164 results about "Orbit/Orbital" patented technology

The invention discloses an initial orbit determination method for passive angle measurement of a ring fire orbit for Mars sampling return, and the method comprises the steps: building an absolute dynamic model, obtaining an absolute motion state of a spacecraft, and deducing a relative motion state X; constructing a data set by using the measurement quantities H and X of the sensor; a network architecture is designed, input data I are position and speed coordinates of the tracker in a Mars equator inertial coordinate system and angle information converted from H, and output data are X; and finally, adjusting an observation interval dt and times k according to a ring fire observation condition, training a physical information neural network by using limited observation data, defining a total loss function F comprising a loss function FY and a physical law penalty term FD, and optimizing a weight coefficient and bias by minimizing the loss function through an F gradient descent method. According to the method, the neural network is trained offline, a network mapping model from a target line-of-sight angle measurement sequence to a target orbit is constructed, and the network mapping model is solidified on a returner for online use, so that sampler short-arc passive measurement rapid orbit determination is realized.

The invention discloses an optical measurement segmental arc correlation method based on probability expression, and the method comprises the steps: obtaining an initial orbit feasible region of a segmental arc through an admissible domain, carrying out the uniform distribution discrete sampling of the admissible domain, and describing the uncertainty distribution of the initial orbit of the segmental arc through a Gaussian mixture model. The uncertainty distribution and propagation of the orbit are described in a corrected spring equinox orbital element space, and the state transition tensor of a gravity potential model of any order is applied, so that the precision and speed of orbit uncertainty propagation are improved. In combination with a Gaussian mixture model and a state transition tensor, sparse sampling points are used as component mean propagation dense sampling points of the Gaussian mixture model, and uncertainty distribution of the initial orbit of the arc section is propagated to a common moment. And giving a definition of association probability measurement, calculating the association probability measurement by utilizing a probability density function of track distribution of the segmental arc at the public moment, and carrying out association judgment on the optical measurement segmental arc based on probability expression.

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

The invention relates to a generalized carrier rocket trajectory calculation method which comprises the steps that a target orbit and constraint conditions corresponding to the target orbit are set, the target orbit comprises a sun-synchronous orbit or a low-dip-angle orbit, and the constraint conditions at least comprise the geocentric vector diameter, the absolute speed and the flight path angle; input parameters are simplified, wherein the input parameters comprise configuration parameters, mass parameters, pneumatic parameters and engine parameters; and based on the constraint condition of the target orbit, the corresponding control variable and the input parameter, calculating the value of the control variable through a Newton iteration method so as to obtain the trajectory of the carrier rocket. The invention provides a universal carrier rocket trajectory calculation method. The carrying capacity of a carrier rocket is ensured by optimizing a typical target orbit; according to the method, input parameters of trajectory calculation are simplified and carded, the trajectory is obtained through calculation of the Newton iteration method, the adaptability to different configuration schemes is improved on the premise that the calculation accuracy is guaranteed, rapid trajectory iteration calculation can be conducted, and generalization is achieved.

The invention relates to a method for designing a transfer orbit from a DRO orbit to a lunar low-orbit circular orbit. The method comprises the following steps: establishing a mapping data table and an orbital transfer speed increment three-dimensional data pair according to a preset DRO orbit size parameter and a DRO aircraft initial moment phase angle; dividing the orbital transfer speed increment three-dimensional data pair into two groups according to positive and negative conditions of an orbital transfer speed increment elevation angle, acquiring ascending node longitude and transfer flight time corresponding to the data pair of the speed increment in the minimum transfer process of each group to calculate orbital transfer moments, and selecting a minimum value in the two orbital transfer moments as a selected orbital transfer moment; correcting the phase at the selected orbital transfer moment, and calculating an orbital transfer speed increment projection according to the corrected phase; and adopting a Newton iteration method to take the projection of the selected orbital transfer moment and orbital transfer speed increment as an iteration initial value to carry out iteration solution, and obtaining the transfer orbital transfer moment and the transfer orbital transfer speed increment. By adopting the method, the transfer track design precision can be improved.

The invention discloses a satellite orbit space environment disturbance coping method and device and a computer program product, and the method comprises the steps: obtaining the current position data and current space environment data of a target orbit, the current space environment data comprises a first high-energy particle integral flux at a high-energy particle flux target position and a first upper atmosphere daily average atmospheric density; according to the current position data, querying a second high-energy particle integral flux and a second upper atmosphere daily average atmospheric density at the high-energy particle flux target position when the target orbit environment is quiet; calculating the relative disturbance quantity of the high-energy particle radiation environment of the target orbit and the maximum relative disturbance quantity of the atmospheric density of the upper layer; determining the current high-energy particle radiation environment disturbance degree and the current atmospheric environment disturbance degree in combination with the disturbance degree judgment range; and generating risk coping suggestions. According to the invention, the operation environment disturbance condition of the spacecraft can be accurately known, and the in-orbit operation safety of the spacecraft is improved.

A pipeline deviation full-orbit fitting method of a strict regression orbit satellite can determine single-point pipeline deviation on the basis of alignment of a reference trajectory coordinate system and an orbit cutting plane according to the characteristics of a strict regression orbit, and can also perform calculation and curve fitting on data of at least one orbit. And the fitted maximum deviation is used as the pipeline deviation of the satellite, so that comprehensive analysis of strict regression precision of the satellite is carried out, limitation of single-point pipeline deviation calculation is effectively overcome, and meanwhile, the influence of single-point error data on pipeline deviation calculation can be inhibited by adopting a curve fitting method, so that the accuracy of pipeline deviation calculation is improved. And the calculation result is more accurate and reliable.

The invention relates to an orbit identification method based on earth-moon collinear translation point orbit parameter characterization. The method comprises the following steps: describing a circular restrictive three-body problem as a chaotic Hamiltonian power system, converting a motion equation into a new coordinate system for series expansion, and taking a quadratic term of a Hamiltonian function in a polynomial form as a linear model of a translation point in the circular restrictive three-body problem; and carrying out complex transformation and regular transformation on the linearized model transformation to realize the decoupling of the hyperbolic unstable direction and the central direction of the central manifold, and constructing a mapping relation between CRTBP coordinates and characterization parameters to carry out parameter characterization. Selecting spacecraft coordinates from parameter characterization to calculate characterization parameters of a reference orbit and integral initial values of the coordinates, and then calculating mean square errors of a real orbit and the reference orbit; and by taking the mean square error as a target function and a preset constraint condition, constructing an orbit identification solving model and solving to obtain an orbit identification result. By adopting the method, the track identification accuracy can be improved.

The invention discloses a method and device for planning fixed-time fixed-point landing and fixed-time takeoff of lunar sampling, and the method comprises the steps: determining the six orbit elements and orbit injection time of an orbit in which a first control stage is located, and the nominal parameters of fixed-time fixed-point landing and fixed-time takeoff in the lunar sampling task according to the first control stage in the lunar sampling task; and according to the nominal parameter, the six orbit elements and the orbit injection time, performing joint planning on the first control stage and the subsequent control stage to obtain a first deviation, and performing iteration on the first deviation according to the first variable. And when the first deviation is smaller than a convergence threshold, determining a target variable of the first control stage, so that the satellite achieves fixed-time fixed-point landing and fixed-time takeoff based on the target variable of each control stage, thereby realizing dynamic joint planning by taking fixed-time fixed-point landing and fixed-time takeoff as a final target in each control. The control reconstruction capability of the task is effectively improved, additional control is avoided, and meanwhile resource consumption is reduced.

The invention provides a low-energy transfer window rapid design method and system for earth-moon DRO, and relates to the technical field of aerospace. Based on a double-circle restrictive four-body model, an earth-moon transfer orbit database is constructed, the earth-moon transfer orbit database comprises a plurality of transfer orbits from a target near-earth orbit to a target long-distance retrograde orbit; determining a target track point from the track distribution diagram corresponding to each track leveraging category according to the track information track distribution diagram of each transfer track and the input near-earth track information and epoch time of the target spacecraft, and correcting the track information of the transfer track corresponding to the target track point based on an ephemeris model, and obtaining a target transfer orbit of the target spacecraft. According to the method provided by the invention, rapid and accurate planning of the earth-moon DRO low-energy transfer window can be realized, the processing time is reduced, and the planning efficiency is improved.

The invention provides a space target orbit maneuver identification method and system based on adaptive LSM, and the method comprises the steps: obtaining the orbit data of a space target, calculating the six orbit elements based on the orbit data, and coding the six orbit elements into a pulse sequence; and inputting the pulse sequence into an improved adaptive LSM for detection to obtain an orbit maneuver behavior identification result of the space target. According to the method provided by the invention, six orbital elements are calculated through the speed and the position of the orbital, then the orbital parameters are input into an improved LSM, and the improved LSM adopts an adaptive threshold neuron as a calculation unit, so that the threshold can be dynamically adjusted according to the current pulse response and attenuated along with time; meanwhile, an improved particle swarm optimization algorithm is combined to carry out optimization setting on an initial value of a neuron, so that the initial performance of the model is improved; and a learnable neuron connection weight based on an attention mechanism is introduced, so that the accuracy and reliability of orbit prediction are improved.

The invention discloses a real-time closed-loop pursuit game method and device for an elliptical orbit of a spacecraft, belongs to the crossing field of spacecraft orbit dynamics and intelligent game control, and particularly relates to the real-time closed-loop pursuit game method for the elliptical orbit of the spacecraft based on a convex optimization theory. The problems that a traditional orbit game control method is difficult in equilibrium solution, poor in strategy convergence and difficult in on-orbit application are solved. The method comprises the following steps: solving an equivalent discount unilateral minimum time transfer problem to obtain an open-loop optimal strategy pair; according to the optimal control strategy, the state of the spacecraft is subjected to recursion, the state of the escape spacecraft under unknown control is subjected to recursion, and a target point is updated. The real-time closed-loop pursuit game method and device for the elliptical orbit of the spacecraft are suitable for the real-time pursuit game of the elliptical orbit of the spacecraft.

The embodiment of the invention provides a satellite ephemeris updating method and device and spaceborne equipment, and relates to the technical field of aerospace, and the method comprises the steps: obtaining the current position and current space environment parameters of a satellite; searching a grid adjacent to the current position from the plurality of grids, wherein the grid is associated with the position, the space environment parameter interval and the atmospheric density; predicting the atmospheric density of the orbit where the satellite is located by using the position, the space environment parameter interval and the atmospheric density associated with the adjacent grids, the current position and the current space environment parameters; and updating the ephemeris of the satellite by using the predicted atmospheric density. By applying the technical scheme provided by the embodiment of the invention, the prediction precision of the atmospheric density can be improved and the satellite ephemeris prediction precision can be improved under the condition that the computing power and storage of satellite-borne equipment are limited.

The application provides an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite orbiting, comprising the following steps: acquiring a satellite-rocket separation orbit and a station parameter, calculating an elevation angle of the satellite and the station, and then determining a range of a true anomaly angle of pulse ignition meeting a measurement and control requirement; in combination with the range of the true anomaly angle of pulse ignition, generating a fuel most-probable pulse orbit change strategy of the satellite from an initial orbit to a target orbit, determining an orbit change speed increment, an ignition position and an ignition frequency; determining ignition-on and ignition-off conditions of each orbit change and a drift circle number between adjacent two orbit changes; according to the pulse orbit change strategy, respectively determining an initial value of an ignition time length, an ignition time and a thrust direction of each orbit change, and determining the final ignition time length, the ignition time and the thrust direction of each orbit change. The application adopts a non-apogee double-pulse orbit change strategy to design a fuel optimal orbit transfer strategy of a large-ellipse frozen orbit satellite orbiting, and solves the design method problem of the orbit transfer strategy.

The present invention provides a satellite constellation electric propulsion deorbiting control method and system, which relates to the technical field of spacecraft orbit design and control. The method includes: Step S1: According to the deorbiting mission requirements, select the characteristic point height of the deorbiting orbit, and calculate the semi-major axis and eccentricity of the target orbit; Step S2: Adjust the semi-major axis of the orbit, continuously fire along the tangential direction until the semi-major axis is reduced to the target value; Step S3: Adjust the eccentricity of the orbit, fire throughout the orbit, and fire along the direction perpendicular to the apsidal line until the eccentricity meets the target value; Step S4: Fire at the orbit characteristic point until the perigee height and apogee height meet the deorbiting requirements. The present invention can achieve rapid evaluation and optimization of the deorbiting orbit parameter set; by adopting the strategy of firing in different coordinate systems in stages, the coupling between the semi-major axis and the eccentricity is reduced, which is beneficial to simplifying the complexity of the actual mission deorbiting scheme.

In an orbital attitude control device (1150), an ideal thrust axis direction calculator (1505) calculates an ideal thrust axis direction based on information of a predetermined orbit, an ideal attitude calculator (1506) calculates an ideal attitude of the satellite based on the ideal thrust axis direction and a solar direction, and a control torque calculator (1510) calculates an ideal control torque that makes the attitude of the satellite follow the ideal attitude and a torque restraint plane in which the solar direction is orthogonal to a rotational axis of the solar array panel, defines an evaluation function obtained by weighting a distance from the ideal control torque and a distance from the torque restraint plane and then summing the weighted distances, and calculates the control torque that allows the drive constraint to be satisfied and the evaluation function to be minimized.

The invention provides an on-satellite autonomous task planning method, device and equipment for earth target imaging, and relates to the technical field of spaceflight, and the method comprises the steps: carrying out the fitting of each circle of strict regression orbit of an earth observation satellite, and obtaining a target function corresponding to each circle of strict regression orbit; constructing a virtual orbit passing through the target observation point based on the position information of the specified target observation point and the orbit parameters of each circle of strict regression orbit so as to determine a corresponding latitude depression angle sequence when the earth observation satellite runs on each circle of virtual orbit and passes through the target observation point; determining an initial reference orbit point based on the latitude depression angle sequence through the target function, and performing multiple times of compensation on the initial reference orbit point to obtain a target reference orbit point; and if the target reference orbit point satisfies the preset observation constraint, executing a ground target imaging task. According to the method, the satellite autonomy and the task planning real-time performance in the process of planning the earth target imaging task can be remarkably improved.

The invention discloses a strict regression frozen orbit design method. The method comprises the following steps: calculating an initial estimated value of a target strict regression frozen orbit mean element number based on an analytic formula of a regression orbit and a frozen orbit under a third-order gravity field; under the high-order gravity field dynamic model, iteratively correcting the orbit semi-major axis and the orbit inclination angle based on the right ascension change rate of the ascending node and the deviation value of the longitude of the revisit ascending node; taking the geocentric vector difference of the revisit ascending node as an optimization target, taking the eccentricity rate as an optimization variable, and adopting an optimization algorithm to obtain a current optimal frozen eccentricity rate; iteratively correcting the semi-major axis of the orbit and the inclination angle of the orbit, and optimizing the eccentricity ratio to obtain the orbital number of the strictly frozen regression orbit.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

To provide an orbit estimation apparatus, an orbit estimation method and an orbit estimation program for an artificial celestial body, which can accurately estimate an orbit element, even by estimating a short circular arc through an astronomical survey.SOLUTION: An orbit estimation method for a plurality of artificial celestial bodies, which performs orbit estimation by Bayes' estimation using MCMC, on the basis of an astronomical survey, comprises: a low-accuracy posterior distribution calculating step of obtaining a low-accuracy posterior distribution of orbit elements and sizes of celestial bodies from short circular arcs of respective orbits of a plurality of artificial celestial bodies obtained through an astronomical survey; an artificial celestial body identifying step of identifying a survey of the same artificial celestial body by comparing the posterior distributions with each other, with respect to different surveys, of surveys through the astronomical survey; and a high-accuracy posterior distribution calculating step of obtaining a high-accuracy posterior distribution of the orbit elements using surveys of three or more short circuit arcs of the same artificial celestial body.SELECTED DRAWING: Figure 5

The application discloses a kind of space target ultra-short arc initial orbit intelligent determination method, belong to aerospace technology field.The application is based on artificial intelligence technology to realize space target ultra-short arc initial orbit determination, there is no strict requirement to complex nonlinear problem, and can develop more complex perturbation condition and initial orbit determination under the operation law of multiple body, coverage is wider;The application generates multiple groups of target orbits randomly and calculates the observation and state quantity of each visible arc segment, constructs the sample data set with ultra-large capacity, which is helpful for neural network to learn the nonlinear relationship between observation and state quantity using rich samples;The application builds and trains long short-term memory network model, determines the optimal network model weight parameter by designing and adjusting hyperparameter, can obtain the fast prediction solution of unknown space target ultra-short arc initial orbit according to observation;The application can meet more space target ultra-short arc initial orbit determination tasks by considering different inputs and outputs.

The invention discloses a space-based angle-measurement-only target orbit determination method for compensating orbit position errors, and belongs to the technical field of satellite application measurement. The method comprises the following steps: establishing an error model of an empirical orbit dynamics model error introduction measurement error according to a real orbit dynamics model of a target; training the initial neural network according to a preset training sample to obtain an orbit position error estimation model; wherein the training sample is determined through historical orbit data and an empirical orbit dynamics model; and according to the orbit position error estimation model, compensating the error model, and processing the initial state of a to-be-measured target by using a deviation compensation least square method to obtain an optimal estimation value of the orbit of the to-be-measured target. The method can effectively improve the determination precision of the target track.

The invention provides a method for determining an autonomous orbit of a high-orbit satellite based on combination of inter-satellite relative position vector measurement and a GNSS (Global Navigation Satellite System), and belongs to the field of autonomous navigation. The method comprises the following specific steps: constructing a GEO orbit (geostationary orbit) environment of a satellite, and establishing a gravity model by using an EGM2008 model; establishing a relative position vector measurement model between the spacecrafts and a relative distance measurement model between the spacecrafts and the GNSS satellites; analyzing and fusing the system and measurement errors; performing linearization processing on the measurement equation; and estimating the state of the high-orbit spacecraft by adopting an extended Kalman filter, and determining the absolute position and speed of the spacecraft. On the basis of determining the orbit of the spacecraft by using a GNSS satellite traditionally, direction finding measurement is added between high-orbit spacecrafts, so that the orbit determination precision of the high-orbit spacecrafts is improved, and high-precision orbit determination is realized.

The utility model discloses an assembled astronavigation star orbit, and relates to the field of simulation models. The split mounting type astronavigation star orbit comprises a simulation star orbit unit, a driving unit is installed in the simulation star orbit unit, and the simulation star orbit unit is matched with the driving unit to be in transmission connection with a simulation celestial body unit; the simulation star orbit unit comprises a bottom plate and a simulation orbit, a plurality of clamping plates are clamped to the outer side of the bottom plate, and clamping grooves are formed in the clamping plates; the outer edge of the simulation track is attached to the inner walls of the multiple clamping grooves, and the simulation track is in transmission connection with the driving unit. According to the split mounting type astronavigation star orbit, in the using process, the simulation orbit can be mounted only by clamping the multiple clamping plates to the outer side of the bottom plate and enabling the outer edge of the simulation orbit to be attached to the inner walls of the multiple clamping grooves, at the moment, the driving unit and the simulation celestial body unit are correspondingly mounted, and then overall mounting of the astronavigation star orbit can be completed; installation is simple, and the device is suitable for users of different ages.

The application relates to a design method of a moon global positioning constellation based on a lunar space DRO reference orbit, wherein the initial orbit state and the orbit period of any DRO reference orbit in a DRO reference orbit to be deployed are determined based on the initial phase of the DRO reference orbit; the initial orbit state of any satellite to be deployed on the initial phase of the DRO reference orbit is determined based on the orbit state corresponding to the target phase; the orbit period of the any satellite is determined based on the orbit period of the any DRO reference orbit; the target orbit state corresponding to the any satellite is obtained based on the orbit period of the any satellite and the initial orbit state of the any satellite; and the moon global positioning constellation under a four-body model is constructed based on the target orbit state corresponding to each satellite. The moon global positioning constellation constructed by the application has higher stability and higher moon surface coverage capacity compared with the existing near-moon constellation.

An object is to appropriately manage disclosure of information on the orbit of a satellite constellation. An information management device is installed in at least one of a plurality of management business devices each of which conducts a management business for a plurality of space objects flying in space. A storage unit stores orbit forecast information including an orbit forecast, which is a forecast value of an orbit of each of the plurality of space objects, and a forecast error that is forecast for the orbit. An information disclosure unit determines whether the orbit forecast information is to be disclosed to a different information management device, based on a disclosure threshold for determining whether the orbit forecast information is to be disclosed and the forecast error.

The present invention discloses a method for designing a companion satellite orbit for space station aimed at early warning of space debris threats, and belongs to the technical field of aerospace. The present invention provides a method for designing a companion satellite orbit for space station aimed at early warning of space debris threats, which writes the CW equation into a parameterized form by combining the parametric equation of the space ellipse, and then uses the geometric meaning of the parameters in combination with the azimuth and distance constraints to give the orbit design result. The design method proposed in the present invention makes full use of the geometric characteristics and mathematical geometry methods of the CW equation, and has the advantages of being intuitive, having obvious physical meaning, and being easy to calculate; at the same time, the present invention supplements the design method of the flyby orbit for early warning of space debris impact in the existing field, and provides ideas, references and technical support for subsequent references in engineering fields.

This invention discloses a method and apparatus for planning a timed, fixed-point landing and timed takeoff for lunar sampling missions. The method includes: determining the orbital root number and orbital insertion time of the orbit in the first control phase of the lunar sampling mission, as well as the nominal parameters for timed, fixed-point landing and timed takeoff. Based on the nominal parameters, orbital root number, and orbital insertion time, the first control phase is jointly planned with subsequent control phases to obtain a first deviation, and the first deviation is iterated based on a first variable. When the first deviation is less than a convergence threshold, the target variable for the first control phase is determined, allowing the satellite to achieve timed, fixed-point landing and timed takeoff based on the target variables of each control phase. This achieves dynamic joint planning with timed, fixed-point landing and timed takeoff as the final goal for each control phase, effectively improving the mission's control reconfiguration capability, avoiding additional control operations, and reducing resource consumption.

This invention relates to the field of collision warning technology, specifically to a spacecraft collision warning method for a large number of space targets. The method includes: selecting a multi-data parallel screening warning approach; screening secondary target orbits within the TLE data set of the complete collision target orbits using set pre-screening thresholds for date expiration, apogee / perigee pre-screening safety margins, orbital intersection pre-screening thresholds, and time pre-screening thresholds; performing SGP4 orbit prediction on the primary target orbit and the initially screened secondary target orbits; obtaining the minimum spatial distance and collision probability between spacecraft within the primary and secondary target orbits; summarizing the primary and secondary target orbit numbering information, as well as the minimum spatial distance, collision probability, and the time corresponding to the minimum spatial distance, into a collision warning result; and summarizing all collision warning results to output a collision warning set information. This invention can filter out targets that are unlikely to intersect with the spacecraft to the greatest extent possible, reducing the computational load of subsequent orbit prediction.

The invention discloses a borehole trajectory generation method and system, electronic equipment and a storage medium, and relates to the technical field of borehole trajectory designing.The method comprises the steps that on the basis of six-section type borehole trajectory design, a space arc method and a cylindrical spiral method are combined to generate a borehole trajectory; combining an orbit constraint equation with to-be-optimized parameters and a vertical depth position and depth used for limiting a stratum, taking the shortest total length of a borehole orbit as a target, and optimizing an initial borehole orbit by utilizing a heuristic algorithm; and judging whether the appointed stratum is overlapped with the deflecting section of the initial borehole trajectory or not to obtain a judgment result, and determining the initial borehole trajectory as a final borehole trajectory if the appointed stratum is not overlapped with the deflecting section of the initial borehole trajectory. According to the method, on the basis of the six-section type borehole trajectory design, the stratum factors can be comprehensively considered through the mode of matching the trajectory constraint equation, the initial borehole trajectory is optimized through the heuristic algorithm, the design period of the borehole trajectory can be effectively shortened, and the construction difficulty and the construction cost are reduced.