46 results about "Bearing (navigation)" patented technology

The invention provides an attitude measurement method, relates to the technical field of directional drilling measurement while drilling, and aims to improve the observability of an inertial instrument error, inhibit the repeatability error of a gyroscope and improve the attitude measurement precision. According to the method, initial alignment is performed by adopting a method of respectively performing fine alignment at multiple positions. The method comprises the following steps of S1, taking the current attitude and speed data of the gyroscope as a first initial value, and performing finealignment at a first position; S2, enabling a strapdown inertial navigation system to be trans-positioned to the n position, and performing attitude and speed updating according to the last fine alignment result in the trans-positioning process; and S3, taking attitude updating and speed updating results as the n initial value, and performing n fine alignment at the n position to complete initial alignment of the strapdown inertial navigation system so as to realize attitude measurement. The scheme provided by the invention is suitable for horizontal attitude and azimuth measurement of a whole well inclined section of the horizontal well, and is especially suitable for attitude measurement application of directional drilling gyro measurement while drilling in large well inclined shafts and horizontal wells.

The invention relates to a complementary filtering attitude calculation method for a navigation attitude reference system, and the method comprises the steps: collecting gyroscope output, accelerationoutput and magnetometer output, and carrying out the normalization; carrying out horizontal alignment by utilizing the first group of accelerometer data, calculating a pitch angle theta, a roll anglegamma and an attitude conversion matrix from the system to a horizontal system, carrying out azimuth alignment by utilizing the first group of magnetometer data, calculating an initial course angle,and acquiring gyroscope output, acceleration output and magnetometer output data; calculating a gyroscope error by utilizing the acceleration and magnetometer output data, calculating a magnetic fieldintensity reference value of a local navigation system, and calculating a projection value fn of acceleration output on the navigation system; calculating an acceleration crossover result fX, calculating a magnetic vector crossover result calculation error, calculating a gyroscope data error correction value, and updating the attitude after output correction of the gyroscope; and calculating an attitude angle according to the updated attitude information.

The invention discloses a method for separating an attitude angle error of an inertial measurement system based on centrifugal machine excitation. The method includes that the characteristic that the orientation of a centrifugal machine reverse rotation platform relative to the ground is unchanged is utilized, and inertial platform system frame angle information is taken as observed quantity; an attitude angle error with an error can be obtained by comparing the observed quantity with an attitude angle calculated by an inertial navigation attitude, and the error is mainly caused by dynamic errors of a gyroscope and an inertial measurement system; various errors related to the attitude angle error can be separated by adopting a least square method, and the precision of inertial navigation can be improved through error compensation.

The embodiment of the invention relates to an inter-ship relative field acquisition method based on a ship cluster situation. The inter-ship relative field acquisition method comprises the following steps: dividing an interest perception area of a target ship into a plurality of interest perception subareas according to a ship safety meeting distance; according to the navigational speeds and the navigational directions of the target ship and the interference ship, the relative directions of the target ship and the interference ship, and the collision avoidance response time, in combination with the ship length of the target ship and / or the interference ship, calculating to obtain inter-ship variable advancing distances of the target ship under different meeting situations; determining a weight factor of the inter-ship variable carry distance in the plurality of interest perception sub-regions; calculating to obtain an inter-ship variable path distance according to the inter-ship variable advance distance and the weight factor; and obtaining a plurality of boundary points according to the inter-ship variable path distances of the interest perception sub-regions, and fitting to obtain the inter-ship relative field based on the ship cluster situation. According to the inter-ship relative field acquisition method, the inter-ship variable pitch is calculated according to the inter-ship variable pitch under different meeting types of the target ship and the interference ship in combination with the weight factor, and the inter-ship relative field based on the ship cluster situation is established in real time.

The invention relates to a method and a device for calibrating an error of an SINS / DVL integrated navigation system. The method comprises the following steps: modeling a nonlinear optimization function; parameters of the nonlinear optimization function comprising a scale factor error of the speed of the DVL and an azimuth installation error angle between the SINS and the DVL; and solving the nonlinear optimization function to complete calibration of the two errors. According to the method for calibrating the errors of the SINS / DVL integrated navigation system, a nonlinear optimization function is modeled by aiming at the two parameters, the nonlinear optimization function is solved by utilizing a Gauss-Newton algorithm, the two errors are accurately calibrated, and the calibration precision of the method can reach the optimal solution through comparison with an enumeration method. According to the method for calibrating the error of the SINS / DVL integrated navigation system, the precision of SINS / DVL integrated navigation can be improved, and the method has very high practical value.

The invention belongs to the technical field of precision rolling bearings, and relates to a matching method for special bearings for inertial navigation platforms. The matching steps are: (1) preselection, (2) pairing, (3) measuring and correcting the protrusion of the end face of the single set of bearings, (4) measuring and correcting the deformation of the inner ring of the paired bearing, (5) testing the paired bearing starting friction torque. The method of the invention can reduce the error of the pre-tightening force of the paired bearings, make the starting friction torque of the bearings meet the requirements, and greatly improve the performance of the inertial navigation platform.

The invention relates to a navigation painted abacus, which comprises panels 1, a movable rotary disk 2, a logarithmic calculating disk 3, a movable scale 4 and a logarithmic ratio scale 5, wherein, the panels 1 are divided into an upper layer and a lower layer; the movable rotary disk 2 is a circular disk and fixed between the upper and lower panels 1, and can freely rotate around the center of the circle; the movable rotary disk 2 is engraved with two main index lines 6 which pass through the center of the circle and are perpendicular to each other, and a plurality of auxiliary index lines 7 parallel to the main index lines 6; the disk surface is engraved with a distance / speed circle 8 used for measuring the distance / speed, and the outer ring of the rotary disk is engraved with an azimuth scale line 9 of 0 to 360 degrees. The upper and the lower panels 1 are made of organic glass, lined by copper strips, and fixed by screws; and each corner is subjected to slow round processing. The logarithmic calculating disk 3 consists of an inner disk 10 and an outer disk 11 which have the same circle center, wherein, the inner disk 10 can freely rotate, and the outer disk 11 is a fixing disk. The movable scale 4 is made of copper materials, concentric with the movable rotary disk 2 and fixed on the upper panel 1 by the inside thread screws, and can freely carry out circumferential rotation. The logarithmic ratio scale 5 is a strip-shaped plexiglass ruler, one side of which is engraved with logarithmic scale, and can be used as a logarithmic slide rule by matching with the movable scale (4) on the movable rotary disk (2).

The invention discloses a manual inertial navigation test turntable with high sensitivity. The manual inertial navigation test turntable comprises a U-shaped base, the upper end of the inner side of the U-shaped base is provided with a table top, the two ends of the table top are connected with a left main shaft and a right main shaft respectively, the left main shaft is sleeved with a base bearing assembly and an angle measurement assembly which are arranged in the left end of the U-shaped base, the right main shaft is sleeved with a right bearing assembly and a worm wheel which are arrangedin the right end of the U-shaped base, and the lower end of the worm wheel is connected with a worm in the adjusting assembly in a meshed mode. The manual inertial navigation test turntable reasonablyand effectively solves the problems that an existing inertial navigation test turntable shakes after being positioned, is not high in self-positioning sensitivity and cannot meet the test and calibration requirements of a high-precision inertial measurement unit, a gyroscope and an accelerometer, and has the characteristics of compact structure, small occupied space and high sensitivity.

The invention discloses a method for quickly correcting deep sea inertial navigation orientation, which belongs to the technical field of inertial navigation, and aims to correct the orientation of the deep sea inertial navigation, greatly shorten the correction time and greatly improve the correction precision. The method comprises the following steps of S1, carrying a strapdown inertial navigation system and a Doppler log DVL on a carrier, wherein the strapdown inertial navigation system comprises two parallel strapdown integral resolvers, namely a first resolver I and a second resolver II,S2, constructing a horizontal fine alignment filter, and estimating a horizontal misalignment angle according to the outputs of the first resolver I and the Doppler log DVL, S3, feeding back the horizontal misalignment angle to the second resolver II, and controlling the carrier to start turning maneuver, S4, constructing an orientation fine alignment filter, and estimating an orientation misalignment angle according to the output of the second resolver II and the Doppler log DVL, and S5, feeding back the orientation misalignment angle to the second resolver II, completing carrier attitude correction, and feeding back the corrected carrier attitude to the first resolver I.

The invention discloses a method for determining a safe course interval of a ship and a related system. The method comprises the steps: obtaining a relative motion speed vector of critical safe avoidance according to a safe distance to the closest point of approach DCPAS and the current direction of a target ship; calculating a course interval of the ship meeting the condition that DCPA is greaterthan or equal to DCPAS; solving any course value C''o in the safe course range of the ship, and expressing the calculated safe course range of the ship in an interval form; and when multiple ships meet, calculating a safe course interval of the ship for each target, taking the intersection of all the safe course intervals as the safe course interval of the ship for avoiding collision with multiple ships, and selecting the safe course with the minimum steering amplitude as a new navigation direction for steering collision avoidance. The safe course of the ship is expanded to one or two interval sections from one constant value, the collision avoidance amplitude selection margin is larger, and the flexibility is higher; and meanwhile, based on the current collision avoidance situation, theresult calculated by the method has the advantages of good real-time performance, high effectiveness and good flexibility.

The invention discloses a garden playing navigation device which comprises a shell, a damping mechanism and a navigator are arranged on the inner wall of the shell, rotating mechanisms are arranged onthe two sides of the shell, a shell cover is arranged in the shell, the shell cover is connected with the shell through the rotating mechanisms, a lifting belt mechanism is arranged in the shell cover, the lifting belt mechanism comprises a cavity, the cavity is formed in the shell cover, first bearings are fixedly installed on the two sides of the cavity, a winding roller is arranged on the first bearings, a lifting belt body is arranged on the shell cover, the two ends of the lifting belt body penetrate through the cavity and are wound around the winding roller, a spring groove is formed inthe shell cover, a first spring is arranged in the spring groove, a limiting plate is arranged at one end of the first spring, a second bearing is arranged on the limiting plate, the second bearing penetrates through the cavity, a rotating rod is arranged on the second bearing, one end of the rotating rod penetrates through the shell cover, a first handle is arranged at one end of the rotating rod, and a gear ring is arranged at the bottom end of the rotating rod. By arranging the shell cover and the rotating mechanisms, the purposes of conveniently opening the shell cover and facilitating use of a navigator are achieved, meanwhile, the shell cover can be conveniently fixed to different angle positions, the navigator can be conveniently supported and placed through the shell cover, and the purpose of convenient use is achieved.

The invention provides a self-calibration method for an optical fiber strapdown inertial navigation system, which includes: northward speed error, eastward speed error, horizontal misalignment angle, azimuth misalignment angle, gyro constant value offset and accelerometer constant value offset The quantity is the state quantity, and the state equation is established; the northward velocity error and the eastward velocity error are used as the quantity measurement, and the measurement equation is established; the state equation and the measurement equation constitute a Kalman filter model; When the SINS is located at the initial position, the state quantity is calculated according to the Kalman filter model; after a preset time period, the heading angle of the optical fiber SINS is rotated by 180 degrees, and then according to the Kalman filter The filter model calculates the state quantity, realizes the self-calibration of the constant value offset of the gyro and the constant value of the accelerometer, improves the observability of the system, and can improve the calibration accuracy while meeting the requirement of rapidity.

The embodiment of the invention provides an accurate positioning and parking method of a trackless navigation AGV. The accurate positioning and parking method comprises the steps that the AGV scans a reflection plate through laser radar to obtain coordinates of preset positions on the reflection plate; a position for positioning is selected from the preset positions; the current position of the AGV is calculated according to preset rules and the selected position. According to the method, three special reflection points are extracted to achieve three-point positioning, the coordinate orientation of the AGV under a global coordinate system is determined mainly through a least square method, and then the direction angle of the AGV is calculated through multiple quadrants. The relative position and the direction angle of the AGV with respect to the reflection plate can be also calculated through outline detection so that the AGV can be accurately parked at the designated of the reflection plate. By adopting the method, the defect of poor positioning accuracy of an SLAM technology is overcome to the most degree, and a reliable solution is also provided for automatic charging of the trackless navigation AGV.

The invention provides a mine inertial navigation and magnetic compass error angle calibration and compensation method. According to the uniqueness principle of a state cosine matrix between two space coordinate systems, a calculation result obtained after measurement in a first state is used for calculation of a measurement result in a second state; the limitation condition of 90-degree rotation around the X axis does not exist between the two states of the mine in the Y-axis direction and the Z-axis direction, the requirements for calibration equipment and operation are simplified, and implementation is easy; a calibration result obtained by aiming at each mine is written into a trajectory control computer, and unification of azimuth and attitude references of a target detection system and a trajectory control system is realized through compensation, so that system errors caused by equipment installation are effectively eliminated, and the accuracy of the active attack mine to implement guidance attack on a target is ensured; and the method is verified by lake and sea tests at a certain model scheme stage.

The invention discloses a civil aviation navigation equipment sight distance coverage analysis method combined with variable height of a flight path, belongs to the technical field of civil aviation navigation equipment, and solves the problem that the flight safety cannot be ensured due to the absence of a method for carrying out sight distance coverage analysis on a variable height flight path in the prior art. The method comprises the following steps: S1, calculating azimuth angles B of all obstacles relative to navigation equipment by taking true north as 0 degree; s2, screening obstacles in the same direction as the flight path; and S3, calculating whether the n obstacles shield the flight path or not, and calculating the shielding position. The invention provides a method for calculating the shielding of the obstacle to the variable-height flight path for the first time, which comprises the specific shielding position, namely the critical position of the received signal from existence to absence, and the signal coverage condition on the flight path after shielding, so that a reference basis is provided for the height control of the obstacle and the adjustment of the flight height of the aircraft; and the flight safety is powerfully ensured.

The invention provides an automatic navigation and positioning cutting method and system for a cantilever type heading machine, and relates to the technical field of intelligent mining of coal mines, and the method comprises the following steps: setting the design orientation, inclination angle and section parameters of a heading roadway, automatically leveling and north-seeking a full-automatic gyroscopic total station, searching a roadway traverse point prism in a rearview manner, and calculating the geodetic coordinates of a set site; the geodetic coordinates of a rack prism are searched, tracked and measured in a foresight mode, the geodetic coordinates of a gun head are calculated in real time based on lifting and rotating oil cylinder displacement values and the physical size of the heading machine, and lifting and rotating oil cylinder actions of the heading machine are controlled by combining the heading roadway design orientation, the dip angle, section parameters and the cutting process. The heading machine is enabled to perform positioning cutting according to set parameters and technological processes, and position and heading can be dynamically rectified. The absolute geodetic coordinates of the cutting head and the relative coordinates under the cutting section coordinate system are calculated in real time, unification of the absolute coordinates and the relative coordinates is guaranteed, and errors caused by posture changes of a machine body in the working process of the cantilever type heading machine can be dynamically corrected and compensated.

The invention belongs to the technical field of flight navigation, and particularly relates to a method and system for intelligently selecting a TACAN station through inertial navigation combination correction. The method comprises the following steps: S1, determining a nearest TACAN station in a preset range; s2, acquiring the position information, the TACAN orientation, the TACAN slant distance and the aircraft height information of the TACAN; s3, calculating first longitude and latitude information according to data given by the TACAN station; s4, acquiring second longitude and latitude information given by the inertial navigation system; s5, determining the deviation between the first longitude and latitude information and the second longitude and latitude information; and step S6, judging whether the deviation is smaller than a given value, and if the deviation is smaller than the given value, correcting second longitude and latitude information given by the inertial navigation system based on the deviation in the subsequent flight process of the aircraft. According to the method, the inertial information given by the inertial navigation system in the flight process of the TACAN data correction aircraft can be corrected, so that the navigation accuracy is improved.

The invention discloses a method for analyzing the line-of-sight coverage of civil aviation navigation equipment combined with flight track and variable altitude, which belongs to the technical field of civil aviation navigation equipment, and solves the problem that there is no method for line-of-sight coverage analysis for the flight track with variable altitude in the prior art, and flight safety cannot be guaranteed The problem. The method of the present invention comprises the following steps: S1. Taking the true north as 0 degree, calculate the azimuth angle B of all obstacles relative to the navigation equipment; S2. screen the obstacles with the same orientation as the flight track; S3. calculate n obstacle pairs Whether the flight path causes occlusion, and calculate the occlusion position. The present invention proposes for the first time the blockage calculation method of the obstacle to the variable altitude flight track, including the specific block position caused, that is, the critical position where the received signal changes from presence to non-existence, and the signal coverage on the flight track after blockage. Altitude control and aircraft flight altitude adjustment provide a reference basis, which effectively guarantees flight safety.

The invention discloses a light, small, high-precision composite inertial navigation system and a navigation mode switching method. The navigation system includes a base, an outer rotating frame mounted on the base through a first bearing rotation, and an outer rotation frame installed on the outer rotation frame through a second bearing rotation. The inner rotating frame on the rotating frame and the inertial measurement unit arranged on the inner rotating frame, one end of the inner rotating frame and the outer rotating frame are respectively provided with a driving assembly, and the other end is respectively provided with a locking mechanism, and the driving assembly includes an ultrasonic motor and an angle The sensor and the locking mechanism include a moving plate, a disc spring, a coil, a casing and a fixed plate. The moving plate is in contact with the inner rotating frame and the outer rotating frame respectively. The disc spring is arranged between the moving plate, the fixing plate and the housing. The magnetic field generated by the coil is stretched to realize the support and distance of the moving plate, and the coil changes the magnetic field by controlling the current input. Ultrasonic motor and electromagnetic lock can be quickly locked and unlocked in harsh mechanical environments to meet the needs of complex mechanical environments.

A method for determining an azimuth angle of a wind power installation is provided. The method includes attaching at least two global navigation satellite system (GNSS) receivers to a nacelle, comparing reception signals of the GNSS receivers, deriving the azimuth angle from the comparison. The at least two GNSS receivers may be attached to a wind measuring supporting frame. Provided is a method including attaching a telescopic sight to a nacelle, determining a bearing of a bearing object at a bearing angle, deriving the azimuth angle from a comparison of the bearing angle with coordinates of the bearing object and / or the wind power installation. The telescopic sight may be attached to a wind measuring supporting frame. The at least one GNSS receiver may be attached to a wind measuring supporting frame. A wind power installation is also provided.

The invention relates to a ship-borne inertial navigation attitude angle time sequence detection method based on fixed star observation. The method comprises the following steps: 1, using inertial navigation attitude angle data and angle measurement data of a theodolite to calculate the actual measurement orientation of fixed stars under the horizon system; 2, using a least square method to fit and calculate the inertial navigation attitude angle data at any time; 3, calculating the inertial navigation attitude data which exceeds or lags behind different time, respectively combining the data with data observed by using the theodolite, and finding a set of inertial navigation attitude data with the smallest standard deviation of the azimuth and pitch angle error of the fixed stars under theground level, wherein the corresponding time sequence is an inertial navigation attitude angle time sequence error. The ship-borne inertial navigation attitude angle time sequence detection method based on fixed star observation has higher inertial navigation time sequence error detection accuracy and can effectively improve the inertial navigation attitude angle measurement accuracy.

The invention discloses an inertial navigation-based aerostat occultation prediction method, which includes first obtaining the positions of all GNSS navigation stars in the WGS84 coordinate system according to the positioning equipment of the occultation detector, and then calculating the position, longitude, and Latitude and altitude, calculate the pitch angle, azimuth angle of all GNSS navigation stars and the azimuth angle of the aerostat body, judge whether the obtained pitch angle Elv, azimuth angle Azm and antenna installation azimuth meet the threshold of the occultation event, if so, then It is determined that the GNSS navigation star is an occultation. The present invention judges occultation events by adopting continuous arc segments by inertial navigation equipment, thereby enhancing the accuracy of forecast results, the method is simple and easy, and has important application value for real-time forecast.

The present invention discloses a solar TDOA measurement method and an integrated navigation method oriented to formation flight, including establishing a solar TDOA measurement model as follows, executing an orbital dynamics model oriented to formation flight for detectors in formation flight, and establishing a Mars direction measurement model , establish the inter-satellite link ranging model; use the extended Kalman filter EKF as the navigation filter, use the star sensor to obtain the celestial body orientation information, the inter-satellite link measurement value and the relative distance in the radial direction of the sun, combined with the formation The orbital dynamics model of the flying probe, under the action of the EKF, obtains the estimated position and velocity information of the probe. The present invention proposes a solar TDOA measurement technical solution for formation flight, which can provide relative position information in the radial direction; the present invention makes full use of a variety of navigation measurement information, and can be used for formation flight in the entire space The sensor provides absolute and relative navigation information with higher accuracy.

The invention discloses an AGV navigation sensing device based on navigation map design and a scheduling method. The AGV navigation sensing device comprises an AGV body, a wheel groove is formed in the bottom of the AGV body, wheels are arranged in an inner cavity of the wheel groove, bearings are arranged in the inner cavity and on the right side of the AGV body, protection covers are fixedly connected to the two ends of the bearings, a driven wheel is arranged in an inner cavity of the protection cover, a processor is arranged in the middle of an inner cavity of the assembly groove, a landmark sensor is arranged on the left side of the processor, a magnetic navigation sensor is arranged on the right side of the processor, a supporting rod is fixedly connected to the top of the protection cover, a spring groove is formed in an inner cavity of the supporting rod, springs are arranged in inner cavities of the spring grooves, damping gaskets are fixedly connected to the tops of the supporting rods, a loading plate is arranged at the tops of the damping gaskets, a connecting plate is fixedly connected between the front protection cover and the rear protection cover, screw holes are formed in inner cavities of the loading plate and the connecting plate, and screw rods are fixedly connected to inner cavities of the screw holes.

The invention discloses a laser aiming reference object-free aircraft magnetic heading calibration method, which belongs to the technical field of aviation airborne strapdown heading attitude navigation, and is characterized by comprising the following steps of: a, mounting a target plate on a longitudinal axis below an aircraft nose; b, reading the course of emitted laser; c, emitting laser; d, dragging the airplane to rotate around the center of a calibration site; e, emitting laser to the target plate through a laser azimuth instrument, and calculating to obtain a real magnetic heading angle of the airplane; and f, carrying out magnetic heading calibration on an emergency magnetic compass and a strapdown backup navigation system. The method does not need to coordinate a pilot and start an aircraft engine, an operator does not need to get on and off an aircraft back, operation is easy, safety risks are reduced, the method is not influenced by visibility, and accurate calibration can be conducted even under the weather condition of low visibility.

The invention provides an inertial navigation resolving method suitable for steady-state navigation and capable of saving a two-axis strapdown gyroscope. According to the invention, two horizontal axial gyroscopes are removed; a single-axis gyroscope of an azimuth channel is reserved, an accelerometer in a magnetic compass is used for extracting the difference between a pitching horizontal attitude angle and a rolling horizontal attitude angle in the non-maneuvering navigation process to form an angle increment, the angle increment and the single-axis gyroscope jointly form complete description of the moving work of the aircraft along three axial spaces, and updating of a mathematical platform is achieved. Due to the fact that strapdown inertial navigation calculation of the single-axis gyroscope magnetizing compass can be achieved, a two-axis strapdown gyroscope is omitted on the basis of traditional strapdown inertial navigation configuration, the size can be reduced, and cost is saved.

The invention belongs to the technical field of attitude calibration of inertial navigation systems, and in particular relates to a method for calibrating the attitude of a moving base based on self-collimation tracking. Install the inertial navigation system on the inner frame table of the three-axis swing table, so that the rotation of the outer frame of the three-axis swing table excites the heading angle change of the inertial navigation system, the rotation of the middle frame stimulates the pitch angle change of the inertial navigation system, and the rotation of the inner frame stimulates the change of the pitch angle of the inertial navigation system. The roll angle of the inertial navigation system changes; the eyepiece of the dynamic auto-collimation tracking measuring instrument is firstly self-aligned with the north reference mirror, and the output value of the dynamic auto-collimating and tracking measuring instrument is recorded at this time; The measuring instrument is self-collimated with the azimuth reference mirror of the inertial navigation system, and the output value of the dynamic self-collimation tracking measuring instrument is recorded again. The invention utilizes the automatic tracking and dynamic angle measurement functions of the dynamic self-collimation tracking measuring instrument to realize the real-time calibration of the inertial navigation system attitude under static and dynamic conditions.

The invention discloses a detachable shipborne navigational flow measurement support, and belongs to the field of ocean observation. The detachable shipborne navigational flow measurement support comprises a steel pipe; the steel pipe is of an L-shaped structure; the pipe wall of the steel pipe is sleeved with a lantern ring; a first clamping plate is fixedly installed at the bottom of the lanternring and is of an L-shaped structure; a threaded hole is formed in one side of the first clamping plate in a penetrating mode; a threaded rod is arranged in the threaded hole in a threaded fit mode;the rod wall of the threaded rod is fixedly sleeved with a second clamping plate through a bearing; the second clamping plate and the first clamping plate are arranged correspondingly; a mounting block is fixedly mounted at the bottom end of the steel pipe; a groove is formed in the bottom end of the mounting block; and a stud is fixedly mounted at the bottom of the groove. Ocean current can be observed in the advancing process of a ship, other operations are not delayed, data can be conveniently inspected at any time, and different ships can be carried to perform observation near shore.