A launch window dynamic reconstruction design method and system for lunar orbit

Abstract

The invention discloses a method and system for dynamically reconstructing a launch window for a lunar orbit, and relates to the technical field of lunar exploration launch orbit design. The method comprises the following steps: receiving the initial transfer orbit elements of the lunar orbit, and calculating the orbit elements at the moment of satellite-rocket separation according to the initial transfer orbit elements of the lunar orbit; inputting the orbit elements at the moment of satellite-rocket separation into an accurate dynamics equation, iteratively obtaining the accurate orbit elements of the first lunar orbit at the take-off moment, changing the take-off moment, and iteratively obtaining the accurate orbit elements of multiple lunar orbits; performing polynomial fitting on the bound elements of the accurate orbit elements of the multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and binding the elements of a launch vehicle as the polynomial coefficients corresponding to each orbit, so as to realize dynamic reconstruction of the lunar orbit according to the bound elements of the launch vehicle.

Description

Technical Field

[0001] The present invention relates to the technical field of lunar exploration launch trajectory design, and in particular to a launch window dynamic reconstruction design method and system for a lunar exploration trajectory. Background Art

[0002] The design technology of lunar launch orbit is a key technology for implementing lunar exploration missions. Traditional orbit design methods usually design nominal lunar orbits based on optimization analysis methods such as conic section splicing and differential precision correction. Due to the revolution of the moon and the rotation of the earth, the ideal lunar orbit corresponds to the launch time one by one. If the launch cannot be carried out on time due to weather or improper launch operation at the launch time, it can only wait for the next window to implement the launch, so it has the characteristics of low launch efficiency and high risk. Summary of the invention

[0003] In order to solve the deficiencies mentioned in the above background technology, the purpose of the present invention is to provide a launch window dynamic reconstruction design method and system for the lunar orbit, which can improve fuel utilization efficiency and increase the payload loading capacity of the lunar probe platform.

[0004] In a first aspect, the purpose of the present invention can be achieved by the following technical solution: a launch window dynamic reconstruction design method for a lunar orbit, the method comprising the following steps:

[0005] Receive the initial transfer orbital elements of the lunar orbit, and calculate the orbital elements at the time of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit;

[0006] Input the orbital elements at the separation point of the rocket and the satellite into the precise dynamics equation, iterate to get the precise orbital elements of the first lunar orbit at the take-off time, change the take-off time, iterate to get the precise orbital elements of multiple lunar orbits;

[0007] Polynomial fitting is performed on the precise orbital elements of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and the launch vehicle binding elements are used as the polynomial coefficients corresponding to each orbit, so as to realize dynamic reconstruction of the lunar window based on the launch vehicle binding elements.

[0008] In combination with the first aspect, in some implementations of the first aspect, the method further includes: a process of obtaining the number of initial transfer orbit elements of the lunar orbit:

[0009] Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows:

[0010]

[0011] hp represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E Indicates the true anomaly angle; r 0 represents the position vector; v 0 Represents the velocity vector.

[0012] In combination with the first aspect, in some implementations of the first aspect, the method further includes: the orbital element number at the satellite-rocket separation point is:

[0013] Flight time f And the number of first orbital elements at the separation point:

[0014]

[0015] Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation; the pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the launch plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in the near-circular parking orbit T h (T q ) seconds, and realize the dynamic orbital position requirements through secondary ignition.

[0016] In combination with the first aspect, in some implementations of the first aspect, the method further includes: forward integrating the flight orbit of the probe to the perigee according to the first orbit element number of the separation point:

[0017]

[0018] Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth; and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively. E (δ M ) is the switching function; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration;

[0019] Get the perigee T L The position velocity at the moment is recorded as:

[0020] (rt (r 0 (t),v 0 (t)),v t (r 0 (t),v 0 (t))).

[0021] In combination with the first aspect, in some implementations of the first aspect, the method further includes: the precise orbital elements of the first lunar orbit at the take-off time are obtained by differential correction based on the near-moon end orbit constraint and the near-Earth end departure constraint.

[0022] In combination with the first aspect, in certain implementations of the first aspect, the method further includes: the near-moon end orbit constraint:

[0023]

[0024] And the near-Earth orbit constraints:

[0025]

[0026] Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a M is the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r 0 =(x 0 ,y 0 ,z 0 ) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true near point angle constraint, and the constraint given above is

[0027] G=(g 1 ,g 2 ,g 3 ,g 4 ,g 5,g 6 )

[0028] Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints:

[0029]

[0030] The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows:

[0031]

[0032] Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E Indicates the longitude of the ascending node, f E is the true anomaly angle, T q is the launch vehicle take-off time. By adjusting T q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal.

[0033] In combination with the first aspect, in some implementations of the first aspect, the method further includes: the process of changing the take-off time and iteratively obtaining the precise orbital elements of multiple lunar orbits:

[0034] Adjust the departure time of the lunar orbit to Repeated calculations generate n precise Earth-Moon transfer orbits. In the n orbits, the launch directions of different orbits are consistent. The yaw program is added after the launch vehicle leaves the atmosphere. Achieve precise orbit insertion;

[0035] Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions.

[0036] In combination with the first aspect, in some implementations of the first aspect, the method further includes: performing polynomial fitting on the precise orbital elements of the plurality of lunar orbits, and using a least squares algorithm to fit the lunar orbital elements (h a ,i,ω,T) to perform polynomial fitting, obtain the corresponding polynomial coefficients, and bind the polynomial coefficient matrix C(ΔT q(n)), according to the accuracy requirements, the order of the first or second order polynomial is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window orbit elements to the moon.

[0037] In a second aspect, in order to achieve the above-mentioned purpose, the present invention discloses a launch window dynamic reconstruction design system for a lunar orbit, comprising:

[0038] The first processing module is used to receive the initial transfer orbital elements of the lunar orbit, and calculate the orbital elements at the time of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit;

[0039] The second processing module is used to input the orbital elements at the time of the satellite-rocket separation point into the precise dynamics equation, iteratively obtain the precise orbital elements of the first lunar orbit at the take-off time, change the take-off time, and iteratively obtain the precise orbital elements of multiple lunar orbits;

[0040] The dynamic reconstruction module is used to perform polynomial fitting on the precise orbital element binding parameters of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and use the polynomial coefficients corresponding to each orbit as the launch vehicle binding parameters, thereby realizing dynamic reconstruction of the lunar landing window based on the launch vehicle binding parameters.

[0041] In conjunction with the second aspect, in some implementations of the second aspect, the system further includes: a process of obtaining the number of initial transfer orbit elements of the lunar orbit in the first processing module:

[0042] Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows:

[0043]

[0044] h p represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E Indicates the true anomaly angle; r 0 represents the position vector; v 0 Represents the velocity vector.

[0045] Or the orbital elements at the moment of separation of the rocket and satellite in the first processing module:

[0046] Flight time f And the number of first orbital elements at the separation point:

[0047]

[0048] Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation; the pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the launch plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in the near-circular parking orbit T h (T q ) seconds, and realize the dynamic position requirement of orbital insertion through secondary ignition;

[0049] Among them, the first round of orbital elements of the separation point in the first processing module, the forward integrated flight orbit of the probe to the perigee:

[0050]

[0051] Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth; and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively. E (δ M ) is the switching function; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration;

[0052] Get the perigee T L The position velocity at the moment is recorded as:

[0053] (r t (r 0 (t),v 0 (t)),v t (r 0 (t),v 0 (t)))

[0054] The precise orbital elements of the first lunar orbit at the take-off time in the second processing module are obtained by differential correction from the near-moon orbit constraint and the near-Earth departure constraint.

[0055] Orbital constraints near the moon in the second processing module:

[0056]

[0057] And the near-Earth orbit constraints:

[0058]

[0059] Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a M is the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r 0 =(x 0 ,y 0 ,z 0 ) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true near point angle constraint, and the constraint given above is

[0060] G=(g 1 ,g 2 ,g 3 ,g 4 ,g 5 ,g 6 )

[0061] Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints:

[0062]

[0063] The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows:

[0064]

[0065] Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E Indicates the longitude of the ascending node, f E is the true anomaly angle, Tq is the launch vehicle take-off time. By adjusting T q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal;

[0066] In the second processing module, the take-off time is changed to iteratively obtain the precise orbital elements of multiple lunar orbits:

[0067] Adjust the departure time of the lunar orbit to Repeated calculations generate n precise Earth-Moon transfer orbits. In the n orbits, the launch directions of different orbits are consistent. The yaw program is added after the launch vehicle leaves the atmosphere. Achieve precise orbit insertion;

[0068] Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions.

[0069] In the dynamic reconstruction module, polynomial fitting is performed on the precise orbit elements of multiple lunar orbits. The least squares algorithm is used to fit the lunar orbit elements (h a ,i,ω,T) to perform polynomial fitting, obtain the corresponding polynomial coefficients, and bind the polynomial coefficient matrix C(ΔT q (n)), according to the accuracy requirements, the order of the first or second order polynomial is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window orbit elements to the moon.

[0070] Beneficial effects of the present invention:

[0071] The present invention is beneficial to improving the orbit insertion accuracy of the lunar orbit, reducing the fuel consumption required for mid-course corrections, and thus improving the lunar orbit delivery capability. The dynamic reconstruction of the launch window method greatly improves the orbit insertion accuracy, avoids additional mid-course corrections caused by launch time deviations, and significantly saves fuel consumption. Compared with the traditional multi-trajectory setting method, the dynamic adjustment of the launch window when the launch is delayed by 5 minutes can save about 35m / s of speed increment, which is approximately equivalent to saving 1.2% of the total mass of the probe in propellant weight, and the saved fuel can be used to carry more payloads. In addition, compared with the traditional method, the launch window is further expanded, and several discrete trajectories set in advance are expanded to an infinite number of continuous trajectories, which improves the redundancy and reliability of the mission. BRIEF DESCRIPTION OF THE DRAWINGS

[0072] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, for those skilled in the art, other drawings can be obtained based on these drawings without creative work.

[0073] Figure 1 It is a schematic flow chart of the method of the present invention;

[0074] Figure 2 It is a schematic diagram of the apogee height discrete sample points and the fitting curve of the present invention;

[0075] Figure 3 It is a schematic diagram of the discrete sample points of orbital inclination and the fitting curve of the present invention;

[0076] Figure 4 It is a schematic diagram of discrete sample points of perigee argument and fitting curve of the present invention;

[0077] Figure 5 It is a schematic diagram of the discrete sample points and fitting curve of the sliding time of the present invention;

[0078] Figure 6 It is a schematic diagram of the system structure of the present invention. DETAILED DESCRIPTION

[0079] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0080] Embodiment 1:

[0081] The following is an introduction to the relevant terms involved in the embodiments of the present application:

[0082] Least squares method: The least squares method is a mathematical tool that is widely used in many disciplines such as error estimation, uncertainty, system identification, prediction, and forecasting, and other data processing fields.

[0083] Lunar orbit: The lunar orbit is the trajectory of the center of mass of a spacecraft launched from the Earth and heading to the Moon.

[0084] like Figure 1 As shown, a launch window dynamic reconstruction design method for a lunar orbit includes the following steps:

[0085] S101: receiving the initial transfer orbital elements of the lunar orbit, and calculating the orbital elements at the time of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit;

[0086] The process of obtaining the initial transfer orbit elements of the lunar orbit:

[0087] Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows:

[0088]

[0089] h p represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E Indicates the true anomaly angle; r 0 represents the position vector; v 0 Represents the velocity vector.

[0090] Orbital elements at the time of separation of the rocket and satellite:

[0091] Flight time f And the number of first orbital elements at the separation point:

[0092]

[0093] Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation, which are used to control the thrust direction; the pitch angle represents the angle between the thrust direction in the shooting plane and the shooting direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the shooting plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in a near-circular parking orbit at an altitude of about 200 km. h (T q ) seconds, and realize the dynamic orbital position requirements through secondary ignition.

[0094] According to the first orbital element number at the separation point, the forward integrated probe's flight orbit to the perigee is:

[0095]

[0096] Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth (j=1 for the Sun, j=2 for the Moon); and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively.E (δ M ) is the switch function, which takes 1 when the detector is within the gravitational influence sphere of the Earth (Moon), and takes 0 otherwise; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration;

[0097] Get the perigee T L The position velocity at the moment is recorded as:

[0098] (r t (r 0 (t),v 0 (t)),v t (r 0 (t),v 0 (t))).

[0099] S102: Inputting the orbital elements at the time of separation of the satellite and the rocket into the precise dynamics equation, iteratively obtaining the precise orbital elements of the first lunar orbit at the time of takeoff, changing the takeoff time, and iteratively obtaining the precise orbital elements of multiple lunar orbits;

[0100] The precise orbital elements of the first lunar orbit at the take-off time are obtained by differential correction based on the orbit constraints near the moon and the departure constraints near the earth.

[0101] Near-moon orbit constraints:

[0102]

[0103] And the near-Earth orbit constraints:

[0104]

[0105] Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a M is the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r 0 =(x 0 ,y 0 ,z 0 ) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true near point angle constraint, and the constraint given above is

[0106] G=(g 1 ,g 2 ,g 3 ,g 4 ,g 5 ,g 6 )

[0107] Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints:

[0108]

[0109] The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows:

[0110]

[0111] Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E Indicates the longitude of the ascending node, f E is the true anomaly angle, T q The launch time of the carrier rocket is the launch time of the carrier rocket aiming at the desired orbit. q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal.

[0112] Changing the take-off time, iteratively obtaining the precise orbital elements of multiple lunar orbits:

[0113] Adjust the departure time of the lunar orbit to Repeated calculations were performed to generate n precise Earth-Moon transfer orbits. In order to make the n lunar orbits of the launch vehicle-rocket separation point more regular, after research and analysis, it was found that the launch directions of different orbits in the n orbits were consistent. By adding a yaw program after the launch vehicle leaves the atmosphere, the launch vehicle was able to achieve a smooth launch. Achieve precise orbit insertion;

[0114] Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions.

[0115] S103: Perform polynomial fitting on the precise orbital element binding parameters of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and use the polynomial coefficients corresponding to each orbit as the launch vehicle binding parameters, so as to realize dynamic reconstruction of the lunar orbit window according to the launch vehicle binding parameters.

[0116] The precise orbital elements of multiple lunar orbits are bound by polynomial fitting. The least squares algorithm is used to fit the lunar orbit binding elements (h a ,i,ω,T) to get the corresponding polynomial coefficients. Studies have shown that good fitting results can be obtained by linear fitting or quadratic polynomial fitting. If more complex Chebyshev polynomial fitting is used, the accuracy is higher. In this example, the perigee height of the earth is set to 200km when it departs.

[0117] Binding polynomial coefficient matrix C(ΔT q (n)), according to the accuracy requirements, the polynomial order such as linear or quadratic terms is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window lunar orbit elements.

[0118] Table 1 Statistics of linear fitting binding parameters errors

[0119]

[0120] Note: Increasing the apogee height of the lunar orbit by 5 km is equivalent to increasing the perigee speed by 0.0011 m / s, which is far below the orbit insertion accuracy indicator.

[0121] Table 2 Statistics of errors in quadratic fitting binding parameters

[0122]

[0123] Note: Increasing the apogee height of the lunar orbit by 5 km is equivalent to increasing the perigee speed by 0.0011 m / s, which is far below the orbit insertion accuracy indicator.

[0124] Embodiment 2: In the second aspect, as Figure 6 As shown, in order to achieve the above-mentioned purpose, the present invention discloses a launch window dynamic reconstruction design system for a lunar orbit, comprising:

[0125] The first processing module 11 is used to receive the initial transfer orbital elements of the lunar orbit, and calculate the orbital elements at the moment of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit;

[0126] The second processing module 12 is used to input the orbital elements at the time of the satellite-rocket separation point into the precise dynamics equation, iteratively obtain the precise orbital elements of the first lunar orbit at the take-off time, change the take-off time, and iteratively obtain the precise orbital elements of multiple lunar orbits;

[0127] The dynamic reconstruction module 13 is used to perform polynomial fitting on the precise orbital element binding parameters of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and use the launch vehicle binding parameters as the polynomial coefficients corresponding to each orbit, so as to realize the dynamic reconstruction window to the moon according to the launch vehicle binding parameters.

[0128] In combination with the second aspect, in some implementations of the second aspect, the system further includes: a process of obtaining the number of initial transfer orbit elements of the lunar orbit in the first processing module 11:

[0129] Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows:

[0130]

[0131] h p represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E Indicates the true anomaly angle; r 0 represents the position vector; v 0 Represents the velocity vector.

[0132] Or the orbital elements at the moment of satellite-rocket separation in the first processing module 11:

[0133] Flight time f And the number of first orbital elements at the separation point:

[0134]

[0135] Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation; the pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the launch plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in the near-circular parking orbit T h (T q ) seconds, and realize the dynamic position requirement of orbital insertion through secondary ignition;

[0136] Among them, the first round of orbital elements of the separation point in the first processing module 11, the forward integral of the flight orbit of the probe to the perigee:

[0137]

[0138] Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth; and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively. E (δ M ) is the switching function; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration;

[0139] Get the perigee T L The position velocity at the moment is recorded as:

[0140] (r t (r 0 (t),v 0 (t)),v t (r 0 (t),v 0 (t)))

[0141] The precise orbital elements of the first lunar orbit at the take-off time in the second processing module 12 are obtained by differential correction from the near-moon end orbit constraint and the near-earth end departure constraint;

[0142] The near-moon end orbital constraints in the second processing module 12:

[0143]

[0144] And the near-Earth orbit constraints:

[0145]

[0146] Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a Mis the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r 0 =(x 0 ,y 0 ,z 0 ) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true near point angle constraint, and the constraint given above is

[0147] G=(g 1 ,g 2 ,g 3 ,g 4 ,g 5 ,g 6 )

[0148] Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints:

[0149]

[0150] The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows:

[0151]

[0152] Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E Indicates the longitude of the ascending node, f E is the true anomaly angle, T q is the launch vehicle take-off time. By adjusting T q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal;

[0153] The second processing module 12 changes the take-off time and iterates to obtain the precise orbital elements of multiple lunar orbits:

[0154] Adjust the departure time of the lunar orbit to Repeated calculations generate n precise Earth-Moon transfer orbits. In the n orbits, the launch directions of different orbits are consistent. The yaw program is added after the launch vehicle leaves the atmosphere. Achieve precise orbit insertion;

[0155] Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions.

[0156] The dynamic reconstruction module 13 performs polynomial fitting on the precise orbital elements of multiple lunar orbits and uses the least squares algorithm to reconstruct the lunar orbital elements (h a ,i,ω,T) to perform polynomial fitting, obtain the corresponding polynomial coefficients, and bind the polynomial coefficient matrix C(ΔT q (n)), according to the accuracy requirements, the polynomial order such as linear or quadratic terms is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window lunar orbit elements.

[0157] Based on the same inventive concept, the present invention also provides a computer device, which includes: one or more processors, and a memory for storing one or more computer programs; the program includes program instructions, and the processor is used to execute the program instructions stored in the memory. The processor may be a central processing unit (CPU), or other general-purpose processors, digital signal processors (DSP), application-specific integrated circuits (ASIC), field-programmable gate arrays (FPGA) or other programmable logic devices, discrete gates or transistor logic devices, discrete hardware components, etc. It is the computing core and control core of the terminal, which is used to implement one or more instructions, specifically for loading and executing one or more instructions in a computer storage medium to implement the above method.

[0158] It needs to be further explained that, based on the same inventive concept, the present invention also provides a computer storage medium, on which a computer program is stored, and the computer program is executed by a processor to execute the above method. The storage medium can adopt any combination of one or more computer-readable media. The computer-readable medium can be a computer-readable signal medium or a computer-readable storage medium. The computer-readable storage medium can be, for example, but not limited to, an electrical, magnetic, optical, electrical, magnetic, infrared, or semiconductor system, device or device, or any combination of the above. More specific examples (non-exhaustive list) of computer-readable storage media include: an electrical connection with one or more wires, a portable computer disk, a hard disk, a random access memory (RAM), a read-only memory (ROM), an erasable programmable read-only memory (EPROM or flash memory), an optical fiber, a portable compact disk read-only memory (CD-ROM), an optical storage device, a magnetic storage device, or any suitable combination of the above. In the present invention, a computer-readable storage medium can be any tangible medium containing or storing a program, which can be used by an instruction execution system, device or device or used in combination with it.

[0159] In the description of this specification, the description with reference to the terms "one embodiment", "example", "specific example", etc. means that the specific features, structures, materials or characteristics described in conjunction with the embodiment or example are included in at least one embodiment or example of the present disclosure. In this specification, the schematic representation of the above terms does not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in any one or more embodiments or examples in a suitable manner.

[0160] The above shows and describes the basic principles, main features and advantages of the present disclosure. Those skilled in the art should understand that the present disclosure is not limited by the above embodiments, and the above embodiments and descriptions are only for explaining the principles of the present disclosure. Without departing from the spirit and scope of the present disclosure, the present disclosure may have various changes and improvements, and these changes and improvements fall within the scope of the present disclosure to be protected.

Claims

1. A method for dynamically reconfiguring a launch window for a lunar orbit, characterized in that: The method comprises the following steps: Receive the initial transfer orbital elements of the lunar orbit, and calculate the orbital elements at the time of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit; The orbital elements at the separation point of the rocket and the satellite are input into the precise dynamic equation, and the precise orbital elements of the first lunar orbit at the take-off time are obtained by iteration. The precise orbital elements of multiple lunar orbits are obtained by iteration by changing the take-off time. The precise dynamic equation is: Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth; and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively. E and δ M is the switching function; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration; Get the perigee T L The position velocity at the moment is recorded as: (r t ,v t ), r t is the position of the detector at the perigee, v t is the speed of the probe at perigee; Polynomial fitting is performed on the precise orbital elements of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and the launch vehicle binding elements are used as the polynomial coefficients corresponding to each orbit, so as to realize dynamic reconstruction of the lunar window based on the launch vehicle binding elements.

2. The method for dynamically reconfiguring a launch window for a lunar orbit according to claim 1, characterized in that: The process of obtaining the initial transfer orbit elements of the lunar orbit is as follows: Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows: Among them, h p represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E represents the true anomaly angle; r0 represents the position vector; v0 represents the velocity vector.

3. The method for dynamic reconstruction design of a launch window for a lunar orbit according to claim 1, characterized in that: The orbital elements at the moment of separation of the satellite and the rocket are: Flight time T f And the number of first orbital elements at the separation point: Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation; the pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the launch plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in the near-circular parking orbit T h seconds, and realize the dynamic orbital position requirements through secondary ignition.

4. The method for dynamically reconfiguring a launch window for a lunar orbit according to claim 1, characterized in that: The precise orbital elements of the first lunar orbit at the take-off time are obtained by differential correction based on the near-moon orbit constraint and the near-earth departure constraint.

5. The method for dynamic reconstruction design of a launch window for a lunar orbit according to claim 4, characterized in that: The near-moon orbital constraints are: And the near-Earth orbit constraints: Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a M is the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r0 = (x0, y0, z0) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true anomaly constraint, let the constraint given above be G = (g1, g2, g3, g4, g5, g6) Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints: The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows: Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E represents the longitude of the ascending node, f E is the true anomaly angle, T q is the launch vehicle take-off time. By adjusting T q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal.

6. The method for dynamically reconfiguring a launch window for a lunar orbit according to claim 1, characterized in that: The process of changing the take-off time and iteratively obtaining the precise orbital elements of multiple lunar orbits is as follows: Adjust the departure time of the lunar orbit to Repeated calculations generate n precise Earth-Moon transfer orbits. In the n orbits, the launch directions of different orbits are consistent. The yaw program is added after the launch vehicle leaves the atmosphere. Achieve precise orbit insertion; Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions.

7. The method for dynamically reconfiguring a launch window for a lunar orbit according to claim 1, characterized in that: The polynomial fitting of the precise orbit elements of multiple lunar orbits is performed by using the least squares algorithm to fit the lunar orbit elements (h a ,i,ω,T) to perform polynomial fitting, obtain the corresponding polynomial coefficients, and bind the polynomial coefficient matrix C(ΔT q (n)), according to the accuracy requirements, the order of the first or second order polynomial is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window orbit elements to the moon.

8. A launch window dynamic reconstruction design system for lunar orbit, characterized in that: include: The first processing module is used to receive the initial transfer orbital elements of the lunar orbit, and calculate the orbital elements at the time of the satellite-rocket separation point according to the initial transfer orbital elements of the lunar orbit; The second processing module is used to input the orbital elements at the time of the satellite-rocket separation point into the precise dynamic equation, iteratively obtain the precise orbital elements of the first lunar orbit at the take-off time, change the take-off time, and iteratively obtain the precise orbital elements of multiple lunar orbits; wherein the precise dynamic equation is Among them, r is the position vector of the detector relative to the center of the earth, r j is the perturbing body M j Position vector relative to the center of the Earth; and The two terms refer to the non-spherical perturbations of the Earth and the Moon, respectively. E and δ M is the switching function; F drag represents the atmospheric drag acceleration, F SRP represents the solar pressure perturbation acceleration; Get the perigee T L The position velocity at the moment is recorded as: (r t ,v t ), r t is the position of the detector at the perigee, v t is the speed of the probe at perigee; The dynamic reconstruction module is used to perform polynomial fitting on the precise orbital element binding parameters of multiple lunar orbits to obtain the polynomial coefficients corresponding to each orbit, and use the polynomial coefficients corresponding to each orbit as the launch vehicle binding parameters, thereby realizing dynamic reconstruction of the lunar landing window based on the launch vehicle binding parameters.

9. The launch window dynamic reconstruction design system for the lunar orbit according to claim 8, characterized in that: The process of obtaining the initial transfer orbit elements of the lunar orbit in the first processing module: Set the initial epoch, calculate the initial value of the lunar orbit according to the lunar position when reaching the perigee, and combine the lunar orbit constraints to solve the initial transfer orbit elements of the lunar orbit as follows: h p represents the perigee angular distance; a E represents the semi-major axis of the orbit; i E represents the orbital inclination; ω E Indicates the pericentric angular distance; Ω E represents the right ascension of the ascending node; f E represents the true anomaly; r0 represents the position vector; v0 represents the velocity vector; Or the orbital elements at the moment of separation of the rocket and satellite in the first processing module: Flight time T f And the number of first orbital elements at the separation point: Where r is the radius vector from the center of the Earth to the center of mass of the launch vehicle, P, R and g are the thrust, aerodynamic force and gravitational acceleration vectors, and ψ(t) represent the pitch angle and yaw angle of the flight program calculated by the launch vehicle navigation; the pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the moment of launch, and the yaw angle represents the angle between the thrust direction and the launch plane at the moment of launch; in order to meet the launch requirements of the lunar orbit at different lunar positions, the launch vehicle's orbital stage glides in the near-circular parking orbit T h seconds, and achieve the dynamic requirements of orbital position through secondary ignition; Among them, the first orbital elements of the separation point in the first processing module are integrated forward from the flight orbit of the probe to the perigee; The precise orbital elements of the first lunar orbit at the take-off time in the second processing module are obtained by differential correction from the near-moon orbit constraint and the near-Earth departure constraint. Orbital constraints near the moon in the second processing module: And the near-Earth orbit constraints: Among them, r t -r M =(x tM ,y tM ,z tM ) T , is the position and velocity vector of the probe's perigee, r t 、r M are the positions of the probe and the moon at perigee, v t 、v M are the speeds of the probe and the moon at perigee, a M is the lunar equatorial radius, h M is the perigee height constraint, i M is the orbit inclination constraint near the moon; r0 = (x0, y0, z0) T , are the position and velocity vector of the detector at perigee, a E is the Earth's equatorial radius constraint, h p is the perigee height constraint, i E is the near-Earth orbit inclination constraint, f E is the true anomaly constraint, let the constraint given above be G = (g1, g2, g3, g4, g5, g6) Perform differential iterative correction according to the following formula to obtain the precise Earth-Moon transfer orbit that satisfies the constraints: The expected position and velocity of the lunar transfer orbit after iteration is (r E ,v E ), which can be converted to the expected number of orbital elements as follows: Where h p represents the perigee height, a E represents the semi-major axis, i E represents the orbital inclination, ω E represents the argument of perigee, Ω E Indicates the longitude of the ascending node, f E is the true anomaly angle, T q is the launch vehicle take-off time. By adjusting T q , so that the launch vehicle can complete the launch orbit design with the maximum carrying capacity as the optimization goal; In the second processing module, the take-off time is changed to iteratively obtain the precise orbital elements of multiple lunar orbits: Adjust the departure time of the lunar orbit to Repeated calculations generate n precise Earth-Moon transfer orbits. In the n orbits, the launch directions of different orbits are consistent. The yaw program is added after the launch vehicle leaves the atmosphere. Achieve precise orbit insertion; Through repeated iterative calculations, we can obtain the precise lunar orbit elements with strong regularity under n different take-off time conditions. In the dynamic reconstruction module, polynomial fitting is performed on the precise orbit elements of multiple lunar orbits. The least squares algorithm is used to fit the lunar orbit elements (h a ,i,ω,T) to perform polynomial fitting, obtain the corresponding polynomial coefficients, and bind the polynomial coefficient matrix C(ΔT q (n)), according to the accuracy requirements, the order of the first or second order polynomial is selected, the launch vehicle realizes the dynamic reconstruction window to the moon according to the binding launch parameters, and the probe completes the subsequent flight correction according to the real-time launch window orbit elements to the moon.

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