Design method and system for takeoff elements at any time of the Mars exploration launch window

Abstract

The present invention discloses a design method for the elements of achieving takeoff at any time during the Mars exploration launch window, which relates to the technical field of Mars exploration launch orbit design. The present invention includes: calculating the initial values required for the departure parameters of the detector according to the launch time; designing the orbital parameters of the launch vehicle according to the initial values of the detector departure parameters, and combining the calculation amount and design amount in the orbital parameters of the launch vehicle to complete the launch orbit design through iteration; adjusting the target orbit inclination and calculating and generating the target elements of the launch orbit with strong regularity of change with the takeoff time. The present invention can enable the launch vehicle to achieve takeoff at any time within the launch window on the same day, meet the precise departure parameter requirements for the detector to escape and fly on the orbit to Mars, save the additional propellant consumption of the detector caused by the deviation of the takeoff time after the detector separation, improve the transportation capacity of the detector platform, and at the same time improve the design work efficiency.

Description

Technical Field

[0001] The present invention relates to the technical field of Mars exploration launch orbit design, and specifically to a method and system for designing the elements for takeoff at any time in the Mars exploration launch window. Background Art

[0002] The deep space launch orbit design technology is a top-level key technology for deep space exploration activities. During the ground launch process of a Mars probe, usually the upper stage of the launch vehicle first sends the probe into a nearly circular parking orbit at an altitude of about 200 km, then glides to near the re-acceleration point, and the upper stage starts again to send the probe into the predetermined escape orbit. Affected by factors such as the Earth's rotation, the ideal launch window for Mars exploration is a zero window.

[0003] However, there are many influencing factors in the test process of the launch vehicle launch process. Especially for cryogenic rockets, the pre-launch procedures are complex, and the launch process is also affected by meteorological factors and other accidental factors. To ensure the launch reliability, the common practice is to design a launch window with a time width of M (M = 30 - 50 min). Within this window, the launch is carried out aiming at the predetermined orbit injection parameters. The deviation from the predetermined launch time and the deviation of the orbit injection parameters need to be corrected by the probe consuming propellant. Therefore, we propose a method and system for designing the elements for takeoff at any time in the Mars exploration launch window. Summary of the Invention

[0004] The purpose of the present invention is to provide a method and system for designing the elements for takeoff at any time in the Mars exploration launch window, which can enable the launch vehicle to take off at any time within the launch window on the same day, meet the precise departure parameter requirements for the probe to escape and fly in the Mars-bound orbit, and save the additional propellant consumption of the probe caused by the deviation of the takeoff time after separation.

[0005] To achieve the above purpose, the present invention provides the following technical solution: A method for designing the elements for takeoff at any time in the Mars exploration launch window, including the following steps:

[0006] Calculate the initial values required for the probe departure parameters according to the launch time;

[0007] Design the orbit injection parameters of the launch vehicle according to the initial values of the probe departure parameters, and combine the calculated and designed quantities in the orbit injection parameters of the launch vehicle to complete the launch orbit design through iteration;

[0008] Adjust the target orbit inclination, and calculate and generate the target elements of the launch orbit with strong regularity of change with the takeoff time;

[0009] For the generated multiple target elements of the launch orbit, use the least square principle to fit the target elements into a polynomial of the takeoff time relative to the reference takeoff time to obtain the fitting coefficients;

[0010] Bind the obtained polynomial fitting coefficients in the launch vehicle;

[0011] According to the real-time launch window and the polynomial coefficients corresponding to each root number, calculate the real-time zero-window launch elements by the polynomial function to achieve real-time zero-window launch.

[0012] Furthermore, the initial values required to calculate the departure parameters of the detector according to the launch time are as follows:

[0013] From the start of the detector separation to the end of reaching the periapsis, the entire Mars exploration orbit is integrated in the heliocentric coordinate system, and the flight dynamics equation is as follows:

[0014]

[0015] In the formula, μ S is the solar gravitational coefficient, μ j is the gravitational coefficient of the disturbing celestial body in the solar system. For the Mars exploration transfer orbit, the influences of the eight major planets and the moon need to be considered; and the two items respectively refer to the non-spherical perturbation force terms of the Earth and Mars, δ E (δ M ) is a switching function, which takes 1 when the detector is within the sphere of influence of the Earth (Mars) gravity and 0 otherwise. δV(t1, t2) represents the application and end of deep space maneuvers at times t1 and t2, and F SRP represents the solar radiation pressure term, and F others represents other minor influence terms;

[0016] Combined with the requirements of Earth departure and Mars arrival, solve the optimization problem shown in Equation (2) through the particle swarm optimization algorithm to calculate the theoretical departure time of the detector:

[0017]

[0018] In the formula, δV pM represents the velocity increment required to enter the target Mars orbit, G1 represents the equality constraints that need to be satisfied during the transfer process, and G2 represents the inequality constraints that need to be satisfied during the transfer process;

[0019] Combined with Equation (1) and Equation (2), calculate the position and velocity of the detector at the departure and arrival times where i = E and i = M respectively represent the position and velocity at Earth departure and Mars arrival, which can be mutually converted with the orbital elements of the detector as follows:

[0020]

[0021] In Equation (3), H pE represents the perigee altitude, C3 EDenote the Earth escape C3, i E Denote the orbital inclination, ω E Denote the argument of perigee, Ω E Denote the longitude of the ascending node, f E Is the true anomaly. In addition, denote the takeoff moment of the launch vehicle as Tq, where H pE Is usually about the parking orbit altitude, Ω E Is directly related to the flight path of the launch vehicle, f E Is determined by the flight orbit characteristics of the launch vehicle, the orbital inclination i E ∈[i0 i1].

[0022] Furthermore, the calculation quantities of the launch vehicle include the longitude of the ascending node and the true anomaly, and the design quantities include the perigee altitude, the escape C3, the orbital inclination, and the argument of perigee.

[0023] Furthermore, in the launch inertial system, the flight dynamics equation of the rocket in the atmosphere is as follows:

[0024]

[0025] Where r represents the vector radius from the geocenter to the center of mass of the launch vehicle in the launch inertial system, and P, R, and g are the thrust, aerodynamic force, and gravitational acceleration vectors in the launch inertial system, And ψ(t) represents the flight program pitch angle and yaw angle. The pitch angle represents the angle between the thrust direction in the injection plane and the injection direction at the instant of launch, and the yaw angle represents the angle by which the thrust direction deviates from the injection plane at the instant of launch. During the flight, the thrust direction is controlled by the pitch and yaw program angles. Under the constraint of the orbit injection target, the launch orbit design is completed with the maximum launch capacity as the optimization target. The specific method is:

[0026] Adjust Make

[0027] In the formula, Tg(t) represents the shutdown moment of the active section, Th(Tq) represents the coasting time, the subscript 1 represents the orbit injection parameters achieved by the launch vehicle, the subscript 2 represents the orbit injection parameters required by the detector, and Ω E And f E Are adopted in the iteration, and ε represents the desired iteration accuracy.

[0028] Furthermore, the method for generating the launch orbit is specifically as follows:

[0029] Use α o -α L To represent the difference between the escape right ascension and the Greenwich sidereal time angle of the launch site at the takeoff moment. Then α o -α L The relationships with the escape declination, the injection direction, and the relative takeoff moment ΔTq are as follows:

[0030]

[0031] ΔTq = (α o - α L ) / ω e (6)

[0032] In the formula, δ0 represents the escape declination, ω e represents the angular velocity of the Earth's rotation, and φ L represents the latitude of the launch site;

[0033] At the target escape declination, the launch time and the firing direction have a strong correlation, and

[0034] i E = i E (A L (Tq)) (7)

[0035] ω E = ω E (Tq, Th(Tq)) (8)

[0036] That is, the target elements can all be expressed as continuous functions of the takeoff time. At the same time, due to the Earth's launch properties, applying the yaw angle ψ(Tq) in engineering implementation can achieve the equivalence with adjusting the firing direction A L (Tq). Specifically, by applying a constant yaw ψ i (t) in the second-stage acceleration working section of the upper stage and adjusting the takeoff time, coasting time, and shutdown time, the launch orbit of the launch vehicle can meet the orbital injection requirements.

[0037] Furthermore, fitting the orbital inclination, argument of perigee, and coasting time at different takeoff times within the window to polynomials of the difference between the takeoff time and the reference takeoff time, obtaining the fitting coefficients, the calculation method is as follows:

[0038] c = (A T A) -1 A T b (9)

[0039] In the formula, c represents the fitting coefficient, b represents the sample point result sequence, and A is a matrix composed of launch times. For n sample points (n ≥ 4), taking the fitting of the orbital inclination by a cubic curve as an example, the expression is as follows:

[0040] i EΔt = c3Δt 3 + c2Δt 2 + c1Δt + c0 (10)

[0041] In the formula, Δt is the time difference between the launch time and the reference launch time within the launch window, and i EΔtIt is the orbital inclination corresponding to the launch at a time difference of Δt from the reference launch time, c k , where k = 0, 1, 2, 3 are fitting coefficients. When there are no less than 4 groups of sample points, the fitting coefficients can be solved through Equation (9). Similarly, the fitting coefficients of the argument of perigee, launch C3, and coasting time can be calculated.

[0042] Furthermore, the real-time zero-window launch elements are calculated by a polynomial function to achieve real-time zero-window launch:

[0043] Within the launch window, for any given launch time, according to the bound fitting coefficients and the deviation of the launch time relative to the reference time, the target orbit parameters are quickly calculated according to Equation (10), guiding the launch vehicle to fly according to the orbit elements fitted by aiming, and achieving precise orbit insertion.

[0044] According to the first aspect of the present invention, the present invention provides an element design system for realizing takeoff at any time in the Mars exploration launch window, including:

[0045] The first calculation module: used to calculate the initial values required for the detector departure parameters according to the launch time;

[0046] The launch orbit design module: used to design the launch vehicle orbit insertion parameters according to the initial values of the detector departure parameters, and complete the launch orbit design through iteration in combination with the calculated and designed quantities in the launch vehicle orbit insertion parameters;

[0047] The second calculation module: used to calculate and generate a launch orbit with strong regularity of orbit insertion parameters by adopting the method of adjusting the orbital inclination;

[0048] The third calculation module: used to perform polynomial fitting on the generated multiple launch orbits according to the least squares principle to obtain the polynomial fitting coefficients of the designed quantity parameters varying with the launch time under a fixed perigee altitude;

[0049] The binding module: used to bind the obtained polynomial fitting coefficients to the launch vehicle;

[0050] The fourth calculation module: used to calculate the real-time zero-window launch elements by a polynomial function according to the real-time launch window and the polynomial coefficients corresponding to each element, and achieve real-time zero-window launch.

[0051] According to the second aspect of the present invention, the present invention provides a terminal device, including a memory, a processor, and a computer program stored in the memory and capable of running on the processor. The memory stores a computer program capable of running on the processor. When the processor loads and executes the computer program, the element design method for realizing takeoff at any time in the Mars exploration launch window is adopted.

[0052] According to the third aspect of the present invention, the present invention provides a storage medium containing computer-executable instructions, and when the computer-executable instructions are executed by a computer processor, they are used to execute the method for designing the elements for taking off at any moment within the Mars exploration launch window.

[0053] The present invention has at least the following beneficial effects:

[0054] The method for determining the launch elements according to the present invention enables the launch vehicle to take off at any moment within the same day's launch window, meets the precise departure parameter requirements for the detector to escape into the Mars orbit, saves the additional propellant consumption of the detector caused by the deviation of the take-off moment after separation, improves the transportation capacity of the detector platform, and at the same time improves the design work efficiency.

[0055] Of course, it is not necessary for any product implementing the present invention to achieve all the above-mentioned advantages simultaneously. BRIEF DESCRIPTION OF THE DRAWINGS

[0056] Figure 1 Schematic diagram of the flight process of a launch vehicle for Mars exploration detector (looking down from the North Pole);

[0057] Figure 2 Schematic diagram of the overall flow of the method according to the present invention;

[0058] Figure 3 Schematic diagram of the relationship between the firing direction, the required take-off moment and the declination of escape;

[0059] Figure 4 Fitting curve graph of the argument of perigee of a certain launch vehicle for the Earth-Mars transfer orbit in the embodiment of the present invention;

[0060] Figure 5 Fitting curve graph of the orbital inclination of a certain launch vehicle for the Earth-Mars transfer orbit in the embodiment of the present invention;

[0061] Figure 6 Fitting curve graph of C3 of a certain launch vehicle for the Earth-Mars transfer orbit in the embodiment of the present invention;

[0062] Figure 7 Fitting curve graph of the coasting time of a certain launch vehicle for the Earth-Mars transfer orbit in the embodiment of the present invention. DETAILED DESCRIPTION OF THE EMBODIMENTS

[0063] Next, the technical solutions in the embodiments of the present disclosure will be clearly and completely described in conjunction with the accompanying drawings in the embodiments of the present disclosure. Obviously, the described embodiments are only a part of the embodiments of the present disclosure, rather than all of the embodiments. Based on the embodiments in the present disclosure, all other embodiments obtained by those of ordinary skill in the art without making creative efforts belong to the protection scope of the present disclosure.

[0064] Please refer to Figure 1-7 , the present invention provides a technical solution: a method for designing the elements for taking off at any time in the Mars exploration launch window, including the following steps:

[0065] Calculate the initial values required for the detector departure parameters according to the launch time;

[0066] According to the initial values of the detector departure parameters, design the orbital injection parameters of the launch vehicle, and combine the calculation amount and design amount in the orbital injection parameters of the launch vehicle to complete the launch orbit design through iteration;

[0067] By adopting the method of adjusting the orbital inclination, calculate and generate a launch orbit with strong regularity of the orbital injection parameters;

[0068] For the generated multiple launch orbits, perform polynomial fitting according to the least square principle to obtain the polynomial fitting coefficients of the design amount parameters varying with the launch time under a fixed perigee altitude;

[0069] Bind the obtained polynomial fitting coefficients into the launch vehicle;

[0070] According to the real-time launch window and the polynomial coefficients corresponding to each element, calculate the real-time zero-window launch elements by the polynomial function to achieve real-time zero-window launch.

[0071] The specific embodiments are as follows:

[0072] S1. Design the initial values required for the detector departure parameters according to the departure date;

[0073] From the separation of the detector to the arrival at the periapsis of Mars, the entire Mars exploration orbit can be integrated in the heliocentric coordinate system, and the flight dynamics equation is as follows:

[0074]

[0075] In the above formula, μ S is the solar gravitational coefficient, μ j is the gravitational coefficient of the disturbing celestial body in the solar system. For the Mars exploration transfer orbit, the influences of the eight major planets and the moon need to be considered; and the two terms respectively refer to the non-spherical perturbation force terms of the Earth and Mars, δ E (δ M ) is a switching function, which takes 1 when the detector is within the sphere of influence of the Earth (Mars) gravity and 0 otherwise. δV(t1,t2) represents the application and end of the deep space maneuver at times t1 and t2, and F SRP represents the solar radiation pressure term, and F others represents all other minor influence terms, such as the upper atmospheres of the Earth and Mars, tidal effects, radiation pressure of the Earth and Mars, etc.;

[0076] Combined with the requirements of departure from the Earth and arrival at Mars, by solving the optimization problem shown in Equation (2), the theoretical departure time of the detector can be obtained:

[0077]

[0078] In the formula, δV pM represents the velocity increment required to enter the target Mars orbit, G1 represents the equality constraints that need to be satisfied during the transfer process, and G2 represents the inequality constraints that need to be satisfied during the transfer process; the above formula can be solved by various methods such as the B-plane method, the iterative differential correction algorithm, the direct optimization algorithm, or other optimization algorithms, and the specific process will not be elaborated. Due to the Earth's rotation and the location limitations of the launch site, it is usually necessary to launch when the orbit passes over the launch site. Combining Equation (1) and Equation (2), the position and velocity of the detector at the departure and arrival times can be calculated Among them, when i = E and i = M, they represent the position and velocity at the departure from the Earth and the arrival at Mars respectively, and they can be mutually converted with the orbital elements of the detector entering the orbit as follows:

[0079]

[0080] In Equation (3), H pE represents the perigee altitude, C3 E represents the Earth escape C3, 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. In addition, the takeoff time of the launch vehicle is denoted as Tq; among them, H pE is usually approximately the altitude of the parking orbit, Ω E is directly related to the flight path of the launch vehicle, f E is determined by the flight orbit characteristics of the launch vehicle and changes little; the orbital inclination i E ∈[i0 i1]. To make full use of the Earth's rotation to improve the launch capacity and take into account the landing area constraints and the orbit injection requirements, the launch direction should preferably be the eastward launch direction;

[0081] S2. According to the initial values of the detector departure parameters, design the launch vehicle orbit injection parameters (including the takeoff time). Combining the calculation amounts and design amounts in the launch vehicle orbit injection parameters, complete the launch orbit design through iterative optimization;

[0082] In the launch inertial system, the flight dynamics equation of the rocket in the atmosphere is as follows:

[0083]

[0084] where r represents the vector from the geocenter to the center of mass of the launch vehicle in the inertial launch system, and P, R, and g are the thrust, aerodynamic force, and gravitational acceleration vectors in the inertial launch system. And ψ(t) represents the pitch angle and yaw angle of the flight program. The pitch angle represents the angle between the thrust direction in the firing plane and the firing direction at the instant of launch, and the yaw angle represents the angle by which the thrust direction deviates from the firing plane at the instant of launch. During the flight, the thrust direction is controlled by the pitch and yaw program angles. Under the constraint of the orbit insertion target, the launch orbit design is generally completed with the maximum payload capacity as the optimization goal.

[0085] The specific method is as follows:

[0086] Adjust to make

[0087] In the formula, Tg(t) represents the shutdown time of the powered flight phase, Th(Tq) represents the coasting time, the subscript 1 represents the orbit insertion parameters achieved by the launch vehicle, the subscript 2 represents the orbit insertion parameters required by the detector, and Ω E and f E are adopted in the iteration, and ε represents the desired iteration accuracy.

[0088] S3. Adjust the target orbit inclination and calculate the elements of the launch orbit target that change regularly with the takeoff time;

[0089] Theoretical research shows that if α o -α L represents the difference between the escape right ascension and the Greenwich sidereal time angle of the launch site at the takeoff time, then α o -α L has the following relationship with the escape declination, the firing direction, and the relative takeoff time ΔTq:

[0090]

[0091] ΔTq = (α o -α L ) / ω e (6)

[0092] In the formula, δ0 represents the escape declination, ω e represents the angular velocity of the earth's rotation, and φ L represents the latitude of the launch site. From this, the relationship between the firing direction, the required takeoff time, and the escape declination can be obtained as shown in Figure 2 ;

[0093] It can be seen that under the target escape declination, the launch time and the firing direction have a good correlation, and

[0094] i E = i E (A L (Tq)) (7)

[0095] ω E = ω E (Tq, Th(Tq)) (8)

[0096] That is, all the target elements can be expressed as continuous functions of the takeoff time. At the same time, due to the earth launch property, applying the yaw angle ψ(Tq) in engineering implementation can achieve and adjust the firing direction A L (Tq) is equivalent. Specifically, by applying a constant yaw ψ i (t) in the second-stage acceleration working section of the upper stage and adjusting the takeoff time, glide time, and shutdown time, the launch orbit of the launch vehicle can meet the orbit injection requirements;

[0097] The present invention does not strongly constrain the parking orbit. In the calculation example, the parking orbit has an apogee altitude of 245 - 290 km and a perigee altitude of 160 - 175 km. The calculation results are as follows:

[0098] Table 1 Calculation results of target elements at different takeoff times

[0099]

[0100]

[0101] ΔTq represents the deviation of the takeoff time. Under different takeoff time deviations, the target elements have good regularity;

[0102] S4. For the multiple transfer orbits generated in S3, polynomial fitting is performed according to the least squares principle to obtain the polynomial fitting coefficients of parameters such as the orbit inclination, escape C3, and argument of perigee varying with the launch time at a fixed perigee altitude;

[0103] The main process is as follows: According to the results of S3, the orbit inclination, argument of perigee, and glide time within the window at different takeoff times are fitted to polynomials of the difference between the takeoff time and the reference takeoff time to obtain the fitting coefficients. The calculation method is as follows:

[0104] c = (A T A) -1 A T b (9)

[0105] In the formula, c represents the fitting coefficient, b represents the result sequence of the sample points, and A is a matrix composed of the launch times. For n sample points (n ≥ 4), taking the fitting of the orbit inclination by a cubic curve as an example, the expression is as follows:

[0106] i EΔt = c3Δt 3 + c2Δt 2 + c1Δt + c0 (10)

[0107] Where Δt is the time difference between the launch time in the launch window and the reference launch time, i EΔt is the orbital inclination corresponding to the launch at a time Δt different from the reference launch time, c k , k = 0, 1, 2, 3 are fitting coefficients. When no less than 4 sets of sample points are known, the fitting coefficients can be solved by equation (9). Similarly, the fitting coefficients of target parameters such as perigee argument and glide time can be calculated. Taking a rocket launch Earth-Mars transfer orbit as an example, the result Figure 4-7 As shown;

[0108] The fitting accuracy is verified by using sample points. The fitting errors of the elements at each sample point in S3 are shown in the following table:

[0109] Table 2 Fitting errors of various parameters

[0110]

[0111]

[0112] In the above table, Vp200 represents the escape velocity corresponding to the perigee of 200km. It has a one-to-one correspondence with C3. For the sake of intuitiveness, this parameter is used instead of C3 in error analysis. When implementing it, it is selected in combination with the actual needs of the project. It can be seen from the above table that the fitting errors are all high-order small quantities relative to the parameter orbital accuracy deviation; therefore, the present invention has very high fitting accuracy, and both the quadratic curve fitting and the cubic curve fitting meet the accuracy requirements of engineering applications, and the cubic polynomial fitting accuracy is higher. If fewer sample points are used for fitting, and the fitting accuracy is verified by interpolation of other points, the fitting accuracy is equivalent to the results in the above table, and it can still be guaranteed that the fitting deviation is a high-order small quantity required for the orbital accuracy deviation;

[0113] It is particularly noted that the present invention does not strictly limit the apogee and perigee heights of the parking orbit. If the parking orbit is strictly limited to a 200km circular orbit, the fitting accuracy of the relevant parameters is higher;

[0114] S5. The fitting polynomial coefficients in the launch vehicle bound into S4;

[0115] Bind the S4 fitting coefficients to the launch vehicle;

[0116] S6. Generate real-time zero window transmission elements according to the real-time transmission window and the polynomial coefficients corresponding to each number of roots, and complete the transmission:

[0117] Within the launch window, for any given launch time, the target orbit parameters are quickly calculated according to the fitting coefficients bound in S5, and the launch vehicle is guided to fly according to the orbital elements obtained by the aiming fitting to achieve precise orbit entry.

[0118] It should be noted that, in this document, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprising", "including" or any other variant thereof are intended to cover non-exclusive inclusion, such that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device.

[0119] For those of ordinary skill in the art, the specific meanings of the above terms in the present invention can be understood according to specific circumstances. When an element is referred to as being "assembled on", "installed on", "fixed to" or "disposed on" another element, it can be directly on the other element or there may also be an intermediate element. When an element is considered to be "connected" to another element, it can be directly connected to the other element or there may be an intermediate element at the same time. The terms "vertical", "horizontal", "upper", "lower", "left", "right" and similar expressions used herein are for illustrative purposes only and do not represent the only embodiments.

[0120] Although the embodiments of the present invention have been shown and described, for those of ordinary skill in the art, it can be understood that various changes, modifications, substitutions and variations can be made to these embodiments without departing from the principles and spirit of the present invention, and the scope of the present invention is defined by the appended claims and their equivalents.

[0121] In the description of this specification, the description with reference to terms such as "one embodiment", "example", "specific example", etc. means that the specific features, structures, materials or characteristics described in connection with the embodiment or example are included in at least one embodiment or example of the present disclosure. In this specification, the schematic representations of the above terms do not necessarily refer to the same embodiment or example. Moreover, the specific features, structures, materials or characteristics described can be combined in a suitable manner in any one or more embodiments or examples.

Claims

1. A method for calculating the elements of a Mars exploration launch window to achieve takeoff at any time, characterized in that, The following steps are involved: Calculate the initial values ​​of the detector's departure parameters based on the launch time; According to the initial values ​​of the probe's departure parameters, the launch vehicle's orbital parameters are designed. The launch trajectory design is completed through iteration by combining the calculated and designed quantities in the launch vehicle's orbital parameters. The corresponding take-off time is used as the reference take-off time. Adjust the target orbit inclination and calculate and generate the launch orbit target parameters with strong regularity in changing with the take-off time; For the target parameters of the generated multiple launch trajectories, the least squares principle is used to fit the target parameters into a polynomial of the take-off time relative to the reference take-off time, and the fitting coefficients are obtained; Bind the obtained polynomial fitting coefficients into the launch vehicle; According to the real-time transmission window and the polynomial coefficients corresponding to each root number, the real-time zero window transmission elements are calculated by the polynomial function to realize the real-time zero window transmission; Among them, the calculated quantities of the launch vehicle include the longitude of the ascending node and the true perigee angle, and the design quantities include the perigee height, escape C3, orbit inclination, and perigee argument; The orbit inclination, perigee argument and glide time at different take-off times in the window are fitted to a polynomial of the difference between the take-off time and the reference take-off time, and the fitting coefficient is obtained. The calculation method is as follows: c = (A T A) -1 A T b(9) Where c represents the fitting coefficient, b represents the sample point result sequence, and A is a matrix composed of the launch time. For n sample points, n ≥ 4, taking the fitting of the orbital inclination by the cubic curve as an example, the expression is as follows: i EΔt = c3Δt 3 + c2Δt 2 + c1Δt + c0 (10) where Δt is the time difference between the launch time within the launch window and the reference launch time, and i EΔt is the orbital inclination corresponding to the launch at a time difference of Δt from the reference launch time, c k , k = 0, 1, 2, 3 are fitting coefficients. When there are no less than 4 groups of sample points, the fitting coefficients can be solved through Equation (9). Similarly, the fitting coefficients of the argument of perigee, escape C3, and coasting time can be calculated.

2. The element design method for achieving takeoff at any time of the Mars exploration launch window according to claim 1, characterized in that: The calculation of the initial values ​​of the detector departure parameters required according to the launch time is as follows: From the separation of the probe to the arrival at the perigee, the entire Mars exploration orbit is integrated in the heliocentric coordinate system, and the flight dynamics equation is as follows: where μ S is the solar gravitational coefficient, and μ j is the gravitational coefficient of the disturbing body under the solar system. For the Mars exploration transfer orbit, the influences of the eight major planets and the moon need to be considered; and the two items respectively refer to the non-spherical perturbation force terms of the Earth and Mars. δ E and δ M are switching functions, which take 1 when the detector is within the sphere of influence of the Earth and Mars gravity and 0 otherwise. δV(t1,t2) represents the application and end of the deep space maneuver at times t1 and t2. F SRP represents the solar radiation pressure term, and F others represents other minor perturbation influence terms. r represents the vector radius from the geocenter to the center of mass of the launch vehicle in the geocentric inertial system; Combined with the requirements of departure from Earth and arrival at Mars, the particle swarm optimization algorithm is used to solve the optimization problem shown in formula (2), and the theoretical departure time of the probe is calculated: where δV pM represents the velocity increment required to enter the target Mars orbit, G1 represents the equality constraint that needs to be satisfied during the transfer process, and C3 E represents the Earth escape C3, and G2 represents the inequality constraint that needs to be satisfied during the transfer process; Combining Equation (1) and Equation (2), the position and velocity of the detector at the departure and arrival times are calculated where \(i = E\) and \(i = M\) represent the position and velocity at the time of departure from the Earth and arrival at Mars respectively, which can be mutually converted with the orbital elements of the detector, specifically as follows: H in Equation (3) pE represents the perigee altitude, C3 E represents the Earth escape C3, 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. In addition, the takeoff time of the launch vehicle is denoted as Tq, where H pE is usually approximately the parking orbit altitude, Ω E is directly related to the flight path of the launch vehicle, f E is determined by the characteristics of the launch vehicle flight orbit.

3. The design method of elements for taking off at any time within the Mars exploration launch window according to claim 2, characterized in that, In the launch inertial system, the flight dynamics equation of the rocket in the atmosphere is as follows: where \(r\) represents the vector from the geocenter to the center of mass of the launch vehicle in the inertial launch system, \(P\), \(R\), and \(g\) are the thrust, aerodynamic force, and gravitational acceleration vectors in the inertial launch system, and \(\psi(t)\) represents the pitch angle and yaw angle of the flight program. The pitch angle represents the angle between the thrust direction in the launch plane and the launch direction at the instant of launch, and the yaw angle represents the angle by which the thrust direction deviates from the launch plane at the instant of launch. During the flight, the thrust direction is controlled by the pitch and yaw program angles. Under the constraints of the orbital injection target, the launch trajectory design is completed with the maximum payload capacity as the optimization objective. The specific method is as follows: Adjust Wherein, Tg(t) represents the shutdown time of the powered flight phase, Th(Tq) represents the coasting time, the subscript 1 represents the orbit injection parameters achieved by the launch vehicle, the subscript 2 represents the orbit injection parameters required by the detector, and Ω E and f E are adopted in the iteration, and ε represents the desired iteration accuracy.

4. The element design method for achieving takeoff at any time of the Mars exploration launch window according to claim 3, wherein: The method for generating the launch trajectory is as follows: Using α o -α L to represent the difference between the right ascension of escape and the Greenwich sidereal hour angle of the launch site at the takeoff moment, then α o -α L and the relationship between the declination of escape, the azimuth of fire, and ΔTq relative to the takeoff moment is as follows: ΔTq = (α o - α L ) / ω e (6) where δ0 represents the escape declination, ω e represents the angular velocity of the Earth's rotation, and represents the latitude of the launch site; At the target escape declination, the launch time and the launch direction have a strong correlation, and i E = i E (A L (Tq)) (7) ω E = ω E (Tq, Th(Tq)) (8) That is, all the target elements can be expressed as continuous functions of the takeoff time. At the same time, due to the Earth launch property, applying the yaw angle ψ(Tq) in the engineering implementation can achieve the equivalence with and adjustment of the launch direction A. L (Tq). Specifically, by applying a constant yaw ψ i (t) in the second-stage acceleration working section of the last stage and adjusting the takeoff time, taxiing time, and shutdown time, the launch orbit of the launch vehicle can meet the orbit insertion requirements.

5. According to the method for designing parameters for taking off at any time in the Mars exploration launch window as described in claim 4, within the launch window, for any given launch time, the target orbit parameters are quickly calculated according to the bound fitting coefficients, and the carrier rocket is guided to fly according to the orbital elements obtained by the aiming fitting to achieve precise orbit entry.

6. A design system for elements of taking off at any time during the Mars exploration launch window, characterized in that include: The first calculation module is used to calculate the initial values ​​of the detector's departure parameters according to the launch time; Launch trajectory design module: used to design the launch vehicle orbit parameters according to the initial values ​​of the probe's departure parameters, and to complete the launch trajectory design through iteration by combining the calculated and designed quantities in the launch vehicle orbit parameters; The second calculation module: adjusts the target orbit inclination, calculates and generates the launch orbit target parameters with strong regularity of change with the take-off time; The third calculation module: It is used to fit the target elements of multiple launch orbits generated into a polynomial of the takeoff moment relative to the reference takeoff moment by using the least square principle, and obtain the fitting coefficients. Among them, the calculation quantities of the launch vehicle include the longitude of the ascending node and the true anomaly, and the design quantities include the perigee altitude, the escape C3, the orbital inclination, and the argument of perigee; fit the orbital inclination, the argument of perigee, and the coasting time at different takeoff moments within the window into a polynomial of the difference between the takeoff moment and the reference takeoff moment, and obtain the fitting coefficients. The calculation method is as follows: c = (A T A) -1 A T b(9) In the formula, c represents the fitting coefficient, b represents the result sequence of the sample points, and A is a matrix composed of the launch times. For n sample points, n≥4, taking the fitting of the orbital inclination by a cubic curve as an example, the expression is as follows: i EΔt = c3Δt 3 + c2Δt 2 + c1Δt + c0 (10) where Δt is the time difference between the launch time within the launch window and the reference launch time, and i EΔt is the orbital inclination corresponding to the launch when it is different from the reference launch time by Δt, and c k , k = 0, 1, 2, 3 are fitting coefficients. When there are no less than 4 sets of sample points, the fitting coefficients can be solved through Equation (9). Similarly, the fitting coefficients of the argument of perigee, escape C3, and coasting time can be calculated; The binding module: It is used to bind the obtained polynomial fitting coefficients into the launch vehicle; The fourth calculation module: It is used to calculate the real-time zero-window launch elements from the polynomial function according to the real-time launch window and the polynomial coefficients corresponding to each root number, so as to realize the real-time zero-window launch.

7. A terminal device, comprising a memory, a processor, and a computer program stored in the memory and capable of running on the processor, characterized in that, The memory stores a computer program that can run on a processor. When the processor loads and executes the computer program, it adopts the method for designing the elements for taking off at any moment of the Mars exploration launch window described in any one of claims 1 to 5.

8. A storage medium containing computer-executable instructions, characterized in that, The computer-executable instructions are used to execute the method for designing the elements for taking off at any moment of the Mars exploration launch window described in any one of claims 1 to 5 when executed by a computer processor.

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