227 results about "Line-of-sight" patented technology

A collaborative engagement system comprises: at least two unmanned vehicles comprising an unmanned air vehicle including sensors configured to locate a target and an unmanned ground vehicle including sensors configured to locate and track a target; and a controller facilitating control of, and communication and exchange of data to and among the unmanned vehicles, the controller facilitating data exchange via a common protocol. The collaborative engagement system controls the unmanned vehicles to maintain line-of-sight between a predetermined target and at least one of the unmanned vehicles.

A gimbal system, including components and methods of use, configured to track moving targets. More specifically, the disclosed system may be configured to orient and maintain the line of sight (“los”) of the gimbal system toward a target, as the target and, in some cases, the platform supporting the gimbal system are moving. The system also may be configured to calculate an estimated target velocity based on user input, and to compute subsequent target positions from previous positions by integrating the estimated target velocity over time. A variety of filters and other mechanisms may be used to enable a user to input information regarding a target velocity into a gimbal controller.

A firearm sighting system includes a rear sight unit having a plurality of separate sight elements adaptable for “zeroing” the same gun with different ammunition. In preferred embodiments, a rear sight unit comprises windage and elevation adjustments for each of two sight elements, so that multiple ammunition types having different trajectories may be fired accurately from a single firearm after zeroing-in one of the plurality of sight elements for each of the different ammunition types. Preferably, the separate sight elements may be connected to each other or to a common pivot arm or movable bracket so that moving one sight element into the sight path automatically removes the other from the line of vision. Elevation adjustments may be done in various ways, for example, by sliding sight elements out along an arm or bracket, or by changing an angle of the arm or bracket relative to the firearm. The preferred sighting system also includes an elevation-adjustable front sight unit, which may act as an extension member for gross adjustment of the front end of the firearm by significantly lowering the barrel position for a given line of sight between the user's eye, the selected rear sight, and the front sight. Alternatively, the rear sight unit multiple sight elements may comprise only one that is windage adjustable and elevation adjustable, which rear sight unit may cooperate with a front sight unit that has one sight element that is windage adjustable and elevation adjustable.

A low orbit optical imaging satellite has a long thin satellite body housing an optical telescope arrangement. A major part of the telescope arrangement has its optical axis roughly parallel to the direction of elongation and includes a mirror arrangement deployed to direct a line of sight of the optical telescope out sideways from the direction of elongation. The transverse dimensions of the satellite body are preferably minimized to be close to the optical aperture dimension of the optical telescope, thereby providing a high ballistic coefficient and high orbit lifetime for orbits in the low thermosphere.

A system and method that guides a device to an object using periodically adaptive guidance. The guidance and control system creates a reconstructed line-of-sight (LOS) vector to avoid system destabilization associated with small angle approximations during high bore sight engagements. The guidance system adaptively estimates the periodic maneuver of evasive objects with a set of harmonically balanced Kalman filter banks. The Harmonically Balanced Kalman filter banks generate a set of probabilities that weight the effect of each individual Kalman filters on a resultant guidance command signal. The guidance command signal generated by the system acts perpendicular to the object LOS. The guidance and control system uses vectored proportional navigation guidance laws, optimal proportion navigation laws and periodically adaptive augmented guidance laws to generate a guidance command signal to supply to an autopilot.

The invention discloses an optic axis and inertia axis superposition-based geographic tracking method which comprises the steps of computing the space pose of a carrier aiming line in real time by using a space two-point geometric algorithm according to the positioning data and the self position information of a target, computing the space angle of the aiming line relative to a carrier coordinate system by being combined with the carrier space pose and the space coordinate transformation principle, sending the angle to a servo control unit, and adjusting corresponding space pose of a turntable by the servo control unit so as to complete the geographic tracking to the target. The method is different from the image tracking method in terms of theory, therefore, the method is not influenced by the cloud, the mist and the obstacles; and when being used on aircrafts, the method is capable of avoiding the problem of losing the target due to strenuous vibration or sight shielding, so that the scouting and searching capabilities of the battlefield weapon systems are remarkably improved. The method can also be used for guide devices of various guided missiles, so as to improve the accurate attacking capability to ground targets.

The invention discloses an attack angle and attack time control method based on trajectory planning. Firstly, a planned missile flying trajectory is assumed to be composed of an arc section SA and a line section AT, the arc section SA has a radius of R, and the line section AT passes across a target point T(xT, yT) and the requirement that the tail attack angle of the missile is thetaf is met; the arc section SA starts at the initial position S (x0, y0) of the missile and is tangent to the line section AT at a to-be-determined point A(xa, ya), and the point A(xa, ya) is obtained by specified missile attack time tf through an iterative algorithm. On the basis of trajectory planning, the invention provides a trajectory tracking control algorithm based on a virtual target; a feedforward and feedback composite control scheme is put forward in the arc section, and a sight line angle PD control scheme is put forward in the line section; and the method has a small calculation amount, and engineering implementation is facilitated.

The invention discloses a guidance method for attack time control. The guidance method comprises the following steps of establishing mathematical models of a missile and a target by considering the three-dimensional space of motion of the missile and the target; obtaining a dynamic equation of the trajectory inclination angle and the trajectory deflection angle of the missile, and obtaining a dynamic equation of the distance between the missile and the target, the angle of sight of the missile and the deflection angle of sight of the missile; on the basis, establishing a dynamic equation of the attack time error of the missile; then, separately designing the guidance law of the missile according to a pitching channel and a yaw channel, wherein the guidance law of the pitching channel is designed as an augmented proportion guidance law, the guidance law of the yaw channel is designed as a guidance law with an attack time error feedback item, and the trajectory track is adjusted to realize time control over attack on the target; and when the yaw angle is zero, adjusting the guidance law of the yaw channel to avoid singularity conditions. According to the guidance method, the attack time error is estimated, the attack time error tends to be zero through a guidance strategy, and the distance between the missile and the target tends to be zero. The guidance method for attack time control has the characteristics of clear structure and high applicability.

An angle tracking radar system particularly for a missile with a steerable antenna and gyros strapped down to the missile body—a ‘partially strapdown’ system. The body rate signals, body acceleration signals where provided, and target position signals are converted into an electronic reference frame which is controlled to align with the target sightline, the above body and target signals being employed to produce estimates of target direction, sightline rate and sightline acceleration for use in controlling the missile.

The invention discloses a guidance information processing method aiming at seeker loss-of-lock conditions. The guidance information processing method comprises the following steps of: calculating a theoretical line-of-sight rate according to the position and speed information of a guided missile, and employing different switching methods between the actually measured line-of-sight rate and the theoretical line-of-sight rate under the condition that the seeker loses a target according to different distances between the guided missile and the target; and calculating a theoretical frame angle according to the position and attitude angle information of the guided missile, and presetting the frame angle of the seeker under the condition that the seeker loses the target.

A system includes an angular rate sensor disposed in a vehicle for providing angular rates of the vehicle, and an instrument disposed in the vehicle for providing line-of-sight control with respect to a line-of-sight reference. The instrument includes an integrator which is configured to integrate the angular rates of the vehicle to form non-compensated attitudes. Also included is a compensator coupled across the integrator, in a feed-forward loop, for receiving the angular rates of the vehicle and outputting compensated angular rates of the vehicle. A summer combines the non-compensated attitudes and the compensated angular rates of the to vehicle to form estimated vehicle attitudes for controlling the instrument with respect to the line-of-sight reference. The compensator is configured to provide error compensation to the instrument free-of any feedback loop that uses an error signal. The compensator may include a transfer function providing a fixed gain to the received angular rates of the vehicle. The compensator may, alternatively, include a is transfer function providing a variable gain as a function of frequency to operate on the received angular rates of the vehicle.

An improved display provides information regarding a projectile trajectory so that a user is informed whether or not there is a clear shot. The improved display provides trajectory path indicator shown over cross hairs. In some embodiments, an intermediate point in the projectile trajectory is an aiming point and is indicated in relation to the visualized target. A rangefinder device is calibrated to a weapon having a sight, such as a bow sight. The aiming point is displayed showing the intermediate point in the projectile trajectory that corresponds with a pin in the bow sight. The rangefinder device may be calibrated using a calibration sheet. The display may include a level. The trajectory path indicator shows the aiming point, which is the highest point in the projectile trajectory as perceived by the user.

The invention provides a heterogeneous aircraft cooperative guidance method considering attack time and attack angle constraints. The method comprises the following steps of S1, establishing aircraftmathematical models and a relative motion relationship; S2, designing a specified time hit guidance law for a main aircraft; and S3, designing guidance laws for slave aircrafts, wherein the aircraft group is composed of a single main aircraft and a plurality of slave aircrafts, and only the main aircraft can receive an attack time instruction. According to the method, firstly, the mathematical models of the main aircraft and the slave aircrafts are established, a motion relationship between the main aircraft and the slave aircrafts relative to a target is given, and the guide law for hitting the target at a specified moment for the main aircraft is designed on the basis; and in order to ensure that the slave aircrafts and the main aircraft hit the target at the same time, a cooperative guidance law is designed to enable the distance between the slave aircrafts and the target to be consistent with the distance between the master aircraft and the target, and the sight angle of the slaveaircrafts relative to the target tends to a preset expected angle.

The invention provides a method and equipment for calibrating a fighter weapon system by a fiber-optic gyroscope. The equipment comprises an azimuth and posture determining instrument and fighter vertical axis azimuth determining equipment, wherein the azimuth and posture determining instrument comprises an azimuth and posture regulator, a second fiber-optic gyroscope module is mounted on the azimuth and posture regulator, a computer for analyzing azimuth angle data is embedded into the second fiber-optic gyroscope module, an electronic optical aligner and a display are mounted on the second fiber-optic gyroscope module, both an integral solving module of the second fiber-optic gyroscope module and the display are connected with the computer through data transmission interfaces, the computer is connected with an integral solving module of a first fiber-optic gyroscope module through a data transmission interface, and the fighter vertical axis azimuth determining equipment can measure the current vertical axis direction of the fighter and indicate the direction through a cross line of sight, displayed in a level eyepiece. The method and the equipment have the advantages that steps in the process for calibrating fighter weapons can be effectively simplified, the calibrating efficiency can be effectively improved and the affect of natural factors can be prevented.

The invention relates to a planetary power descending branch navigation method based on a TDS (total descending sensor) and image measurement and belongs to the technical field of deep space exploration. At a power descending branch, an inertial unit of a landing device navigation system is used for flight position recurrence and is a basic sensor; a Doppler radar on the TDS is used for directly acquiring distances from a landing device to the surface of a target planet in three wave beam directions and relative speeds of the landing device along the three wave beam directions and then calculating the altitude of the landing device and speeds of the landing device in three axial directions of a landing point fixedly-connected coordinate system; a navigation camera serving as a vision auxiliary navigation sensor is used for precisely measuring included angles between sight directions of different characteristic points of a landing area; the positions of the characteristic points relative to the landing point are known; the sight included angles contain horizontal position information of the landing device relative to the landing point, so that the horizontal position precision of the landing device is greatly improved.

The obstacle detection assembly is provided for a rotary wing drone, and comprises an obstacle detection device having a motorized detection rotating support configured to be fastened on the drone, and an obstacle detection unit carried by the detection rotating support, the obstacle detection unit bearing at least one obstacle detection sensor and having a line of sight, and an orientation module configured to command the detection rotating support so as to orient the line of sight of the obstacle detection unit as a function of the movement direction of the drone bearing the detection rotating support.

The invention discloses an attack angle constrained high-speed interception guidance method. The attack angle constrained high-speed interception guidance method comprises the following steps of establishing the relative motion relation between a guided missile and a target with the longitudinal plane where the guided missile and the target are located as an attack plane to obtain an expression ofrelative motion parameters; deducing expressions of the zero effort miss and zero effort speed according to the missile-target relative motion parameters; making the zero effort miss and zero effortspeed as first-layer sliding mode variables according to the hierarchical sliding mode theory and combining a first-layer sliding mode into a second-layer mode; and selecting an exponential reaching law, making the exponential reaching law equal to the second-layer mode about the time derivative and then solving the falling angle constraint missile guidance law of a high-speed flight missile. Through the attack angle constrained high-speed interception guidance method, the missile hits on the target at the preset angle of sight to improve the destroy effect. When the missile intercepts the target at high speed, the missile hits on the target with small zero effort miss. Thus, the attack angle constrained high-speed interception guidance method has good reliability.

Acoustic pyrometry-based active temperature monitoring of gas turbine combustors, including industrial gas turbine (IGT) combustors is incorporated into the combustion monitoring and control system by addition of an acoustic transmitter or acoustic transceiver that transmits a sound wave in a line-of-sight with a plurality of thermoacoustic sensors, such as dynamic pressure sensors. Sound transmission time-of-flight is measured by the controller and correlated with path temperature along the line-of-sight. Path(s) of acoustic transmission serve as absolute temperature measurement that optionally is used for calibrating dominant mode passive bulk temperature measurement. In an integrated thermoacoustic pressure-based sensor and monitoring / control system embodiment, the controller correlates performance of an combustion thermoacoustic properties in order to identify combustion anomalies by wavelet or Fourier analysis techniques, determine bulk temperature characteristics within the combustor with dominant mode frequency analysis techniques and determines absolute active path temperatures within the combustor with acoustic transmission and time-of-flight analysis techniques.

The invention relates to a pre-guidance and attitude stabilization matching guidance method based on nonlinear conversion, and belongs to the technical field of aircraft guidance. The method comprisesthe following steps that the sight angle rate of an aircraft moving relative to a target is measured, and the yaw angle and the yaw angle rate of the aircraft are measured; a sight angle signal is constructed according to the sight angle rate, and an azimuth error signal and a pre-angle error signal are constructed according to the sight angle signal and the yaw angle; an amplitude limiting nonlinear signal is constructed according to the azimuth error signal, and a pre-guidance law is constructed according to the amplitude limiting nonlinear signal and the pre-angle error signal; and an attitude stabilization output signal is constructed according to the pre-guidance law, and an expected course angle of the aircraft is obtained according to the attitude stabilization output signal so that the yaw angle can stably track the expected course angle. According to the method, the problems that guidance law output is rapidly increased, and the missile miss distance is too large are solved.

The invention relates to a missile-borne SAR motion compensation method, which comprises the steps of establishing a motion geometric model of missile-borne diving SAR imaging; integrating the actualmotion track of a carrier by using low-precision inertial measurement unit data according to a northeast sky speed measurement value, fitting an ideal track, calculating a position error, uniformly projecting the position error to the center of a wave beam, and completing consistent compensation of a sight-direction motion error by using a slant distance error of each reference point on a scene center line; combining the echo data and the inertial measurement unit data, giving a Doppler center initial value by the inertial measurement unit data, estimating a residual Doppler center by using acorrelation function method to obtain a real unambiguous Doppler center, and completing large-distance walk correction; after the distance is compressed, carrying out space-variant compensation of sight-direction motion errors caused by different slant distances on each distance unit; and combining the echo data and the inertial measurement unit data, giving a Doppler frequency modulation rate initial value according to the inertial measurement unit data, estimating the residual Doppler frequency modulation rate by an MD algorithm, calculating the azimuth residual phase and compensating to realize fine motion compensation of the missile-borne SAR.

The invention relates to a composite guidance method based on sight angle information and front guidance information, and belongs to the technical field of aircraft guidance. The method is characterized by comprising: measuring sight angle information of relative movement of an aircraft and a target according to an aircraft seeker, comparing with course angle information measured by a gyroscope toobtain an attitude error signal; setting a preposed angle and comparing the preposed angle with a course angle measured by a gyroscope to obtain a preposed error signal; after the two signals are synthesized, forming proportional-integral type pre-guide information through proportional-integral operation, finally, linearly superposing the proportional-integral type pre-guide signal and sight angle information to form a final guide law, driving an aircraft course angle tracking loop, and achieving accurate guide of an aircraft to a target. The problem that in the prior art, the two advantagesof front guiding and sight angle guiding cannot be considered at the same time is solved.

The invention discloses a reminding system for preventing a dead zone caused by keeping out of a heavy vehicle and a work method of the reminding system. The reminding system comprises a travelling vehicle, the traveling or stopped heavy vehicle and a laterally-coming obstacle, front sensors are mounted at the front end of the travelling or stopped heavy vehicle, laser transmitters are mounted onrear view mirrors of the travelling or stopped heavy vehicle, and lateral sensors are mounted on the left side and the right side of the traveling or stopped heavy vehicle. The specific method comprises the steps of detection of the front sensors, judgment of a dangerous state, judgment of the dangerous state, judgment of the moving direction of the obstacle, work of the laser transmitters and turning off of lasers. The sensors and the laser transmitters are additionally mounted on the heavy vehicle, so that the front or lateral state of the heavy vehicle is perceived, the laser transmitters are utilized to transmit laser beams to provide dead zone information reminding for the peripheral vehicle, thus a driver rapidly obtains dead zone information and decelerates the vehicle in time underthe premise that attention is not changed, and sight does not leave from the front position and the view angle, the traffic accident is avoided, responding is sensitive, and the safety is high.

High-brightness LPP source and method for generating short-wavelength radiation which include a vacuum chamber (1) with an input window (6) for a laser beam (7) focused into the interaction zone (5), an output window (8) for the exit of the short-wavelength radiation beam (9); the rotating target assembly (3), having an annular groove (11); the target (4) as a layer of a molten metal formed by centrifugal force on the surface of the distal wall (13) of the annular groove (11) while the proximal wall (14) of the annular groove is designed to provide a line of sight between the interaction zone and both the input and output windows particularly during laser pulses. A method for mitigating debris particles comprises using an target orbital velocity high enough for the droplet fractions of the debris particles exiting the rotating target assembly not to be directed towards the input and output windows.

The invention discloses an aircraft composite guidance method based on a non-linear sliding mode and preposition, and belongs to the technical field of aircraft flight guidance. The method comprises the following steps: firstly, measuring a sight angle of an aircraft by adopting seeker equipment, and measuring a yaw angle of the aircraft by adopting a gyroscope; constructing a sight angle approximate differential signal by adopting a first-order filter, and forming a nonlinear sliding mode signal by the sight angle approximate differential signal and the sight angle signal; meanwhile, settinga precondition to obtain a preangle, comparing the preangle with a sight angle and a yaw angle to obtain a preerror signal and a sight attitude error signal, conducting integration to obtain an integral signal, and then forming a nonlinear sliding mode signal based on errors; finally, conducting nonlinear recombination on the two types of nonlinear sliding mode signals based on the sight angle andthe error, and obtaining a final nonlinear sliding mode and front comprehensive guiding signal and transmitting to an attitude angle stabilizing system to achieve precise guiding. The method has theadvantages of being high in guiding precision, good in robustness and high in adaptability.

The windbreak eye shield type assembly (10) comprises members (19, 19A) for mounting the shield (10) so as to be mobile on a helmet (1), so that the shield (10) can be moved away from the line of sight, under the control of wind pressure sensing members, remaining opposed to the wind.

The invention provides a high-precision multi-rotor aircraft cooperative positioning method and system. The method comprises the steps: selecting a target needing to be located; acquiring informationdata by n unmanned aerial vehicles, the information data comprising parameter information of the unmanned aerial vehicles, target sight angle information and distance information between the unmannedaerial vehicles and a target; and processing the information data to obtain an estimated value of the target position. According to the method provided by the invention, multiple unmanned aerial vehicles are adopted to cooperatively acquire information, and the installation error between the photoelectric pod and the integrated navigation module is introduced into the least square model, so that high-precision positioning of a long-distance target in a short time is realized under the condition that the precision of an inertial navigation element is relatively low, the positioning precision ishigh, and the time is short.