moderate; and a high cost efficiency ratio is provided; a high value precision through track extrapolation calculating by fourth-order Runge-Kutta method is provided; the estimate deviation of the launch point position caused by an active segment is removed through model deviation updating; and the ballistic missile launch point estimate precision is substantially improved.

The invention provides a rapid evaluation method of the disturbing gravity along a flight trajectory for the first time. The rapid approximation method with ballistic missiles as research objects aims at solving the problem of rapid evaluation on the disturbing gravity in the missile flight process. The rapid evaluation method comprises generating a standard trajectory according to a launch task; generating a flight pipeline with the standard trajectory as a benchmark and performing airspace subdivision; performing evaluation on the node disturbing gravity by a point quality method or a high-order spherical harmonic function method; calculating a disturbing gravity value of any point on an actual trajectory based on a network function approximation theory inside a calculation unit and achieving the whole rapid evaluation calculation on the disturbing gravity along the flight trajectory. Compared with the existing method, the rapid evaluation on the disturbing gravity along any flight trajectory can be implemented, the trajectory calculation requirements for the evaluation precision of the rapid evaluation method are met, the evaluation speed, the data storage and other indexes are excellent, and the real-time evaluation on the disturbing gravity missiles is implemented.

The invention provides a trajectory boosting segment state deviation parsing and calculating method under the action of disturbance gravitation. In particular, a ballistic missile serves as a research object, and the problem of rapidly working out the ballistic missile boosting segment state deviation is solved. The method includes the specific steps that firstly, a standard trajectory is designed and generated according to a launch mission; secondly, the state deviation of each point of the boosting segment is worked out through a state deviation parsing method induced by the perturbation thought; then, the apparent acceleration deviation of the corresponding point is calculated based on the state deviation of each point of the boosting trajectory, the deviation is regarded as a high-order item of disturbance gravitation, and meanwhile compensation is conducted; finally, the state deviation result is amended through a Newton iteration method. By means of the method, fast calculation of the state deviation of the boosting segment of any flight trajectory can be achieved, the calculation accuracy can meet the requirement for trajectory calculation, the calculation method is far superior to that of an existing method, and a foundation is laid for achieving fast mobile lunching of the trajectory.

The invention discloses a finite time integral sliding mode terminal guidance law based on a super-twisting algorithm. The guidance law comprises the following steps: step 1, establishing a three-dimensional target-missile relative motion equation; step 2, designing a guidance law. A stability analysis is conducted on the guidance law. The invention provide the finite time integral sliding mode terminal guidance law based on the super-twisting algorithm by combining an integral sliding mode, a super-twisting algorithm and a finite time disturbance observer. According to the finite time integral sliding mode terminal guidance law based on the super-twisting algorithm, the finite time can be converged, the buffeting phenomenon can be eliminated, so that the maneuvering acceleration of a target can be better estimated, and the guidance accuracy is high; and the guidance law can be used for intercepting high-speed large-maneuvering targets such as unmanned aerial vehicles, ballistic missiles and the like, and has a wide application prospect.

A method for guiding an intercept vehicle to intercept a ballistic target includes conceptual setup of a line-of-sight (LOS) extending between the vehicles. The interceptor is accelerated in a direction perpendicular to the LOS until its velocity in that direction equals that of the target. At this time, the thrust of the interceptor accelerates it along the line-of-sight, thereby guaranteeing an intercept.

A guidance method for a powered ballistic missile involves using an onboard computer to numerically simulate the flight path of the missile in real time, using a model with at least 3 degrees of freedom. The results of this simulation are used to update in real time an aim point and / or a predicted intercept point. An iterative process may be used in adjusting the aim point and / or the predicted intercept point. The process may be carried out until a specified number of steps have been completed, and / or until a specified heading error threshold of the aim point and a specified time of flight threshold have been achieved. The use of real time updating of an aim point of the missile advantageously takes into account variations in missile velocity and position due to individual variations in the rocket motor of the missile.

A method for guiding a ballistic missile to a target has a first mode guidance process that drives a magnitude of a velocity-to-go (Vgo) vector toward zero. On regular intervals, a proportional velocity deficit value is calculated as equal to a time constant (Tau) multiplied by a specific force magnitude (sf). When the magnitude of the Vgo vector has been driven to less than or equal to the proportional velocity deficit value by way of the first mode guidance process, a second mode guidance process is initiated. The second mode guidance process constrains the magnitude of the Vgo vector to be equal to the proportional velocity deficit value throughout the remaining portion of powered flight.

The invention discloses an air-to-surface missile control full-flow simulation system, which relates to the technical field of ballistic missile simulation. The air-to-surface missile control full-flow simulation system of the invention comprises a carrier simulator, a high-speed real-time simulation computer, a missile-borne computer, an integrated controller, a steering gear and a power supply.The communication network is established between various parts through CAN bus and 1553B bus. The carrier simulator simulates the real carrier, controls the launching process, sends the carrier navigation information and monitors the flight data. The high-speed real-time simulation computer simulates air-to-surface missile inertial unit, and gives air-to-surface missile navigation information; themissile-borne computer uses aircraft navigation information and air-to-surface missile navigation information to complete the closed-loop solution of navigation and guidance law. The invention realizes the complete verification of the whole work flow, the whole logic chain, the whole data stream and the whole algorithm domain of the air-to-surface missile. The invention also discloses an air-to-surface missile control whole process simulation method, wherein the trajectory generation is closer to the flight process, and the simulation test is more reliable.

The invention provides an analyzing method of free phase ballistic error propagation under nonspherical perturbation effect of the earth. For a typical ballistic missile, firstly, under a condition of considering the nonspherical perturbation effect of the earth, a high precisely simplified model applicable to free phase motion analysis and calculation is built; then according to small error assumption and linear system theories, an equiangular / equal geocentric distance / isochronous propagation model and the state transition matrix analytical solution of the equiangular / equal geocentric distance / isochronous propagation model are exported; further through decoupling lengthways and lateral propagation models, lateral errors resulting from higher order terms are corrected based on a least square fitting method; and the balance between the speediness and the accuracy in calculating the free phase ballistic characteristic under nonspherical perturbation effect of the earth is realized. The method provided by the invention has the advantages of rapid calculation speed, wide application area and accurate correction model and satisfying demands for rapid mobile lunch and on-missile real time calculation.

A hypervelocity impact detection method and system for determining the precise impact location in a detection surface, of impacts such as ballistic missile intercepts, micrometeoroids and orbital debris (MMOD) or other shock events, utilizes a gridless detection surface capable of propagating radio frequency (RF) impact detection signals responsive to receiving hypervelocity impacts from objects, and multiple sensors on the detection surface that directly measure radio frequency RF emissions generated by the hypervelocity impacts on the surface, and a time of arrival (TOA) position measurement technique for determining the precise impact location in the detection surface.

The invention provides an improved active section trajectory estimation algorithm. The method comprises the steps of calculating a thrust acceleration template according to the initial mass of a missile target, the air injection speeds of all stages of engines, the shutdown time and the like; Widening the thrust direction angle into a state vector component, and establishing a fine parameterized dynamic model by utilizing the widened state vector and a thrust acceleration template; Establishing an early warning satellite detection model according to the position vectors of the missile target and the two early warning satellites in the earth fixed connection coordinate system and the coordinate transformation matrix from the earth fixed connection coordinate system to each early warning satellite orbit coordinate system; Estimating a missile target initial state vector and an initial state covariance matrix according to the first three groups of equivalent detection positions; Describing a thrust direction angle by utilizing a first-order Markov process, and constructing a process noise matrix adaptive to thrust direction angle change; And performing filtering processing by using aUKF filtering algorithm, and estimating the motion state of the missile target. The method can improve the precision of trajectory estimation of the ballistic missile detected by the early warning satellite.

The invention discloses a multi-satellite coordinated real-time tracking method for a spatial dynamic target. The method comprises the steps of firstly establishing a low earth orbit satellite and staring detector model, a tracking task model and a visible time window model; then introducing a rolling time domain concept when real-time tracking is performed on the spatial dynamic target, taking the forward-looking fixed time length T as the rolling time domain to obtain a tracking task of a ballistic missile in the rolling time domain, and establishing a single-layer to-be-tracked task set with the single rolling time domain as a reference, and splitting each to-be-tracked task into a plurality of atomic tasks in the rolling time domain by use of a reverse tree structure to obtain a set ofdetector resources corresponding to each atomic task; and finally, constructing a fitness function based on constraints and objective functions of resource scheduling, and then solving the optimal particle individual based on the fitness value of the particle individual through an evolutionary particle swarm optimization algorithm with crossover and mutation operations, and performing resource scheduling of the detector resources on the single-layer atomic task set according to the optimal particle individual.

There is provided an interception system that employs miniature kill vehicles. According to certain embodiments of the invention, a small kill-vehicle (KV) accommodated within an interceptor missile together with additional at least one KV, each KV is equipped with an image acquisition sensor operable in the visible or near IR range, and adapted to be launched from the interceptor missile during exo-atmospheric flight and utilizing said sensor for tracking and homing at least one Re-entry Vehicle (RV) of a Ballistic Missile (BM).

In an aiming or sighting device and an aiming method for determining the impact point of a ballistic flying body such as a rocket or a tube weapon projectile, specific characteristic values of the ballistic flying body are stored in a memory and an evaluating stage produces control signals in response to the specific characteristic values and in response to actual supplied system parameters. The control signals are transmitted to a display and to an adjustment drive. A model of the trajectory of the ballistic flying body is produced on the basis of all respective possible actual system parameters in an evaluating stage, whereby the trajectory is subdivided into at least two phases, each with a respective submodel. The submodel of the first phase is thereby a model with three or six degrees of freedom and the submodel of the second phase is a model with three degrees of freedom. An input unit permits adjusting the mode of operation and the correction of the impact location.

The invention discloses a missile-borne BDS / SINS deep integration self-adaptive CKF filtering method. The method comprises the following steps: firstly, setting a flight path of a ballistic missile and generating IMU data by a path generator; secondly, generating instant medium-frequency data by using the flight path of the ballistic missile through a satellite signal simulator, providing the instant medium-frequency data to a software receiver for satellite navigation solution and carrying out strap-down inertial navigation solution on the generated IMU data; thirdly, establishing a BDS / SINSdeep integration navigation system model under an emission inertial system; finally, introducing a multiple suboptimal fading factor into a standard volume Kalman filter to form a self-adaptive volumeKalman filter; and outputting correction information of the system through optimal estimation of the self-adaptive volume Kalman filter. The missile-borne BDS / SINS deep integration self-adaptive CKFfiltering method disclosed by the invention has the advantages that a better tracking effect on the rapidly-changed system state is realized, and navigation precision and robustness of the BDS / SINS deep integration navigation system in the high dynamic and strong interference environment are improved.

The invention relates to a three-diversity anti-interference satellite navigation receiver, which comprises an anti-interference antenna group, a three-diversity satellite navigation receiver, a satellite signal capturing / tracking unit and a positioning information fusion unit, and is characterized in that the three-diversity satellite navigation receiver is used for receiving three paths of satellite navigation signals of the anti-interference antenna group; the satellite signal capturing / tracking unit is used for performing satellite navigation signal capturing on the three paths of satellite navigation signals, and acquiring original observation information of all satellites of each path of navigation signals; and the positioning information fusion unit is used for fusing the original observation information of all satellites of the three paths of signals, and positioning and calculating positioning information according to the original observation information of the satellites. The three-diversity anti-interference satellite navigation receiver receives signals outputted by three anti-interference satellite antennas simultaneously and performs information fusion on the three signals, so that use requirements of a ballistic missile, especially a whole-course maneuvering ballistic missile, in complex electromagnetic environmental conditions are met.

The invention discloses a ballistic missile three-dimensional trajectory estimation method based on an infrared early warning image, and the method comprises the steps: carrying out the point target detection of a pre-obtained missile infrared remote sensing image in a continuous time sequence, and extracting the pixel coordinates of the mass center of a target of each image point; according to the extracted time and space information of the target point, tracking the target and obtaining a motion trail; conducting ballistic missile trajectory screening on ballistic missile motion characteristic analysis; unifying the attitude data and the orbit data of the missile target in a coordinate system; predicting a target three-dimensional trajectory through a collinear equation and a constraintcondition in a z direction according to the image plane two-dimensional trajectory; and correcting the error track according to the kinematic model of the missile and the constraint condition of the active section of the missile. According to the method, three dimensions are estimated from two dimensions, and a new scheme is provided for tracking and early warning of missiles.

The invention discloses a ballistic missile boosting section tracking method based on dimension-augmenting shifted Rayleigh filtering, and belongs to the technical field of boosting section tracking of a ballistic missile, in order to solve the problem that in the existing boosting section tracking method for the ballistic missile, because the axial acceleration and quality consumption velocity are difficult to amend, and the tracking is invalid. The observability problem of the boosting section tracking of the ballistic missile is solved thoroughly by adopting dual-star navigation / following constellation configuration; secondly, a driving section ballistic missile real model is drawn close by adopting an interactive multi-mode method, and the accurate modeling problem is solved; and finally a modified centralized shifted Rayleigh filtering method is deduced to achieve driving section continuous tracking, and the filtering gain is subjected to self-adaptive adjustment by introducing UT (ultrasonic test) conversion on the basis so as to solve the problem that the traditional shifted Rayleigh filtering can not amend the asymmetrical state of a platform and a target. The ballistic missile boosting section tracking method provided by the invention is suitable for boosting section tracking of the ballistic missile.

The invention discloses a heterogeneous distributed detection information target identification optimization method based on threat assessment, which is applied to an attack and defense countermeasuresystem simulation system. The method is characterized in that a heterogeneous distributed sensor network is formed based on signal-level semi-physical systems such as multiple satellites and multipleradars, the property of an attacking target is evaluated to form a threat sequence, and a ground command control system schedules the satellites and the radars to preferentially track and identify ballistic missile targets with high threat values according to the threat values in the threat sequence, and finally an interception or striking sequence is determined according to a target identification result. According to the method, redundant and complementary information fusion is carried out on collected detection information, target and environment information is collected and processed to agreater extent, accuracy and reliability of battlefield target identification are improved, and therefore the method is of great significance in subsequent situation evaluation, threat estimation andinterception strategy formulation, and the winning probability in confrontation simulation is greatly improved.

A method for identification of one or more launched objects obscured by debris objects within a debris field comprises: directing one or more sensor pulses at the debris field to obtain a plurality of sensor images; identifying objects within the debris field based on the sensor images; determining acceleration characteristics for each of the identified objects within the debris field based on the sensor images; identifying objects exhibiting free fall acceleration characteristics as debris objects; and identifying objects exhibiting centripetal acceleration characteristics as the one or more launched objects.

The invention discloses a method for automatically distributing accuracy indexes of a ballistic missile inertia measurement system. The method includes solving drop point accuracy values and drop point deviation proportions of error coefficients according to set environment function values and initial values of the inertia measurement system error coefficients, and comparing the calculated drop point accuracy values with drop point accuracy threshold values; if the calculated drop point accuracy values are smaller than the set threshold values, not needing to adjust; if not, searching to-be-adjusted error coefficients according to the drop point deviation proportions, performing proportion adjustment, solving the drop point accuracy values and the drop point deviation proportions of the error coefficients after adjustment, and repeating the steps to finally acquire inertia measurement error coefficients meeting requirements of the drop point accuracy threshold values. Compared with other inertia measurement system accuracy index methods, the method has the advantages that automatic distribution of the accuracy indexes of the inertia measurement system and optimal adjustment of the error coefficients are achieved.

The invention specifically discloses an early warning image ballistic missile direction estimation method under a single-satellite condition. The method comprises the following steps: S1, obtaining an original image of a target missile, and carrying out target detection on the obtained original image at a known time interval to extract a centroid coordinate of the target missile in the corresponding original image; s2, performing target tracking according to the center-of-mass coordinates of the target missile and the corresponding time information so as to obtain the motion trail of the target missile; s3, converting the motion trail attitude data of the target missile into an earth coordinate system, and finding out a corresponding trajectory point in the earth coordinate system; s4, establishing a normal vector equation set corresponding to the ballistic trajectory plane according to the ballistic trajectory points; and S5, establishing a coordinate equation of the target missile by using the collinear equation, and solving the missile direction according to the coordinate equation and the normal vector equation set. According to the invention, an accurate result can be obtained without ballistic prior knowledge, and the method has the characteristics that the required data is easy to obtain, the calculation amount is small and the calculation time is short.

The invention provides a closed-circuit guidance on-line compensation method considering any-order earth non-spherical gravitational perturbation. The method comprises the following steps that 1, a disturbance gravity reconstruction model is off-line constructed along a trajectory; 2, a required speed vR is calculated on the basis of a binary dynamics model in each guidance period of a missile; 3,the shutdown point state predicted on the missile in real time is taken as an input condition, and a trajectory terminal state deviation delta Xc is rapidly calculated by adopting a trajectory deviation analysis and forecast model taking consideration of the earth non-spherical gravitational perturbation; 4, a required speed correction amount delta vR is analyzed and calculated according to the trajectory terminal state deviation delta Xc; and 5, a control quantity U considering the earth non-spherical gravitational perturbation is solved. The brand-new closed-circuit guidance on-line compensation method is provided on the basis of the trajectory deviation analysis and forecast model of a missile trajectory free section, and a method support is provided for further improving the precisionof the strategic missile guidance method in China.

The invention discloses an optimal data fusion method suitable for a ballistic missile INS / CNS / GNSS integrated navigation system. The method comprises the steps of constructing an INS / CNS / GNSS integrated navigation system model; respectively introducing an adaptive fading factor and a maximum correlation entropy criterion in a time updating stage and a measurement updating stage of a generalized high-order CKF to carry out local state estimation on an INS / CNS subsystem and an INS / GNSS subsystem; and fusing local estimation of the INS / CNS subsystem and the INS / GNSS subsystem according to a minimum variance principle and a volume criterion to obtain global optimal state estimation. According to the method, the influence of process modeling errors and non-Gaussian measurement noise on state estimation can be suppressed at the same time, the adaptability and robustness of ballistic missile INS / CNS / GNSS integrated navigation are improved, and a globally optimal state estimation value is obtained.

The invention discloses a method for estimating complex micro-motion space cone target parameters by using a deep learning convolutional neural network. The method comprises the following steps of: firstly, a ballistic missile target geometric model is established; then, a single-frequency pulse is emitted to a target; echoes of a target are received, time-frequency analysis is performed on the target echo to obtain a time-frequency distribution diagram of the target; the time-frequency graph is used as the input of the CNN, so that the network learns the features of the time-frequency graph;the cone target height, and estimated values of a bottom surface radius and a precession angle are finally obtained; the method for estimating the space cone target comprises the following steps of: estimating a micro Doppler frequency curve of the target from a time-frequency diagram; the method has poor stability, the estimated micro Doppler frequency has great influence on a parameter estimation result; according to the method provided by the invention, the features are directly extracted from the time-frequency diagram, the estimation stability and precision are improved, the previous method can only estimate the precession target, and the method provided by the invention not only can estimate the precession space cone target, but also can estimate the parameters of the nutation spacecone target.

The invention belongs to the space object identification field category, and especially relates to a novel feature extraction technology applied to space object identification; the method comprises the following steps: using a radar to measure and obtain RCS data; parsing chaotic features of the RCS data, including determining an embedding dimension and time delay; using the estimated embedding dimension and time delay to reconstruct a phase space; calculating association integration and parsing so as to determine the object state stability; counting the RCS data switching frequency if the object state is stable; finally, using the state switching frequency to calculate, thus obtaining features reflecting the ballistic missile object flight state. The method can well reflect the true object, fragment and missile body motion states, thus identifying true objects.

The invention discloses a material level prediction method for a ball mill. The method comprises the following steps that: collecting a vibration signal on the cylinder wall or the bearing of the ballmill and the sound vibration signals of microphones around the ball mill; inputting the vibration signal and a sound vibration signal into a data collection card through a conditioning circuit and ananalog-to-digital conversion circuit; through an MFCC (Mel-Frequency Cepstral Coefficients, independently calculating the feature parameters of the vibration signal and the sound vibration signal inthe data collection card; re-combining the vibration signal with the sound vibration signal; collecting the re-recombined signal by the data collection card; and finally, predicting the material levelof a corresponding signal for data in the data collection card through a TBM (Tactical Ballistic Missile) model established by historical data. By use of the method, the material level of the ball mill can be effectively predicted, and the method has the advantages of high material level prediction accuracy, small fluctuation and good practical value and popularization prospect.