Planning method and device for launch window and launch point location of carrier rocket

Abstract

A method and device for planning a launch window and a launch point location of a carrier rocket relate to the technical field of launch of carrier rockets, and the method comprises the following steps: calculating the position of a satellite at each moment according to basic parameters of a target constellation under a J2000 coordinate system; after position information of satellites is converted into geodetic coordinates, sub-satellite point trajectories of the satellites are calculated according to the geodetic coordinates, time windows of the satellites passing through a launching site are calculated in combination with the center position and the maneuvering radius of the launching site, and key data are provided for subsequent planning of launching windows and launching point locations of a carrier rocket. The model of the carrier rocket, the time window and the sub-satellite point track of the satellite are comprehensively considered, the most suitable launching window and launching point location are obtained through iterative calculation, the tedious process of reversely deducing the launching window only by depending on the fixed launching point location and ascending node / descending node information is avoided, the modeling and calculation process is greatly simplified, and the modeling efficiency is improved. And the execution efficiency of the launching task is obviously improved.

Description

Technical Field

[0001] The present application relates to the field of launch vehicle launch technology, and in particular to a method and device for planning a launch window and launch point of a launch vehicle. Background Art

[0002] In recent years, large-scale low-orbit small satellite constellations have great value in a series of application fields such as space-based global communications and remote sensing. Large-scale low-orbit constellation launch plans have emerged in an endless stream, showing an explosive development trend. Rapid response trajectory design is the basis and prerequisite for improving emergency launch capabilities and emergency large-scale satellite networking efficiency. Without the support of rapid response trajectory, emergency launch or networking launch missions cannot be completed quickly and efficiently.

[0003] In the face of emergency launch missions and large-scale constellation deployment missions, the orbital plane of the target orbit and the position within the orbital plane may have specific requirements. The traditional trajectory planning algorithm needs to reversely calculate the launch window based on the given fixed launch point and ascending node / descending node information. The modeling and calculation process is complicated and it is impossible to quickly plan the launch window and launch point according to the requirements of the launch mission. It is difficult to adapt to emergency launch missions with phase constraints or large-scale constellation networking launch missions. Summary of the invention

[0004] The present application provides a method and device for planning a launch window and a launch point for a carrier rocket, which can quickly plan the launch window and the launch point for the carrier rocket according to the basic parameters of the target constellation and the range information of the launch site.

[0005] In a first aspect, an embodiment of the present application provides a method for planning a launch window and a launch point of a carrier rocket, the method comprising:

[0006] In the J2000 coordinate system, according to the basic parameters of the target constellation, the orbital elements of each satellite of the target constellation at different times are calculated, and the positions of each satellite at different times are calculated according to the orbital elements;

[0007] Convert the positions of each satellite at different times into geodetic coordinates;

[0008] Calculate the sub-satellite point trajectory of each satellite based on the geodetic coordinates of each satellite at different times;

[0009] For any satellite in the target constellation, the time window of the satellite passing through the launch site is calculated according to the sub-satellite point trajectory of the satellite, the central position of the launch site and the maneuvering radius;

[0010] The launch window and launch point of the launch vehicle are iteratively calculated according to the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite.

[0011] In combination with the first aspect, in one implementation, calculating the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation includes:

[0012] The basic parameters of the target constellation include the number of satellites, the number of orbital planes, relative phase parameters, orbital inclination and orbital altitude;

[0013] According to the basic parameters of the target constellation, the preset initial value of the right ascension of the ascending node, the preset initial value of the argument of perigee and the preset initial value of the true anomaly, the orbital elements of each satellite of the target constellation at different times are calculated.

[0014] In combination with the first aspect, in one implementation, calculating the position of each satellite at different times according to the orbital elements includes:

[0015] The orbital elements include orbital semi-major axis, orbital eccentricity, orbital inclination, right ascension of ascending node, argument of perigee and true anomaly;

[0016] Calculating a first matrix and a second matrix according to the ascending node right ascension, orbit inclination and perigee argument;

[0017] The position of each satellite at different times is calculated according to the orbit semi-major axis, orbit eccentricity, true anomaly, the first matrix and the second matrix.

[0018] In combination with the first aspect, in one implementation, converting the position of each satellite at different times into geodetic coordinates includes:

[0019] The position of each satellite at different times is converted from the J2000 coordinate system to the geocentric coordinate system to obtain the geocentric rectangular coordinates of each satellite at different times;

[0020] Convert the geocentric rectangular coordinates of each satellite at different times into geodetic coordinates.

[0021] In combination with the first aspect, in one implementation, for any satellite in the target constellation, calculating the time window for the satellite to pass through the launch site according to the sub-satellite point trajectory of the satellite, and the central position and maneuvering radius of the launch site includes:

[0022] Calculate the shortest distance between the center position of the launch site and the sub-satellite point track of the satellite;

[0023] Determine whether the launch mission is feasible based on the shortest distance and the maneuvering radius of the launch site;

[0024] If so, the time difference between the sub-satellite point trajectory of the satellite passing through the boundaries of the maneuvering launch range of the launch site is the time window for the satellite to pass through the launch site; the launch range is calculated based on the center position of the launch site and the maneuvering radius.

[0025] In combination with the first aspect, in one implementation, judging whether the launch mission is launch-feasible based on the shortest distance and the maneuvering radius of the launch site includes:

[0026] If the maneuvering radius of the launch site is greater than or equal to the shortest distance, the launch mission is feasible;

[0027] If the maneuvering radius of the launch site is smaller than the shortest distance, the launch mission is not feasible.

[0028] In combination with the first aspect, in one implementation, iteratively calculating the launch window and launch point of the launch vehicle according to the model of the launch vehicle, the time window, and the sub-satellite point trajectory of the satellite includes:

[0029] Calculating the launch window of the carrier rocket according to the time window and the orbital parameter requirements of this launch mission;

[0030] Inputting the model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket;

[0031] According to the launch trajectory data of the launch vehicle and the motion parameters of the satellite, the launch point position and orbit entry point parameters of the launch vehicle are adjusted to reduce the geographic latitude and longitude deviation of the sub-satellite point corresponding to the orbit entry point of the launch vehicle and the satellite;

[0032] Determine whether the current launch window and launch point meet the launch requirements. If so, use the current launch window and launch point as the launch window and launch point of the carrier rocket; if not, return to the step of calculating the launch window of the carrier rocket based on the time window and the sub-satellite point trajectory of the satellite.

[0033] In combination with the first aspect, in one implementation, calculating the launch window of the carrier rocket according to the time window and the orbital parameter requirements of this launch mission includes:

[0034] Determine the flight duration and pre-set duration based on the orbital parameter requirements of this launch mission;

[0035] The launch window of the carrier rocket is calculated according to the flight duration, the advance preset duration and the time window.

[0036] In combination with the first aspect, in one implementation, the model of the carrier rocket, the time window, and the sub-satellite point trajectory of the satellite are input into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket, including:

[0037] The model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite are input into a preset solid carrier rocket flight trajectory planning algorithm, and planning calculations are performed in combination with the orbital parameter requirements, wind field conditions, performance of the carrier rocket, ground measurement and control constraints and landing area safety requirements of this launch mission to obtain the launch trajectory data of the carrier rocket.

[0038] In a second aspect, an embodiment of the present application provides a launch window and launch point planning device for a carrier rocket, the device comprising:

[0039] A position calculation module is used to calculate the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation in the J2000 coordinate system, and calculate the position of each satellite at different times according to the orbital elements;

[0040] A coordinate conversion module, used to convert the position of each satellite at different times into geodetic coordinates;

[0041] The time window calculation module is used to calculate the sub-satellite point trajectory of each satellite according to the geodetic coordinates of each satellite at different times; and is also used to calculate the time window of the satellite passing through the launch site for any satellite in the target constellation according to the sub-satellite point trajectory of the satellite, the center position of the launch site and the maneuvering radius;

[0042] The planning module is used to iteratively calculate the launch window and launch point of the launch vehicle according to the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite.

[0043] The beneficial effects brought by the technical solution provided in the embodiments of the present application include:

[0044] This application calculates the orbital elements of each satellite at different times according to the basic parameters of the target constellation in the J2000 coordinate system, and further determines its position, accurately simulates and predicts the satellite orbit. After converting the satellite's position information into geodetic coordinates, the subsatellite point trajectory of each satellite is calculated according to the geodetic coordinates, and the time window for the satellite to pass through the launch site is calculated in combination with the central position and maneuvering radius of the launch site, providing key data for the subsequent planning of the launch window and launch point of the launch vehicle. Finally, taking into account the model of the launch vehicle, the time window and the subsatellite point trajectory of the satellite, the most suitable launch window and launch point are obtained through iterative calculation, avoiding the cumbersome process of reversely deriving the launch window by relying solely on fixed launch points and ascending node / descending node information, greatly simplifying the modeling and calculation process, and significantly improving the execution efficiency of the launch mission. BRIEF DESCRIPTION OF THE DRAWINGS

[0045] Figure 1 A flowchart of a method for planning a launch window and a launch point for a carrier rocket according to an embodiment of the present application;

[0046] Figure 2 This is a schematic diagram of the relative relationship between rocket, satellite and earth in the embodiment of the present application;

[0047] Figure 3 A schematic diagram showing how the shortest distance between the center position of the launch site and the sub-satellite point trajectory of the satellite changes with the movement time in an embodiment of the present application;

[0048] Figure 4 This is a schematic diagram of the launch feasibility analysis when the satellite orbit is raised and crosses the maneuverable launch range in the embodiment of the present application;

[0049] Figure 5 This is a schematic diagram of the launch feasibility analysis when the satellite descends to the orbit and crosses the mobile launch range in the embodiment of the present application;

[0050] Figure 6 This is a basic workflow diagram for the launch vehicle flight trajectory planning of the embodiment of the present application;

[0051] Figure 7 This is a detailed flow chart of step S64 in the embodiment of the present application;

[0052] Figure 8 This is a schematic diagram of the structure of a launch window and launch point planning device for a carrier rocket according to an embodiment of the present application. DETAILED DESCRIPTION

[0053] In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0054] First, please refer to Figure 1 , Figure 1 The flowchart of the launch window and launch point planning method of the carrier rocket of the embodiment of the present application is as follows:

[0055] Step S1: In the J2000 coordinate system, according to the basic parameters of the target constellation, the orbital elements of each satellite in the target constellation at different times are calculated.

[0056] Step S2: Calculate the position of each satellite at different times according to the orbital elements of each satellite in the target constellation at different times.

[0057] Step S3: Convert the position of each satellite at different times into geodetic coordinates.

[0058] Step S4: Calculate the sub-satellite point trajectory of each satellite according to the geodetic coordinates of each satellite at different times.

[0059] Step S5: For any satellite in the target constellation, the time window of the satellite passing through the launch site is calculated according to the sub-satellite point trajectory of the satellite, the central position of the launch site and the maneuvering radius.

[0060] Step S6: Iteratively calculate the launch window and launch point of the launch vehicle according to the model of the launch vehicle, the time window of the satellite passing through the launch site and the sub-satellite point trajectory of the satellite.

[0061] This method calculates the orbital elements of each satellite at different times according to the basic parameters of the target constellation in the J2000 coordinate system, and further determines its position, accurately simulates and predicts the satellite orbit. After converting the satellite's position information into geodetic coordinates, the sub-satellite point trajectory of each satellite is calculated according to the geodetic coordinates. Combined with the central position and maneuvering radius of the launch site, the time window of the satellite passing through the launch site is calculated, providing key data for the subsequent planning of the launch window and launch point of the launch vehicle. Finally, considering the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite, the most suitable launch window and launch point are obtained through iterative calculation, avoiding the tedious process of reversely deriving the launch window by relying solely on fixed launch points and ascending / descending node information, greatly simplifying the modeling and calculation process, and significantly improving the execution efficiency of the launch mission.

[0062] It should be noted that, considering the important application value of low-Earth orbit and the limited carrying capacity of solid-propellant launch vehicles, the target orbit of satellites carried by solid-propellant launch vehicles is usually designed as low-Earth orbit. For low-Earth orbit, its eccentricity is approximately zero. Therefore, the orbital parameters that need to be designed are mainly the semi-major axis, orbital inclination and ascending node right ascension, among which the ascending node right ascension is determined by the parameters of the target orbit or target constellation.

[0063] For launch missions with both phase and orbital parameter constraints, the launch window and launch point planning of the launch vehicle must meet the following two conditions at the same time:

[0064] Plane window condition: the launch point needs to pass through the orbital plane.

[0065] Phase window condition: the satellite needs to reach a specific phase point in a certain orbital plane after entering orbit.

[0066] For typical missions launched from domestic launch sites, please refer to Figure 2 , Figure 2 This is a schematic diagram of the relative relationship between rocket, satellite and earth in the embodiment of this application. Figure 2 In the example, assume that the longitude and latitude coordinates of the transmitting point B are The latitude and longitude coordinates of the point below the arrow corresponding to the launch vehicle's orbital entry point C are: Assuming that the launch vehicle trajectory design takes into account the safety restrictions of the landing area, the launch mission is carried out using the descending orbit launch method (the rocket is launched from north to south). S is the current moving point of the target satellite, and the longitude and latitude coordinates of the sub-satellite point are The position of the target satellite at the time of launch of the carrier rocket is S'. At the time of rocket orbit entry, the position of the rocket orbit entry point C coincides with the position of the satellite's current motion point S. At the same time, the speed and position of the rocket at the time of orbit entry meet the orbital semi-major axis R of the target orbit. bz , speed V bz and inclination angle i bz Size constraints.

[0067] That is, after the launch window and launch point of the carrier rocket are calculated through the above steps S1-S6, assuming that the total flight time of the rocket is tr, the ignition time of the rocket is T0, and the target satellite operation time corresponding to the rocket entering orbit is T1, the following relationship must be satisfied in the end:

[0068]

[0069] In formula (1), V is the velocity of the rocket at the time of entering orbit, H is the altitude of the rocket at the time of entering orbit, and θ is the orbital inclination at the time of entering orbit.

[0070] In some embodiments, in the above step S1, calculating the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation includes the following steps:

[0071] According to the basic parameters of the target constellation, the preset initial value of the right ascension of the ascending node, the preset initial value of the argument of perigee and the preset initial value of the true anomaly, the orbital elements of each satellite of the target constellation at different times are calculated. Among them, the basic parameters of the target constellation include the number of satellites, the number of orbital planes, relative phase parameters, orbital inclination and orbital altitude.

[0072] Assume that the target constellation consists of T satellites, each of which is in a circular orbit with the same semi-major axis a and the same orbital inclination i. The ascending nodes of the P orbital planes are evenly distributed on the equator; in each orbital plane, the satellites are evenly distributed; satellites in adjacent orbital planes have a certain phase relationship, which is represented by the relative phase parameter F.

[0073] In summary, the basic parameters of the target constellation are: the number of satellites T, the number of orbital planes P, the relative phase parameter (also known as the phase factor) F, the orbital inclination i and the orbital altitude h.

[0074] The number of satellites in each orbital plane is S=T / P, so the target constellation can also be described as S / P / F.

[0075] After obtaining the basic parameters of the target constellation, the orbital elements of each satellite in the target constellation at different times can be calculated. The parameters that need to be input include T, P, F, a, i and e (orbital eccentricity, the orbital eccentricity of the circular orbit can be approximated to 0). In addition, the following parameters should also be given during the calculation process:

[0076] The initial value of the right ascension of the ascending node of the first orbital plane (i.e. the preset initial value of the right ascension of the ascending node) Ω01, the initial value of the argument of perigee of the first satellite in the first orbital plane (i.e. the preset initial value of the argument of perigee) ω011, and the initial value of the true anomaly of the first satellite in the first orbital plane (i.e. the preset initial value of the true anomaly) f011.

[0077] For a circular orbit, ω011=u011 (u011 is the latitude argument of the first satellite in the first orbital plane), f011=0.

[0078] In addition, during the calculation process, the following relationship can be inferred based on the above parameters:

[0079] When a satellite in the kth plane is located at the ascending node (latitude argument u = 0), the satellite closest to the satellite in the j+1th plane has a latitude argument u = ΔΦ (ΔΦ is the phase difference), ΔΦ = F (360° / T), and the possible value of F is 0 to P-1. The interval of the ascending node right ascension ΔΩ = 360° / P, and the interval of the satellite positions in the orbital plane Δθ = 360° / S.

[0080] Through the above calculation method, multiple satellite orbits in the target constellation can be calculated simultaneously, which speeds up the calculation efficiency and is more adaptable to the mission requirements of emergency launch and large-scale constellation networking launch.

[0081] In some embodiments, in the above step S2, the orbital elements include the orbital semi-major axis, the orbital eccentricity, the orbital inclination, the right ascension of the ascending node, the argument of perigee and the true anomaly, and calculating the position of each satellite at different times according to the orbital elements of each satellite of the target constellation at different times includes the following steps:

[0082] S21: Calculate the first matrix and the second matrix according to the right ascension of the ascending node, the orbit inclination and the argument of perigee of each satellite in the target constellation at different times.

[0083] S22: Calculate the position of each satellite at different times according to the orbital semi-major axis, orbital eccentricity, true anomaly, first matrix and second matrix of each satellite in the target constellation at different times.

[0084] In the above step S21, the formulas for calculating the first matrix and the second matrix are as follows:

[0085]

[0086] In formula (2), the right ascension of the ascending node is Ω, the orbit inclination is i, the argument of perigee is ω, and the first matrix is ​​P. In formula (3), the second matrix is ​​Q.

[0087] In the above step S22, the position of each satellite at different times is calculated according to the orbital semi-major axis, orbital eccentricity, true anomaly, first matrix and second matrix of each satellite of the target constellation at different times. The specific formula is as follows:

[0088]

[0089] In formula (4), is the position component of the satellite in the J2000 coordinate system, p=a(1-e 2 ), a is the semi-major axis of the orbit, e is the eccentricity of the orbit, and f is the true anomaly.

[0090] In addition, in order to more accurately describe the satellite's motion attitude, the satellite's velocity component in the J2000 coordinate system should also be calculated. The formula is as follows:

[0091]

[0092] In formula (5), is the velocity component of the satellite in the J2000 coordinate system, μ is the gravitational constant of the Earth, and the other mathematical symbols in the formula have the same meaning as in the previous text and will not be repeated here.

[0093] In some embodiments, in the above step S3, converting the position of each satellite at different times into geodetic coordinates comprises the following steps:

[0094] S31: Convert the position of each satellite at different times from the J2000 coordinate system to the geocentric coordinate system to obtain the geocentric rectangular coordinates of each satellite at different times.

[0095] S32: Convert the geocentric rectangular coordinates of each satellite at different times into geodetic coordinates.

[0096] In the above step S31, the formula for converting the geocentric rectangular coordinates of each satellite at different times into geodetic coordinates is:

[0097]

[0098] In formula (6), (X, Y, Z) are the geocentric rectangular coordinates, (B, L, H) are the geodetic coordinates, and a e is the mean radius of the Earth, e is the flattening of the Earth, and f is the true anomaly.

[0099] It should be noted that the geocentric coordinate system is a three-dimensional rectangular coordinate system with the center of mass of the earth as its origin, and its coordinate form is convenient for mathematical calculations and spatial analysis. Converting the position of each satellite at different times from the J2000 coordinate system to the geocentric coordinate system can ensure that the position data of all satellites are based on a unified, earth-centered coordinate reference, making it easier to perform complex data processing tasks such as orbit prediction, trajectory analysis, and collision detection, and provide convenience for subsequent data analysis and processing. Geodetic coordinates (longitude, latitude, altitude) are an intuitive and accurate way to describe the position on the surface of the earth. After converting geocentric rectangular coordinates to geodetic coordinates, it is easier to obtain the position information of the satellite relative to the surface of the earth.

[0100] In some embodiments, in the above step S5, for any satellite in the target constellation, according to the sub-satellite point trajectory of the satellite, the central position of the launch site and the maneuvering radius, the time window of the satellite passing through the launch site is calculated, including the following steps:

[0101] S51: Calculate the shortest distance between the center position of the launch site and the sub-satellite point trajectory of the satellite.

[0102] S52: Based on the above shortest distance and the maneuvering radius of the launch site, determine whether the launch mission is feasible. If so, proceed to step S53; if not, generate a report on the feasibility of the launch based on the judgment process.

[0103] S53: Calculate the time difference between the sub-satellite point trajectory of the satellite passing through the boundaries of the maneuverable launch range of the launch site as the time window for the satellite to pass through the launch site. The maneuverable launch range is calculated based on the center position of the launch site and the maneuverable radius.

[0104] In the above step S51, the formula for calculating the distance between the center position of the launch site and any sub-satellite point on the sub-satellite point trajectory of the satellite is as follows:

[0105] distance = L[(BFSC, LFSC), (B, L)] Formula (7)

[0106] In formula (7), (BFSC, LFSC) is the center position of the launch site, B is the latitude of the sub-satellite point, L is the longitude of the sub-satellite point, L[] is used to calculate the geodetic distance between two points, and distance is the distance between the center position of the launch site and any sub-satellite point on the sub-satellite point trajectory of the satellite.

[0107] When the satellite moves along the subsatellite point trajectory, the distance will change with the passage of time. For specific changes, please refer to Figure 3 , Figure 3 This is a schematic diagram of the shortest distance between the center position of the launch site and the sub-satellite point trajectory of the satellite changing with the movement time according to an embodiment of the present application.

[0108] exist Figure 3 In the figure, the vertical axis d represents the distance between the center of the launch site and the subsatellite point at the current moment, and the horizontal axis t represents the movement time of the satellite. When the movement time of the satellite is t 1 When min It is the shortest distance between the center of the launch site and the sub-satellite trajectory of the satellite.

[0109] In some embodiments, in the above step S52, whether the launch mission is feasible is determined based on the above shortest distance and the maneuvering radius of the launch site, and the specific steps are as follows:

[0110] If the maneuvering radius of the launch site is greater than or equal to the above-mentioned shortest distance, the launch mission is feasible.

[0111] If the maneuvering radius of the launch site is smaller than the above minimum distance, the launch mission will not be feasible.

[0112] For specific judgment rules, please refer to formula (8):

[0113]

[0114] In formula (8), R is the maneuvering radius of the launch site. When R ≥ distance min When R <distance min At this time, the launch mission was not feasible.

[0115] In some embodiments, when the satellite orbit is raised and crosses the maneuverable launch range, to determine whether the launch mission is feasible, please refer to Figure 4 , Figure 4 This is a schematic diagram of the launch feasibility analysis when the satellite orbit is raised and crosses the maneuverable launch range in an embodiment of the present application.

[0116] exist Figure 4In the figure, the intersection points of the satellite's sub-satellite point track (sub-satellite point trajectory) and the boundaries of the launch site's maneuverable launch range are point A and point B respectively, O is the center point of the maneuverable launch range, and R is the maneuvering radius of the launch site. When R is greater than or equal to the shortest distance between point O and the sub-satellite point track, it means that the launch point can be deployed within the maneuverable launch range of the launch site, which means that this launch mission has the feasibility of launch; when R is less than the shortest distance between point O and the sub-satellite point track, it means that the launch point cannot be deployed within the maneuverable launch range of the launch site, which means that this launch mission does not have the feasibility of launch.

[0117] In some embodiments, when the satellite descends through the maneuverable launch range, to determine whether the launch mission is feasible, please refer to Figure 5 , Figure 5 This is a schematic diagram of the launch feasibility analysis when the satellite in the embodiment of the present application descends to the orbit and crosses the mobile launch range.

[0118] exist Figure 5 In the figure, the intersection points of the subsatellite point track line (subsatellite point trajectory) and the two sides of the launch site's maneuverable launch range are point A and point B, O is the center point of the maneuverable launch range, and R is the maneuvering radius of the launch site. The process of inferring the feasibility of launch is the same as when the satellite orbit is raised and crosses the maneuverable launch range, so it will not be repeated here.

[0119] In the above step S53, the time difference between the sub-satellite point trajectory of the satellite passing through the boundaries of both sides of the maneuverable launch range of the launch site is calculated as the time window for the satellite to pass through the launch site, specifically:

[0120] Assume that the intersection points of the satellite's subsatellite point trajectory and the boundaries of the launch site's maneuverable launch range are point A and point B, respectively. Record the corresponding UTC time (YR1_MO1_DY1_HR1_MI1_SE1) when the satellite passes through point A and the corresponding UTC time (YR2_MO2_DY2_HR2_MI2_SE2) when the satellite passes through point B, as well as the geographical locations of point A (B1, L1) and point B (B2, L2). The time interval between the UTC times corresponding to point A and point B is the time window for the satellite to pass through the launch site.

[0121] In some embodiments, in the above step S6, the launch window and the launch point of the launch vehicle are iteratively calculated according to the model of the launch vehicle, the time window of the satellite passing through the launch site, and the sub-satellite point trajectory of the satellite, including the following steps:

[0122] S61: Calculate the launch window of the carrier rocket according to the time window of the satellite passing through the launch site and the orbital parameter requirements of this launch mission.

[0123] S62: Input the model of the carrier rocket, the time window of the satellite passing through the launch site, and the sub-satellite point trajectory of the satellite into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket.

[0124] S63: According to the launch trajectory data of the launch vehicle and the motion parameters of the satellite, the launch point position and orbit entry point parameters of the launch vehicle are adjusted to reduce the geographic latitude and longitude deviation of the sub-satellite point corresponding to the launch vehicle and the orbit entry point of the satellite.

[0125] S64: Determine whether the current launch window and launch point meet the launch requirements. If so, use the current launch window and launch point as the launch window and launch point of the carrier rocket; if not, return to step S61.

[0126] In some embodiments, in the above step S61, the launch window of the carrier rocket is calculated according to the time window of the satellite passing through the launch site and the orbital parameter requirements of this launch mission, including the following steps:

[0127] S611: Determine the flight duration and advance preset duration based on the orbital parameter requirements of this launch mission.

[0128] S612: Calculate the launch window of the carrier rocket based on the above flight duration, the advance preset duration and the time window of the satellite passing through the launch site.

[0129] Specifically, in the above step S611, the flight duration is determined according to the orbital altitude and orbital inclination of this launch mission, with reference to Table 1; and the advance preset duration is determined according to the orbital inclination of this launch mission, with reference to Table 2. Table 1 is a flight duration difference table, and Table 2 is an advance preset duration difference table.

[0130] Table 1 Flight duration interpolation table

[0131]

[0132] Table 2 Preset duration interpolation table

[0133]

[0134] In the above step S612, the specific steps of calculating the launch window of the carrier rocket according to the above flight time, the advance preset time, the time window of the satellite passing through the launch site and the planned time of the launch window are:

[0135] First, convert the start time (YR1, MO1, DY1, HR1, MI1, SE1) of the time window into the corresponding Julian day JD1, and convert the end time (YR2, MO2, DY2, HR2, MI2, SE2) into the corresponding Julian day JD2. The difference between JD1 and JD2 is the first interval time dt1. The calculation formulas for JD1 and JD2 are as follows:

[0136] JD1=DY1-32075+1461*(YR1+4800+(MO1-14) / 12) / 4+367*(MO1-2-(MO1-14) / 12*12) / 12-3*((YR1+4900+(MO1-14) / 12) / 100) / 4-0.5+HR1 / 24.+MI1 / 1440.+SE1 / 86400 formula (9)

[0137] JD2=DY2-32075+1461*(YR2+4800+(MO2-14) / 12) / 4+367*(MO2-2-(MO2-14) / 12*12) / 1 2-3*((YR2+4900+(MO2-14) / 12) / 100) / 4-0.5+HR2 / 24.+MI2 / 1440.+SE2 / 86400 formula (10)

[0138] Subsequently, the second interval time dt2 is determined according to the flight duration and the advance preset duration, and the launch window planning time GHT is calculated according to dt1 and dt2. The calculation formula of GHT is as follows:

[0139] GHT=dt1-dt2 formula (11)

[0140] Subsequently, the new Julian day JD is calculated based on the start time (YR1, MO1, DY1, HR1, MI1, SE1) and CHT, and then the UTC time (Y, M, D, H, Min, Sec) of the launch window of the carrier rocket is calculated based on the Julian day JD and GHT reverse planning. The calculation formula of JD is as follows:

[0141] JD=DY1-32075+1461*(YR1+4800+(MO1-14) / 12) / 4+367*(MO1-2-(MO1-14) / 12*12) / 1 2-3*((YR1+4900+(MO1-14) / 12) / 100) / 4-0.5+HR1 / 24.+MI1 / 1440.+(SE1+GHT) / 86400

[0142] J = int(JD+0.5)

[0143] N = int(4*(J+68569) / 146097)

[0144] L1=J+68569-int((N*146097+3) / 4)

[0145] Y1=int(4000*(L1+1) / 1461001)

[0146] L2=L1-int(1461*Y1 / 4)+31

[0147] M1=int(80*L2 / 2447)

[0148] L3=int(M1 / 11) Formula (12)

[0149] The UTC time of the launch window of the launch vehicle is calculated as follows:

[0150] D=L2-int(2447*M1 / 80)

[0151] M=M1+2-12*L3

[0152] Y=int(100*(N-49)+Y1+L3)

[0153] H = int((JD+0.5-J)*24)

[0154] Min=int(((JD+0.5-J)*24-H)*60)

[0155] Sec=(((JD+0.5-J)*24-H)*60-Min)*60.0 Formula (13)

[0156] In some embodiments, in the above step S62, the model of the carrier rocket, the time window of the satellite passing through the launch site, and the sub-satellite point trajectory of the satellite are input into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket, specifically:

[0157] The model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite are input into a preset solid carrier rocket flight trajectory planning algorithm, and planning calculations are performed in combination with the orbital parameter requirements, wind field conditions, performance of the carrier rocket, ground measurement and control constraints and landing area safety requirements of this launch mission to obtain the launch trajectory data of the carrier rocket.

[0158] It should be noted that, unlike cumbersome mathematical modeling and complex formula deduction and calculation, this embodiment provides a design method and idea that uses standard interface planning and design and can be directly oriented to rocket model objects, while comprehensively considering multiple constraints such as the real-time wind field conditions of the launch site, the safety of the separation body landing area, and the measurement and control requirements, and finally gives the planning and design results. This method can adapt to different orbital parameter requirements, as well as different domestic launch sites and rocket models, and the obtained trajectory planning results can meet the requirements of actual engineering use.

[0159] In a more specific embodiment, the basic workflow of the launch vehicle flight trajectory planning is as follows: Figure 6 As shown, Figure 6 This is a basic workflow diagram for the launch vehicle flight trajectory planning of the embodiment of the present application. In the method proposed in this embodiment, the last stage of the launch vehicle works twice to send the payload from the initial orbital altitude to the target orbital altitude through orbital maneuvers. The basic workflow for the launch vehicle flight trajectory planning of this embodiment specifically includes the following steps:

[0160] Step A01: Analyze the requirements of this launch mission.

[0161] Step A02: Determine all initial input conditions for mobile launch trajectory planning.

[0162] Step A03: Evaluate the launch vehicle's carrying capacity.

[0163] Step A04: Design and calculate the automatic launch trajectory according to the preset solid-propellant rocket flight trajectory planning algorithm.

[0164] Step A05: Determine whether the calculation result of step A04 meets the relevant orbit entry requirements. If so, go to step A06; if not, return to step A04.

[0165] Step A06: Generate the maneuver launch trajectory and flight parameters.

[0166] Step A07: Output the calculation results of the launch vehicle flight trajectory planning.

[0167] In the above step A04, the preset solid launch vehicle flight trajectory planning algorithm is specifically as follows: establish a connection between the target orbit parameter and the control parameter, multiply the deviation of the target orbit parameter by the coefficient to obtain the adjustment amount of the control parameter, select the coefficient based on design experience, and solve the following equation by the Newton iteration method:

[0168] g(x)=f(x)-b=0Formula (14)

[0169] In formula (14), x is a 5×1-dimensional target orbit parameter variable, b is a 5×1-dimensional control parameter variable, f(x) is the mapping from the orbit entry parameter space to the predicted orbit parameter space, and g(x) represents the difference between f(x) and b.

[0170] The specific solution process of formula (14) is as follows:

[0171]

[0172] In formula (15), x k is the approximate root of the kth iteration, x k+1 is the approximate root of the k+1th iteration, g ′ (x k ) is the Jacobian matrix, g ′ (x k ) -1 g ′ (x k ). The finite difference method is used to solve formula (15), that is, a high-order small quantity is applied to each variable, and then a launch orbit integration is performed, and finally the differential of the error variable with respect to the independent variable is calculated to approximate the partial derivative matrix of the launch orbit integral function with respect to the independent variable.

[0173] In some embodiments, in the above step S63, the launch point position and orbit entry point parameters of the launch vehicle are adjusted according to the launch trajectory data of the launch vehicle and the motion parameters of the satellite to reduce the geographic latitude and longitude deviation of the sub-satellite point corresponding to the launch vehicle and the orbit entry point of the satellite, as follows:

[0174] First, according to the launch trajectory data of the carrier rocket, the geographical latitude rgdB and the geographical longitude rgdL of the sub-satellite point corresponding to the launch vehicle's orbital entry point are obtained; according to the motion parameters of the satellite, the geographical latitude rgdB of the sub-satellite point corresponding to the satellite's orbital entry point is obtained. bz and the geographical longitude of the subsatellite point rgdL bz .

[0175] Subsequently, the difference in geographic latitude ΔrgdB and geographic longitude ΔrgdL between the launch vehicle's orbital entry point and the sub-satellite point corresponding to the satellite's orbital entry point is calculated using the following formula:

[0176] ΔrgdB=rgdB-rgdB bz Formula (16)

[0177] ΔrgdL=rgdL-rgdL bz Formula (17)

[0178] When |ΔrgdB|>0.1deg or |ΔrgdL|≥0.1deg, the geographic latitude and longitude of the launch point of the carrier rocket are adjusted. The adjustment amount of the geographic latitude is |ΔrgdB|·k 1 ·R, the adjustment amount of geographic longitude is |ΔrgdL|·k 2 , k 1 Calculate the adjustment function for the geographic latitude of the launch point, k 2 is the adjustment function for calculating the geographic longitude of the launch point, and R is the calculation coefficient for the orbit insertion method.

[0179] k 1 Set according to the launch vehicle's orbit entry method. If the launch vehicle uses the orbit raising launch plan, k 1 =1; if the launch vehicle uses the orbit raising launch plan, then k 1 =-1.

[0180] Usually, k 2 =1.

[0181] R is set according to the direction of movement of the carrier rocket after entering orbit. If the carrier rocket's orbital entry point is located in the southern hemisphere and the direction of movement after entering orbit is from south to north, then R = -1; if the carrier rocket's orbital entry point is located in the southern hemisphere and the direction of movement after entering orbit is from north to south, then R = 1.

[0182] In some embodiments, please refer to Figure 7 , Figure 7 The detailed flow chart of step S64 in the embodiment of the present application is shown in FIG. In the above step S64, judging whether the current launch window and launch point position meet the launch requirements includes the following steps:

[0183] Step S641: Set the initial value of the transmission window.

[0184] Step S642: Setting the initial value of the transmitting point.

[0185] Step S643: Analyze the orbital entry mode of the launch mission (the orbital entry mode is orbit raising or orbit lowering).

[0186] Step S644: Design and calculate the automatic launch trajectory according to a preset solid-propellant rocket flight trajectory planning algorithm.

[0187] Step S645: Determine whether the calculation result meets the relevant orbit entry requirements. If so, proceed to step S646; if not, return to step S644.

[0188] Step S646: Determine whether the calculation result meets the phase constraint of the target satellite. If so, proceed to step S647; if not, return to step S642.

[0189] Step S647: Determine whether the launch point meets the maneuvering range constraint of the launch site. If so, proceed to step S648; if not, return to step S641.

[0190] Step S648: Generate the maneuver launch trajectory and flight parameters.

[0191] Step S649: Output the planning calculation results of the launch window and launch point.

[0192] Second, please refer to Figure 8 , Figure 8 This is a structural schematic diagram of the planning device for the launch window and launch point of the carrier rocket in the embodiment of the present application. The planning device for the launch window and launch point of the carrier rocket provided in this embodiment includes a position calculation module, a coordinate conversion module, a time window calculation module and a planning module.

[0193] The position calculation module is used to calculate the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation in the J2000 coordinate system, and calculate the position of each satellite at different times according to the orbital elements.

[0194] The coordinate conversion module is used to convert the position of each satellite at different times into geodetic coordinates.

[0195] The time window calculation module is used to calculate the sub-satellite point trajectory of each satellite based on the geodetic coordinates of each satellite at different times; it is also used to calculate the time window of a satellite passing through a launch site for any satellite in the target constellation based on the sub-satellite point trajectory of the satellite, the center position of the launch site, and the maneuvering radius.

[0196] The planning module is used to iteratively calculate the launch window and launch point of the launch vehicle according to the model of the launch vehicle, the time window of the satellite passing through the launch site and the sub-satellite point trajectory of the satellite.

[0197] This device calculates the orbital elements of each satellite at different times according to the basic parameters of the target constellation in the J2000 coordinate system through the coordinate conversion module, and further determines its position, accurately simulating and predicting the satellite orbit. The coordinate conversion module converts the satellite's position information into geodetic coordinates, and the time window calculation module calculates the sub-satellite point trajectory of each satellite according to the geodetic coordinates. Combined with the central position and maneuvering radius of the launch site, the time window of the satellite passing through the launch site is calculated, providing key data for the subsequent planning of the launch window and launch point of the launch vehicle. Finally, the planning module comprehensively considers the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite, and obtains the most suitable launch window and launch point through iterative calculation, avoiding the cumbersome process of reversely deriving the launch window by relying solely on fixed launch points and ascending / descending node information, greatly simplifying the modeling and calculation process, and significantly improving the execution efficiency of the launch mission.

[0198] It should be noted that the serial numbers of the above-mentioned embodiments of the present application are only for description and do not represent the advantages or disadvantages of the embodiments.

[0199] The terms "including" and "having" and any variations thereof in the specification and claims of this application and the above-mentioned drawings are intended to cover non-exclusive inclusions. For example, a process, method, system, product or device that includes a series of steps or units is not limited to the listed steps or units, but optionally includes steps or units that are not listed, or optionally includes other steps or units inherent to these processes, methods, products or devices. The terms "first", "second" and "third" are used to distinguish different objects, etc., and do not represent a sequence, nor do they limit "first", "second" and "third" to different types.

[0200] In the description of the embodiments of the present application, "exemplary", "for example" or "for example" are used to indicate examples, illustrations or descriptions. Any embodiment or design described as "exemplary", "for example" or "for example" in the embodiments of the present application should not be interpreted as being more preferred or more advantageous than other embodiments or designs. Specifically, the use of words such as "exemplary", "for example" or "for example" is intended to present related concepts in a specific way.

[0201] In the description of the embodiments of the present application, unless otherwise specified, “ / ” means or, for example, A / B can mean A or B; the “and / or” in the text is merely a description of the association relationship of associated objects, indicating that three relationships may exist, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, “multiple” refers to two or more than two.

[0202] In some processes described in the embodiments of the present application, multiple operations or steps that appear in a specific order are included, but it should be understood that these operations or steps may not be executed in the order in which they appear in the embodiments of the present application or in parallel, and the sequence number of the operation is only used to distinguish the different operations, and the sequence number itself does not represent any execution order. In addition, these processes may include more or fewer operations, and these operations or steps may be executed in sequence or in parallel, and these operations or steps may be combined.

[0203] Through the description of the above embodiments, those skilled in the art can clearly understand that the methods of the above embodiments can be implemented by means of software plus a necessary general hardware platform. Of course, they can also be implemented by hardware, but in many cases the former is a better implementation method. Based on such an understanding, the technical solution of the present application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product is stored in a storage medium as described above (such as ROM / RAM, magnetic disk, optical disk), and includes several instructions for causing a terminal device to execute the methods described in various embodiments of the present application.

[0204] The above are only the preferred embodiments of the present application, and do not limit the patent scope of the present application accordingly. Any equivalent structure or equivalent process transformation made by using the content of the specification and drawings of the present application, or directly or indirectly applied in other related technical fields, shall be equally included in the patent protection scope of the present application.

Claims

1. A method for planning a launch window and a launch point for a carrier rocket, characterized in that: The method comprises: In the J2000 coordinate system, according to the basic parameters of the target constellation, the orbital elements of each satellite of the target constellation at different times are calculated, and the positions of each satellite at different times are calculated according to the orbital elements; Convert the position of each satellite at different times into geodetic coordinates; Calculate the sub-satellite point trajectory of each satellite based on the geodetic coordinates of each satellite at different times; For any satellite in the target constellation, the time window of the satellite passing through the launch site is calculated according to the sub-satellite point trajectory of the satellite, the central position of the launch site and the maneuvering radius; The launch window and launch point of the launch vehicle are iteratively calculated according to the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite.

2. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 1, characterized in that: The step of calculating the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation includes: The basic parameters of the target constellation include the number of satellites, the number of orbital planes, relative phase parameters, orbital inclination and orbital altitude; According to the basic parameters of the target constellation, the preset initial value of the right ascension of the ascending node, the preset initial value of the argument of perigee and the preset initial value of the true anomaly, the orbital elements of each satellite of the target constellation at different times are calculated.

3. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 1, characterized in that: Calculating the position of each satellite at different times according to the orbital elements includes: The orbital elements include orbital semi-major axis, orbital eccentricity, orbital inclination, right ascension of ascending node, argument of perigee and true anomaly; Calculating a first matrix and a second matrix according to the ascending node right ascension, orbit inclination and perigee argument; The position of each satellite at different times is calculated according to the orbit semi-major axis, orbit eccentricity, true anomaly, the first matrix and the second matrix.

4. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 1, characterized in that: The converting the position of each satellite at different times into geodetic coordinates comprises: The position of each satellite at different times is converted from the J2000 coordinate system to the geocentric coordinate system to obtain the geocentric rectangular coordinates of each satellite at different times; Convert the geocentric rectangular coordinates of each satellite at different times into geodetic coordinates.

5. The method for planning a launch window and a launch point for a carrier rocket according to claim 1, characterized in that: The step of calculating, for any satellite in the target constellation, a time window for the satellite to pass through the launch site according to the sub-satellite point trajectory of the satellite, the central position of the launch site, and the maneuvering radius, comprises: Calculate the shortest distance between the center position of the launch site and the sub-satellite point track of the satellite; Determine whether the launch mission is feasible based on the shortest distance and the maneuvering radius of the launch site; If so, the time difference between the sub-satellite point trajectory of the satellite passing through the boundaries of the maneuvering launch range of the launch site is the time window for the satellite to pass through the launch site; the launch range is calculated based on the center position of the launch site and the maneuvering radius.

6. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 5, characterized in that: Based on the shortest distance and the maneuvering radius of the launch site, determine whether the launch mission is feasible, including: If the maneuvering radius of the launch site is greater than or equal to the shortest distance, the launch mission is feasible; If the maneuvering radius of the launch site is smaller than the shortest distance, the launch mission is not feasible.

7. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 1, characterized in that: According to the model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite, iteratively calculating the launch window and the launch point position of the carrier rocket, including: Calculating the launch window of the carrier rocket according to the time window and the orbital parameter requirements of this launch mission; Inputting the model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket; According to the launch trajectory data of the launch vehicle and the motion parameters of the satellite, the launch point position and orbit entry point parameters of the launch vehicle are adjusted to reduce the geographic latitude and longitude deviation of the sub-satellite point corresponding to the orbit entry point of the launch vehicle and the satellite; Determine whether the current launch window and launch point meet the launch requirements. If so, use the current launch window and launch point as the launch window and launch point of the carrier rocket; if not, return to the step of calculating the launch window of the carrier rocket based on the time window and the sub-satellite point trajectory of the satellite.

8. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 7, characterized in that: Calculating the launch window of the carrier rocket according to the time window and the orbital parameter requirements of this launch mission, including: Determine the flight duration and pre-set duration based on the orbital parameter requirements of this launch mission; The launch window of the carrier rocket is calculated according to the flight duration, the advance preset duration and the time window.

9. The method for planning a launch window and a launch point for a carrier rocket as claimed in claim 7, characterized in that: Inputting the model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite into a preset solid carrier rocket flight trajectory planning algorithm to obtain the launch trajectory data of the carrier rocket, including: The model of the carrier rocket, the time window and the sub-satellite point trajectory of the satellite are input into a preset solid carrier rocket flight trajectory planning algorithm, and planning calculations are performed in combination with the orbital parameter requirements, wind field conditions, performance of the carrier rocket, ground measurement and control constraints and landing area safety requirements of this launch mission to obtain the launch trajectory data of the carrier rocket.

10. A launch window and launch point planning device for a carrier rocket based on the method according to any one of claims 1 to 9, characterized in that: The device comprises: A position calculation module is used to calculate the orbital elements of each satellite of the target constellation at different times according to the basic parameters of the target constellation in the J2000 coordinate system, and calculate the position of each satellite at different times according to the orbital elements; A coordinate conversion module, used to convert the position of each satellite at different times into geodetic coordinates; The time window calculation module is used to calculate the sub-satellite point trajectory of each satellite according to the geodetic coordinates of each satellite at different times; and is also used to calculate the time window of the satellite passing through the launch site for any satellite in the target constellation according to the sub-satellite point trajectory of the satellite, the center position of the launch site and the maneuvering radius; The planning module is used to iteratively calculate the launch window and launch point of the launch vehicle according to the model of the launch vehicle, the time window and the sub-satellite point trajectory of the satellite.

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