General carrier rocket trajectory calculation method

Abstract

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

Description

Technical Field

[0001] The present invention relates to the technical field of trajectory calculation, and in particular to a universal launch vehicle trajectory calculation method. Background Art

[0002] In the development process of launch vehicles, the overall design needs to consider multiple key elements of multiple disciplines, such as rocket stages, diameter, stage ratio, power composition of each stage, and overall layout, and combine them into multiple overall plans. It also conducts comparative analysis of multiple plans, and completes iterative optimization of plans through comprehensive overall performance evaluation such as carrying capacity and overload. However, the traditional ballistic calculation method is compiled according to a certain overall plan of a launch vehicle. In the process of comparing and demonstrating multiple plans, it is necessary to provide a corresponding ballistic calculation method for each plan, which leads to a long trajectory calculation cycle and low work efficiency, which seriously affects the iterative efficiency of the overall plan demonstration of the launch vehicle. Summary of the invention

[0003] The present application aims to solve the technical problem that when comparing and demonstrating multiple schemes in the related art, a corresponding ballistic calculation method is provided according to each scheme, resulting in a long ballistic calculation cycle and low efficiency.

[0004] The present application embodiment provides a generalized launch vehicle trajectory calculation method, comprising the following steps:

[0005] Setting a target orbit and constraints corresponding to the target orbit, wherein the target orbit includes a sun-synchronous orbit or a low-inclination orbit, and the constraints include at least a geocentric vector radius, an absolute velocity, and a flight path angle;

[0006] Simplifying input parameters, wherein the input parameters include configuration parameters, mass parameters, aerodynamic parameters, and engine parameters;

[0007] Based on the constraint conditions of the target orbit, the corresponding control variables and the input parameters, the values ​​of the control variables are calculated by Newton's iteration method to obtain the trajectory of the launch vehicle.

[0008] In one embodiment, when the target orbit is a sun-synchronous orbit, the corresponding constraints include the geocentric vector radius, absolute speed and flight path angle, and the control variables include the maximum negative angle of attack of the programmed turn, the glide time and the payload mass.

[0009] In one embodiment, the launch vehicle is a fully solid launch vehicle or a launch vehicle that adopts a direct-to-orbit scheme and whose last stage is liquid.

[0010] In one implementation, when the target orbit is a sun-synchronous orbit, the Newton iteration method is a three-dimensional Newton iteration method.

[0011] In one embodiment, when the target orbit is a low-inclination orbit, the corresponding constraints include the geocentric vector radius, absolute velocity, flight path angle, and transfer orbit perigee height, and the control variables include the maximum negative angle of attack of the programmed turn, glide time, payload mass, and first working time.

[0012] In one embodiment, when the target orbit is a low-inclination orbit, the corresponding constraints include the geocentric vector radius, absolute speed, flight path angle, and total flight time, and the control variables include the maximum negative angle of attack of the programmed turn, the glide time, the payload mass, and the first working time.

[0013] In one embodiment, the launch vehicle is a launch vehicle that adopts a transfer-to-orbit scheme and whose last stage is liquid.

[0014] In one implementation, when the target orbit is a low-inclination orbit, the Newton iteration method is a four-dimensional Newton iteration method.

[0015] In one embodiment, the configuration parameters include any one of a one-stage and a half rocket, a two-stage rocket, a two-stage and a half rocket, a three-stage rocket, a three-stage and a half rocket, and a four-stage rocket; the mass parameters include the take-off mass excluding the payload corresponding to the configuration parameters; the aerodynamic parameters include the axial force coefficient and the aerodynamic reference area; the engine parameters include thrust type, thrust, propellant consumption per second, and nozzle area.

[0016] In one embodiment, the shape of the target orbit is circular or elliptical, and the orbit inclination is 0°-180°.

[0017] The beneficial effects brought by the technical solution provided in the embodiments of the present application include:

[0018] The present application provides a universal launch vehicle trajectory calculation method. Since a certain configuration rocket has the largest carrying capacity under a typical orbit, the carrying capacity of the configuration rocket to execute other target orbits is also better. By optimizing the typical target orbit, the target orbit range is narrowed, and the carrying capacity of the launch vehicle is guaranteed. Then, the input parameters of the launch vehicle trajectory calculation are simplified and sorted out, and the necessary parameters for the trajectory calculation and the relevant parameters of concern for the overall scheme demonstration are screened out. Finally, the trajectory is calculated by the Newton iteration method. Under the premise of ensuring the accuracy of the calculation, the adaptability to different configuration schemes is improved, and a set of programs is implemented to support the launch vehicle trajectory calculation process of multiple configuration schemes. Fast trajectory iteration calculation can be performed and the carrying capacity can be obtained, which is convenient for developers to select a better overall scheme. BRIEF DESCRIPTION OF THE DRAWINGS

[0019] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0020] Figure 1 The figure is a flow chart of a generalized launch vehicle trajectory calculation method in one embodiment of the present invention. DETAILED DESCRIPTION

[0021] In order to enable those skilled in the art to better understand the solution of the present application, the technical solution in the embodiments of the present application will be clearly and completely described below in conjunction with the drawings in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of this application.

[0022] like Figure 1 As shown, Figure 1 The figure is a flow chart of a generalized launch vehicle trajectory calculation method in one embodiment of the present invention.

[0023] This embodiment provides a generalized launch vehicle trajectory calculation method, comprising the following steps:

[0024] Step S1, setting a target orbit and constraints corresponding to the target orbit, the target orbit includes a sun-synchronous orbit or a low-inclination orbit, and the constraints include at least the geocentric vector radius, the absolute velocity, and the flight path angle;

[0025] Step S2, simplifying input parameters, the input parameters include configuration parameters, mass parameters, aerodynamic parameters, and engine parameters;

[0026] Step S3: Based on the constraints of the target orbit, the corresponding control variables, and the input parameters, the values ​​of the control variables are calculated by Newton's iteration method to obtain the trajectory of the launch vehicle.

[0027] This embodiment provides a universal launch vehicle trajectory calculation method for multidisciplinary optimization. Since a rocket of a certain configuration has the largest carrying capacity under a typical orbit, the carrying capacity of the rocket of this configuration for executing other target orbits is also relatively good. By optimizing the typical target orbit, the target orbit range is narrowed, and the carrying capacity of the launch vehicle is guaranteed. Then, the input parameters of the launch vehicle trajectory calculation are simplified and sorted out, and the necessary parameters for the trajectory calculation and the relevant parameters concerned by the overall scheme demonstration are screened out. Finally, the trajectory is calculated by the Newton iteration method. Under the premise of ensuring the accuracy of the calculation, the adaptability to different configuration schemes is improved, and a set of programs is implemented to support the launch vehicle trajectory calculation process of multiple configuration schemes. Fast trajectory iteration calculation can be performed and the carrying capacity can be obtained, which is convenient for developers to select a better overall scheme, improve the overall design performance of the launch vehicle system, and shorten the design cycle.

[0028] Each step is described and explained in detail below.

[0029] Step S1, setting a target orbit and constraints corresponding to the target orbit, the target orbit includes a sun-synchronous orbit or a low-inclination orbit, and the constraints include at least the geocentric vector radius, absolute velocity, and flight path angle.

[0030] In the early demonstration process of a certain launch vehicle, the target orbit was as simple and common as possible. Since the carrying capacity of the circular sun-synchronous orbit (SSO) at an altitude of 700 km and the carrying capacity of the nearly circular low-inclination orbit (LEO, a commonly used target orbit for manned space flight) with a perigee altitude of 200 km, an apogee altitude of 400 km and an inclination of 42° are more concerned, these two orbits are preferred as typical target orbits to ensure that the carrying capacity of the launch vehicle is better.

[0031] In one embodiment, when the target orbit is a sun-synchronous orbit, the corresponding constraints include the geocentric radius, the absolute velocity, and the flight path angle.

[0032] When the target orbit is a sun-synchronous orbit, the corresponding terminal constraints of the trajectory calculation include: the radius of the earth's center vector R0 at the entry point, the absolute velocity V0, the flight path angle F0, and the orbital inclination I0. Since only the motion within the launch plane is considered, the orbital inclination is almost only affected by the launch direction. Therefore, in order to simplify the calculation, only the three constraints of the radius of the earth's center vector R0 at the entry point, the absolute velocity V0, and the flight path angle F0 are adopted. For the case where the target orbit is a circular sun-synchronous orbit (SSO), a 191° launch direction is given, and the obtained orbital inclination is basically the same as the target orbital inclination, which has little impact on the carrying capacity and can greatly reduce the complexity of the calculation.

[0033] In one embodiment, the sun-synchronous orbit is suitable for a fully solid launch vehicle or a launch vehicle that adopts a direct-to-orbit scheme and whose last stage is liquid.

[0034] In one embodiment, when the target orbit is a low-inclination orbit, the corresponding constraints include the geocentric vector radius, the absolute velocity, the flight path angle, and the transfer orbit perigee altitude (or the total flight time).

[0035] When the target orbit is a low-inclination orbit, the corresponding terminal constraints of the trajectory calculation include: the radius of the earth's center vector R0 at the entry point, the absolute velocity V0, the flight path angle F0, and the orbital inclination I0. Since only the movement within the launch plane is considered, the orbital inclination is almost only affected by the launch direction. Therefore, in order to simplify the calculation, the radius of the earth's center vector R0 at the entry point, the absolute velocity V0, the flight path angle F0, and the transfer orbit perigee height (or total flight time) are taken as constraints. In the case where the target orbit is a near-circular low-inclination orbit (LEO), it is specified to be launched at a 90° launch direction at the Jiuquan Satellite Launch Center. The orbital inclination obtained is basically the same as the target orbital inclination, which has little impact on the carrying capacity and can greatly reduce the complexity of the calculation.

[0036] In one embodiment, a low inclination orbit (LEO) is suitable for a launch vehicle that adopts a transfer-to-orbit scheme and whose last stage is liquid.

[0037] Through the above scheme, the rocket types suitable for typical target orbits and the constraints corresponding to each target orbit are further clarified.

[0038] Step S2, simplifying input parameters, the input parameters include configuration parameters, mass parameters, aerodynamic parameters, and engine parameters.

[0039] Specifically, simplify the flight timing: Based on the needs of ballistic expertise and demonstration, the flight timing of the passive stage is simplified. For most rockets currently under demonstration, the stage separation and shield ejection timings are located in the passive stage and the atmospheric density is relatively thin. These timings have little effect on the carrying capacity. Therefore, the corresponding timings can be merged. That is, after the engine of the upper stage is exhausted, the fully loaded takeoff mass of the lower stage is used as the initial value of the ballistic integral in the ballistic calculation process after a predetermined period of unpowered flight.

[0040] In terms of ballistic scheme, the sun-synchronous orbit (SSO) adopts the transfer-to-orbit scheme, that is, the last-stage engine works twice, enters the elliptical transfer orbit after the first work, and enters the target orbit after the second work; the low-inclination orbit (LEO) adopts the direct-to-orbit scheme, that is, the last-stage engine works once. For N-stage rockets, the flight sequence is simplified to: first-stage active segment (may include boosters), first-stage passive segment, second-stage active segment, second-stage passive segment, ..., N-1-stage passive segment, N-stage active segment 1, N-stage passive segment (LEO does not have this sequence), N-stage active segment 2 (LEO does not have this sequence).

[0041] The dynamic model of the rocket is appropriately simplified. For example, in the demonstration phase, the movement within the launch plane is mainly considered. The ballistic calculation model is simplified. For example, based on the principle of instantaneous balance, the action of the air rudder (or grid rudder) or the swing of the engine nozzle is not considered, and the dynamics is simplified to a differential equation in a two-dimensional plane. Only the movement within the XOY plane of the launch system is considered, and the Z-direction movement is not considered.

[0042] In combination with the existing models under development and pre-research models, all input parameters for ballistic calculations are sorted out, summarized and simplified: parameters of overall professional concern and parameters that have a significant impact on the ballistic calculation results must be retained, such as rocket configuration, shape and mass parameters of each sub-stage, engine performance, etc.; parameters that are of less overall professional concern and have little impact on the ballistic calculation are appropriately simplified or deleted, such as deleting the transverse normal force coefficient in the aerodynamic performance.

[0043] In one embodiment, the configuration parameters include any one of a one-and-a-half-stage rocket, a two-stage rocket, a two-and-a-half-stage rocket, a three-stage rocket, a three-and-a-half-stage rocket, and a four-stage rocket; the mass parameters include the take-off mass excluding the payload corresponding to the configuration parameters; the aerodynamic parameters include the axial force coefficient and the aerodynamic reference area; the engine parameters include thrust type, thrust, propellant consumption per second, and nozzle area.

[0044] Specifically, the configuration parameters are used to define the configuration of the rocket; the mass parameters are used to define the takeoff mass of each stage without payload, and the dimension of the parameters corresponds to the number of stages of the rocket; the aerodynamic parameters are used to define the aerodynamic performance of each stage of the rocket. Since the Z-direction motion is not considered, the program turning section mainly relies on the component of the engine thrust on the Y-axis of the velocity system, and the zero angle of attack flight is maintained after the program turning, the aerodynamic parameters only consider the axial force coefficient. Therefore, the corresponding aerodynamic parameters can be simplified into the axial force coefficient plus the aerodynamic reference area. Considering that only the aerodynamics are estimated in the initial demonstration stage, the aerodynamic parameters of the same benchmark rocket are used, and the axial force amplification coefficient is multiplied on the basis of the benchmark; the engine parameters are used to define the performance of the engines at each stage. The engine performance includes thrust type (sea level or vacuum), thrust, propellant consumption per second, and nozzle area. For solid engines with internal ballistic data, the internal ballistic data is directly used. For solid engines and liquid engines without internal ballistic data, the conversion is based on the propellant mass, thrust and specific impulse.

[0045] In one embodiment, the shape of the target orbit is circular or elliptical, and the orbital inclination is 0°-180°. The rocket type can also be pure solid, pure liquid, or a mixture of solid and liquid.

[0046] Through the above scheme, only necessary input parameters are retained, which is more universal and can adapt to different rocket configuration schemes, engine types, aerodynamic shapes and target orbits.

[0047] When the above input parameters are used for trajectory calculation, they are transmitted through interface files, such as overall parameter files and engine parameter files.

[0048] (1) Overall parameter file

[0049] The overall parameter file format convention is shown in Table 1.

[0050] Table 1 Overall parameter file format

[0051] G0 G1 G2 G3 G4 G5 M0 M1 M2 M3 M4 M5 S0 S1 S2 — — — K0 K1 K2 — — —

[0052] in:

[0053] G0~G5 are configuration identifiers, representing whether the rocket has a booster stage, a first stage, a second stage, a third stage, a fourth stage, and a final stage. When Gi(i=0,1,2,3,4,5)=0, it means that the rocket does not contain the i-th stage. In particular, when G0=1, it means that there is a booster, and the booster is turned on first, and the first stage is ignited and started after the work is completed.

[0054] M0~M5 represent the take-off masses (excluding payload mass) of the booster stage, first stage, second stage, third stage, fourth stage and final stage respectively.

[0055] S0, S1, and S2 represent the aerodynamic reference areas of the booster stage, the first stage, and the second stage, respectively; the third stage and beyond are located in the vacuum section, so the aerodynamics are not considered after the third stage;

[0056] K0, K1, and K2 represent the axial force amplification coefficients of the booster stage, first stage, and second stage respectively; the third stage and beyond are located in the vacuum section, so aerodynamics are not considered after the third stage.

[0057] (2) Engine parameter file

[0058] The engine file includes 6 files, named FDJ0, FDJ1, FDJ2, FDJ3, FDJ4, FDJ5, corresponding to the booster stage, the first stage, the second stage, the third stage, the fourth stage and the final stage. When the rocket does not have a certain stage, the program does not read the corresponding internal ballistic data. The format of each engine file is shown in Table 2.

[0059] Table 2 Engine file format

[0060]

[0061] For solid rocket engines with internal ballistic data, simply replace the 2nd to Nth lines; for solid engines and liquid engines without internal ballistic data, directly use the average thrust and specific impulse. In this case, the internal ballistic data only needs to write 2 lines of internal ballistics, namely the ignition time and the shutdown time.

[0062] Through the above scheme, a unified interface file is used for the selected input parameters, which can adapt to different types and levels of launch vehicles.

[0063] Step S3: Based on the constraints of the target orbit, the corresponding control variables, and the input parameters, the values ​​of the control variables are calculated by Newton's iteration method to obtain the trajectory of the launch vehicle.

[0064] In one embodiment, when the target orbit is a sun-synchronous orbit, the control variables include the maximum negative angle of attack of the programmed turn, the glide time, and the payload mass, which are iteratively calculated using a three-dimensional Newton iteration method.

[0065] In one embodiment, when the target orbit is a low-inclination orbit, the control variables include the maximum negative angle of attack of the programmed turn, the glide time, the payload mass and the first working time, which are iteratively calculated using a four-dimensional Newton iteration method.

[0066] Specifically, based on the constraints of the target trajectory, the corresponding control variables, and the input parameters, a set of control variable guess values ​​are given through experience, and the trajectory is iteratively calculated with the control variable guess values.

[0067] Whether it is a trajectory calculation problem with three control variables corresponding to three constraints or a trajectory calculation problem with four control variables corresponding to four constraints, it is abstracted into a mapping from n-dimensional independent variables to n-dimensional dependent variables, that is, a functional relationship, which can be described in mathematical language as follows:

[0068] The independent variable x consists of n control variables = (x1, x2, ..., x n ) T ;

[0069] The dependent variable y consists of n constraints = (y1, y2, ..., y n ) T ;

[0070] The mapping from x to y is the function y=f(x) of trajectory calculation, that is, for any independent variable x, a unique y can be obtained.

[0071] To achieve the desired trajectory constraint y e (the aforementioned constraints), the corresponding control variable x needs to be solved e , that is, determine the trajectory of the launch vehicle. This engineering problem is a strongly nonlinear equation, which is solved by Newton iteration method (when the independent variable is 3 and 4, it corresponds to three-dimensional Newton iteration and four-dimensional Newton iteration respectively). The specific steps are as follows:

[0072] Step a: For the expected y e =f(x e ), a control variable guess value x0 is given based on experience,

[0073] Step b: Calculate y0=f(x0), and proceed to step c;

[0074] Step c: When ||y0-y e When ||≤Δ (Δ is the allowable error vector), the calculation ends and x0 is the desired value; when ||y0-y e When ||>Δ, go to step d;

[0075] Step d: Correct the variables as follows:

[0076] In the above formula: x0 on the right side of the equal sign is the independent variable of the previous round, f(x0) is the dependent variable corresponding to the previous round, and f′(x0) is the derivative of the previous round; the left side of the equal sign is the updated independent variable x0, and return to step b.

[0077] The specific calculation process of the Newton iteration method is an existing technology and will not be repeated here. The final output is the required ballistic data and optimization target (such as payload mass or remaining propellant mass).

[0078] In the description of the present application, it should be noted that the terms "upper", "lower", etc. indicate orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, rather than indicating or implying that the device or element referred to must have a specific orientation, be constructed and operated in a specific orientation, and therefore cannot be understood as a limitation on the present application. Unless otherwise clearly specified and limited, the terms "installed", "connected", and "connected" should be understood in a broad sense, for example, it can be a fixed connection, a detachable connection, or an integral connection; it can be a mechanical connection or an electrical connection; it can be a direct connection, or it can be an indirect connection through an intermediate medium, or it can be a connection between the two elements. For those of ordinary skill in the art, the specific meanings of the above terms in the present application can be understood according to the specific circumstances.

[0079] It should be noted that, in this application, 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 such actual relationship or order between these entities or operations. Moreover, the terms "include", "comprise" or any other variants thereof are intended to cover non-exclusive inclusion, so that a process, method, article or device including a series of elements includes not only those elements, but also other elements not explicitly listed, or also includes elements inherent to such process, method, article or device. In the absence of further restrictions, the elements defined by the sentence "comprise a ..." do not exclude the presence of other identical elements in the process, method, article or device including the elements.

[0080] In the description of the embodiments of the present application, unless otherwise specified, “ / ” means or, for example, A / B can mean A or B; the “and / or” in the text is merely a description of the association relationship of associated objects, indicating that three relationships may exist, for example, A and / or B can mean: A exists alone, A and B exist at the same time, and B exists alone. In addition, in the description of the embodiments of the present application, “multiple” refers to two or more than two.

[0081] The above is only a specific implementation of the present application, so that those skilled in the art can understand or implement the present application. Various modifications to these embodiments will be apparent to those skilled in the art, and the general principles defined herein can be implemented in other embodiments without departing from the spirit or scope of the present application. Therefore, the present application will not be limited to the embodiments shown herein, but will conform to the widest range consistent with the principles and novel features applied for herein.

Claims

1. A universal launch vehicle trajectory calculation method, characterized in that: The following steps are involved: Setting a target orbit and constraints corresponding to the target orbit, wherein the target orbit includes a sun-synchronous orbit or a low-inclination orbit, and the constraints include at least a geocentric vector radius, an absolute velocity, and a flight path angle; Simplifying input parameters, wherein the input parameters include configuration parameters, mass parameters, aerodynamic parameters, and engine parameters; Based on the constraint conditions of the target orbit, the corresponding control variables and the input parameters, the values ​​of the control variables are calculated by Newton's iteration method to obtain the trajectory of the launch vehicle.

2. The universal launch vehicle trajectory calculation method according to claim 1, characterized in that: When the target orbit is a sun-synchronous orbit, the corresponding constraints include the geocentric vector radius, the absolute speed and the flight path angle, and the control variables include the maximum negative angle of attack of the programmed turn, the glide time and the payload mass.

3. The universal launch vehicle trajectory calculation method as claimed in claim 2, characterized in that: The carrier rocket is a fully solid carrier rocket or a carrier rocket that adopts a direct-to-orbit scheme and whose last stage is liquid.

4. The universal launch vehicle trajectory calculation method as claimed in claim 2, characterized in that: When the target orbit is a sun-synchronous orbit, the Newton iteration method is a three-dimensional Newton iteration method.

5. The universal launch vehicle trajectory calculation method according to claim 1, characterized in that: When the target orbit is a low-inclination orbit, the corresponding constraints include the geocentric vector radius, absolute velocity, flight path angle, and transfer orbit perigee height; and the control variables include the maximum negative angle of attack of the programmed turn, the glide time, the payload mass, and the first working time.

6. The universal launch vehicle trajectory calculation method according to claim 1, characterized in that: When the target orbit is a low-inclination orbit, the corresponding constraints include the geocentric vector radius, absolute speed, flight path angle, and total flight time, and the control variables include the maximum negative angle of attack of the programmed turn, the glide time, the payload mass, and the first working time.

7. The universal launch vehicle trajectory calculation method according to claim 5 or 6, characterized in that: The launch vehicle is a launch vehicle that adopts a transfer-to-orbit scheme and has a liquid last stage.

8. The universal launch vehicle trajectory calculation method according to claim 5 or 6, characterized in that: When the target orbit is a low-inclination orbit, the Newton iteration method is a four-dimensional Newton iteration method.

9. The universal launch vehicle trajectory calculation method according to claim 1, characterized in that: The configuration parameters include any one of a one-and-a-half-stage rocket, a two-stage rocket, a two-and-a-half-stage rocket, a three-stage rocket, a three-and-a-half-stage rocket, and a four-stage rocket; the mass parameters include the take-off mass excluding the payload corresponding to the configuration parameters; the aerodynamic parameters include the axial force coefficient and the aerodynamic reference area; the engine parameters include thrust type, thrust, propellant consumption per second, and nozzle area.

10. The universal launch vehicle trajectory calculation method according to claim 1, characterized in that: The shape of the target orbit is circular or elliptical, and the orbit inclination is 0°-180°.

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