A rapid simulation modeling method for the whole process dynamics of launch vehicles based on the Modelica unified modeling language

Abstract

This paper proposes a rapid simulation method for the full-process dynamics of a launch vehicle based on the Modelica language, aiming to address the limited model reusability and adaptability of traditional simulation methods. By leveraging the object-oriented nature of the Modelica language, this method constructs a comprehensive dynamic model that integrates multiple subsystems, including the propulsion system, control system, and aerodynamic model. The specific steps include: first, constructing an overall model of the launch vehicle and defining time-varying parameters to accurately calculate key parameters such as propellant consumption, moment of inertia, and center of mass position; second, analyzing the impact of different design parameters on rocket performance through simulation to proactively identify and resolve potential issues; and finally, analyzing the simulation results to ensure the accuracy and reliability of the model. The innovation of this method lies in its modular design and multi-domain collaborative simulation capabilities, which not only significantly improve the model's reusability and adaptability but also effectively support rocket design optimization, mission planning, and control strategy formulation. This method not only reduces R&D costs but also accelerates the advancement of aerospace technology, possessing broad application prospects.

Description

Technical Field

[0001] The present invention relates to the field of aerospace engineering, and in particular to a fast dynamics simulation method for a launch vehicle, in particular to a full-process simulation technology based on the Modelica language. Background Art

[0002] A launch vehicle is a multi-stage space transport vehicle used to deliver payloads such as artificial satellites, manned spacecraft, space stations, and space probes into a predetermined orbit. The design, performance evaluation, and testing of launch vehicles are critical components of aerospace engineering. Dynamics simulation can help engineers analyze the performance of a rocket under different design parameters during the design phase, thereby optimizing the design and ensuring flight safety and reliability.

[0003] Although relatively comprehensive mathematical models are currently available for rocket dynamics simulation, building simulation models for specific problems in practice remains a time-consuming and labor-intensive task. Significant differences exist between different models and applications, and traditional software design methods struggle to provide a high level of abstraction. This results in low model reusability and an inability to flexibly adapt to new situations.

[0004] The full-process simulation process involves multiple disciplines, including structural dynamics, multibody dynamics, attitude control, and aerodynamics. It is crucial for aspects such as launch vehicle layout and shape selection, as well as parameter design, and requires the integration of multiple disciplines for multi-physics coupled simulation. It should be understood that the Modelica language is specifically designed for unified modeling and simulation across multiple domains and offers unique advantages in multi-physics coupled simulation. Due to its object-oriented, equation-based, reusable model, and hierarchical structure, the Modelica language effectively enables full-process system modeling of launch vehicles. Therefore, a solution is urgently needed to construct a full-process simulation model for launch vehicles based on the Modelica unified modeling language. Summary of the Invention

[0005] The purpose of the present invention is to provide a method for rapid simulation of the full-process dynamics of a launch vehicle based on the Modelica language, and to establish a full-process simulation model of a launch vehicle based on a modeling method of the Modelica causal modeling language.

[0006] The technical solution for achieving the purpose of the present invention is: a method for rapid simulation of the full-process dynamics of a launch vehicle based on the Modelica language, comprising the following steps:

[0007] Step 1: First, construct the parametric rigid body models of the first and second stages of the launch vehicle with time-varying parameters. The mass, center of mass, and moment of inertia of the first and second stages of the launch vehicle are dynamically changed over time. Then, for the satellite and fairing with time-invariant parameters, respectively, the parametric rigid body models of the satellite and fairing are established.

[0008] Step 2: Based on the selected rocket engine model and the corresponding guidance, navigation and control system GNC, the functional units of the launch vehicle's first and second stage engine models, the swing engine control system model, the geophysical model, the timing model, the atmospheric parameter model, the launch point information model and the aerodynamic model are established through the object-oriented modeling method. The two functional units of the geophysical model and the launch point information model are used to configure a gravity model with optional gravity field types. The gravity field types can be divided into zero gravity field, uniform gravity field and point gravity field. The point gravity field type is the earth's gravity field considering the J2 item. Finally, the force model acting during the first and second stage separation, fairing separation and satellite-rocket separation is established.

[0009] Step 3: Connect the first and second stage parameterized rigid body models of the launch vehicle's time-varying parameters, the satellite rigid body model, and the fairing rigid body model through connectors in the Modelica standard library. Based on the connectors, constrain the motion parameters such as displacement, force, and torque to construct a constraint device model. This constraint device model can be reused during the first and second stage separation, fairing separation, and satellite-rocket separation after calling different separation signals.

[0010] Step 4. Based on the above rigid body model, functional unit and constraint device model, use the building block and connection method to close the input and output parameters of each module and quickly build a full-process dynamic simulation model of the launch vehicle.

[0011] Furthermore, the parametric rigid body models of the first and second sub-stages of the launch vehicle with time-varying parameters described in step one include the parameterized design parameters of time-varying center of mass, time-varying mass, time-varying moment of inertia, length, and diameter, and the output parameters are speed, position, acceleration, and attitude angle. Subsequently, the satellite and fairing rigid body models are parameterized with time-invariant center of mass, time-invariant mass, time-invariant moment of inertia, length, and diameter as design parameters, and the output parameters are speed, displacement, and acceleration.

[0012] Furthermore, the functional units in step 2 include: a launch vehicle first and second stage engine model, a swing engine control system model, an aerodynamic model, an atmospheric parameter model, a geophysical model, a launch point information model, and a timing model;

[0013] The functional unit is constructed by object-oriented modeling method; the input parameters of the first and second stage engine model of the carrier rocket are the atmospheric pressure at the current altitude, the engine swing angle and the rocket attitude angle, the output parameter is the thrust vector, and the design parameters are the engine second consumption, jet speed, and nozzle area; the input parameters of the swing engine control system model are the attitude angles of the first and second stages, and the output parameters are the swing angles of the first and second stage engines of the carrier rocket, and the design parameters are the static amplification coefficients of the pitch, yaw, and roll channels; the timing model includes the design parameters of the first and second stage engine power on and off time, the engine full start time, the first and second stage separation, the fairing separation, and the satellite-rocket separation time. The first and second stage engines and the constraint device of the carrier rocket are called through the Constant function to complete the control of the flight progress by the timing model; the input parameters of the atmospheric parameter model are the rigid body position vectors of the first and second stage of the carrier rocket, and the output parameters are the atmospheric pressure at the current altitude, the atmospheric density, and the atmospheric pressure at the current altitude. degree and atmospheric temperature, and the design parameters are the atmospheric pressure and density at zero sea level; the input parameters of the aerodynamic model are the atmospheric density at the current altitude, the velocity vector of the first stage, the direction vector of the rocket axis, and the attitude angle of the first stage; the output parameters are the aerodynamic vector in the rocket body coordinate system, and the design parameters are the aerodynamic reference length and area of ​​the rocket; in the gravity model configured by the two functional units of the geophysical model and the launch point information model, the design parameters in the geophysical model and the launch point information model are used as Parameter parameters to participate in the particle gravity field calculation, so that the gravity can be automatically calculated by dragging any rigid body model into the full-process simulation model containing the gravity model; the force model acting during the separation of the first and second stages, the separation of the fairing, and the separation of the satellite and rocket includes the reverse thrust on the first stage rigid body and the forward thrust on the second stage rigid body during the separation of the first and second stages, the separation force applied to the fairing during separation, and the separation force applied to the satellite during separation of the satellite and rocket.

[0014] Furthermore, the constraint device model in step 3 is composed of a constraint device, a signal generator, and a signal source. The two ends of the constraint device are the action points of the physical frame. Before separation, there is no relative motion or relative rotation between the first and second sub-stages, the satellite, and the fairing. The constraint equation is described as

[0015]

[0016] Where r A 、r B is the position vector of the entity frame interface at both ends of the rigid body to be connected, θ A ,θ B For the attitude angles at both ends of the rigid body to be connected, calculate the first and second order derivatives of the position vector and attitude angle with respect to time;

[0017]

[0018] This is used to constrain the relative position, velocity and attitude of the first and second sub-stages, satellites and fairing rigid bodies, and to enable the constraint device to release the constraint when receiving the signal from the timing model.

[0019] Furthermore, the rigid body models described in step 3 are connected via connectors, wherein the connectors include flow variables and potential variables, and the sum of the flow variables is zero while the potential variables are equal;

[0020] A mathematical model is established based on the physical principles, force balance equations, and torque balance equations of the full-process dynamics rapid simulation model of the launch vehicle; a Modelica model is adopted, and code development and interface definition are performed based on a standard library; wherein the interface is used to transmit force, torque, and attitude angle.

[0021] Furthermore, the launch vehicle full-process dynamics simulation model described in step 4 includes at least the following sub-models:

[0022] Parametric model of the time-varying parameter rigid body structure of the first and second sub-stages, parametric sub-model of the satellite fairing rigid body structure, timing model, constraint device model, first and second stage engine model of the launch vehicle, swing engine control system model, atmospheric parameter model, aerodynamic model, and force model of each separation stage.

[0023] Compared with the prior art, the present invention can achieve the following beneficial effects:

[0024] 1) Developed based on the Modelica language, the launch vehicle model library can be developed with clear model hierarchy, reusable and extensible models, greatly improving modeling efficiency;

[0025] 2) It can quickly perform numerical simulation of the entire process dynamics of the launch vehicle, increase the flexibility and scalability of the model, shorten the development cycle of the launch vehicle and reduce costs. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 The present invention is a flow chart of a launch vehicle full-process dynamics simulation modeling method based on the Modelica unified modeling language.

[0027] Figure 2 The present invention provides a full-process dynamics simulation theoretical model of a launch vehicle based on the Modelica unified modeling language.

[0028] Figure 3 The invention discloses a rigid body theoretical model of the first and second substages of a launch vehicle with time-varying parameters based on the Modelica language.

[0029] Figure 4 The invention discloses a launch vehicle separation and restraint device model based on the Modelica language.

[0030] Figure 5 The present invention provides a model of the first and second stage engines of a launch vehicle based on the Modelica language and a theoretical calculation curve of the model; Figure 5 (a) is a schematic diagram of the first and second stage engine models of a launch vehicle based on the Modelica language. Figure 5 (b) in the figure is the theoretical thrust change curve of the first and second stage engine models of the carrier rocket during the flight of the carrier rocket.

[0031] Figure 6 The present invention is a model of the launch vehicle swing engine control system based on the Modelica language and a schematic diagram of the swing engine control principle; Figure 6 (a) is a schematic diagram of the launch vehicle swing engine control system model based on the Modelica language; Figure 6 (b) in the figure is a schematic diagram of the engine swing of the swing engine control system model.

[0032] Figure 7 The present invention provides a gravity field model based on the Modelica language and a component parameter interface for selecting gravity field types of the model; Figure 7 (a) is the gravity field model component parameter selection interface. Click to select the gravity field model to be used and it will be automatically updated. Figure 7 (b) is a schematic diagram of the gravity field model based on the Modelica language of the present invention. By simply dragging the gravity field model icon with set component parameters into the target simulation model, the gravity of all rigid body models in the target simulation model can be automatically calculated.

[0033] Figure 8 This is a large launch vehicle simulation result diagram drawn by the launch vehicle full-process dynamics rapid simulation model based on the Modelica language of the present invention; Figure 8 (ab) in the figure is the center of mass change curve of the first and second stages of the launch vehicle; Figure 8 (cd) in the figure is the curve of the moment of inertia change of the first and second stages of the launch vehicle; Figure 8 (ef) in the figure is the pitch angle variation curve of the first and second stages of the launch vehicle; Figure 8 (gh) in the figure is the thrust variation curve of the first and second stage engines of the launch vehicle; Figure 8 (ij) in the figure is the lift and drag variation curve of the launch vehicle. DETAILED DESCRIPTION

[0034] The present invention provides a method for rapid simulation of the full-process dynamics of a launch vehicle based on the Modelica language, comprising the following steps: step 1, rigid body modeling of the launch vehicle satellite, fairing, and first and second sub-stages with timely variable center of mass and moment of inertia; step 2, modeling functional units such as the first and second sub-stages, control system, aerodynamic module, and reverse thrust rocket; step 3, constructing constraint device models between the first and second sub-stages, satellite, and fairing rigid bodies; step 4, based on the rigid body models, functional units, and constraint device models, a full-process dynamics simulation model of the launch vehicle is rapidly constructed by building blocks and connecting lines.

[0035] Through dynamic simulation, the rocket's performance under different design parameters is analyzed. The simulation results are then theoretically analyzed to verify the model's accuracy, ensuring it can accurately predict the rocket's flight trajectory, velocity, and attitude changes. This increases the model's flexibility and scalability, shortening the launch vehicle's development cycle and reducing costs. Specifically, the full-process dynamic simulation model is parameterized, and key data such as the rocket's position, velocity, acceleration, attitude, and engine operating status are recorded during the simulation. The processed data is then visualized, such as plotting the rocket's flight trajectory, velocity curves, acceleration curves, and attitude angle curves, to intuitively analyze the rocket's flight performance.

[0036] In the present invention, the functional units include: a model for the tumbling engine control system that affects the launch vehicle's engine thrust, changes the thrust vector, an aerodynamic force model, and an atmospheric parameter model. The launch vehicle's full-process simulation model includes at least the following submodels: a parameterized submodel for the time-varying first and second stage rigid body structures, a satellite rigid body model, a fairing rigid body model, a timing model, launch vehicle first and second stage engine models, a tumbling engine control system model, an atmospheric parameter model, an aerodynamic force model, and a reverse thrust rocket model.

[0037] Specifically, the time-varying parameter parameterized sub-model of the rigid body structure of the first and second sub-stages includes the time-varying center of mass, time-varying mass, time-varying moment of inertia, length, diameter and other parameters of the first and second sub-stages; the timing model includes the parameters such as the start and shutdown time of the first and second stage engines, the time when the engines are fully started, the separation of the first and second stages, the separation of the fairing, and the separation of the satellite and rocket; the thrust model of the first and second stage engines includes the parameters such as the second consumption, jet speed, jet pressure, nozzle area, etc.; the swing engine control system model includes parameters such as the equivalent swing angle of the engine; the atmospheric parameter model includes parameters such as temperature, pressure, atmospheric density, etc.; the aerodynamic model includes parameters such as the characteristic area of ​​the rocket and the aerodynamic length.

[0038] The constraint device model in step three consists of a constraint device, a signal generator, and connectors with the first and second sub-stages, satellites, and fairing rigid bodies. It is used to constrain the relative position, velocity, and attitude of the first and second sub-stages, satellites, and fairing rigid bodies, and enable the constraint device to release the constraint when receiving a signal.

[0039] Furthermore, in step four, the following is performed: a mathematical model is established based on the physical principles, force balance equations, and torque balance equations of the full-process dynamics simulation model; the Modelica model is used to develop code and define interfaces based on the standard library, wherein the interface is used to transmit force, torque, height, and attitude angle, etc.; the parameterized sub-models are connected by connectors, which include flow variables and potential variables, and the sum of the flow variables is zero while the potential variables are equal.

[0040] Furthermore, the parameters of the full-process simulation model are configured. First, the component parameters of the first and second sub-stages, satellite, and fairing rigid body models are opened. Only the numbers need to be modified to set it to a certain model of launch vehicle. Then, the design parameters of the functional units are modified separately to determine the engine and control system. The flight trajectory, velocity curve, acceleration curve, and attitude angle curve of the current model and control are simulated.

[0041] In order to make the purpose, technical solutions and advantages of this application more clear, the following further describes this application in detail with reference to the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain this application and are not intended to limit this application.

[0042] A theoretical model for the full-process dynamics simulation of a launch vehicle based on the Modelica unified modeling language is as follows: Figure 2 As shown, the following steps are included:

[0043] Step 1: Model the rigid bodies of the first and second stages of the launch vehicle with time-varying parameters, specifically:

[0044] During the flight, the launch vehicle not only consumes a large amount of fuel, resulting in time-varying system mass, center of mass, and moment of inertia, but also undergoes multiple separation processes such as separation of the first and second stages, separation of the fairing, and separation of the satellite and rocket. These time-varying characteristics will inevitably affect the dynamic performance of the launch vehicle. In order to solve the time-varying mass problem, the first-stage flight segment is used as an example to derive time-varying parameters such as the center of mass and moment of inertia, in preparation for solving the system dynamics equations. Other flight segments can be derived similarly. Assume that the launch vehicle consists of seven parts, namely the first and second stages, the first and second stage fuel, two fairings, and a satellite. Figure 3 Without loss of generality, it is assumed that the tank and the fuel are both cylindrical.

[0045] Step 1.1, time-varying mass rigid body theoretical model:

[0046] The remaining mass of the first-stage propellant after the first-stage engine has worked for t seconds is expressed as

[0047]

[0048] In the formula is the initial mass of the first-stage propellant, Engine consumption per second

[0049] The time-varying mass of the launch vehicle system can be expressed as

[0050]

[0051] Where M R1 、M R2 They are the first and second level rocket masses respectively. is the initial mass of the secondary propellant, M F is the mass of a single fairing, M s Satellite quality.

[0052] Step 1.2, time-varying center of mass rigid body theoretical model:

[0053] The length of propellant consumed is

[0054]

[0055] In the formula is the initial length of the first stage propellant. The length of the remaining propellant

[0056]

[0057] Time-varying distance between the propellant's center of mass and the theoretical cusp

[0058]

[0059] In the formula is the distance between the center of mass of the first-stage propellant and the theoretical cusp at the initial moment.

[0060] According to the center of mass calculation formula, and considering the first and second stages, the first and second stage fuel, two fairings and a satellite, the distance between the center of mass of the launch vehicle time-varying system (i.e. the origin of the body coordinate system O1) and the theoretical cusp can be expressed as

[0061]

[0062] Where D R1 、D R2 They are the theoretical distances from the center of mass of the first and second stage arrows to the cusp point, D is the distance between the center of mass of the secondary propellant and the theoretical cusp at the initial moment. F is the distance from the fairing to the theoretical apex along the long axis (assuming the two fairings of the launch vehicle are exactly the same), D s is the distance between the satellite's centroid and the theoretical cusp.

[0063] Step 1.3, time-varying moment of inertia rigid body theoretical model:

[0064] According to the parallel axis theorem of moment of inertia, the moment of inertia of the system around the long axis O1X1 of the rocket body can be expressed as

[0065]

[0066] Where I′ Rx1 , I′ Rx2 are the moments of inertia of the first and second stage rocket bodies relative to their own long axis (assuming their long axis coincides with O1X1), R O1 、R O2 I′ is the radius of the first and second stage propellant tanks, which are approximately equal to the radius of the first and second stage rocket bodies; Fx M is the moment of inertia of the fairing relative to the X axis of its own body coordinate (assuming it is parallel to O1X1), F is the mass of a single fairing, R F I' is the distance from the center of mass of the fairing to O1X1; Sx It is the rotation of the satellite's moment of inertia relative to its own body coordinate X axis (assuming it coincides with the long axis of the rocket body).

[0067] The moment of inertia of the first-stage rocket body relative to the O1Y1 axis of the body coordinate system is

[0068] I Ry1 =I′ Ry1 +M R1 (D System -D R1 ) 2 (8)

[0069] Where I′ Ry1 It is the moment of inertia of the first-stage rocket body relative to the Y-axis of its own coordinate system (parallel to the O1Y1 axis).

[0070] The moment of inertia of the remaining fuel in the first stage relative to the O1Y1 axis of the body coordinate system is:

[0071]

[0072] The moment of inertia of the secondary rocket body relative to the O1Y1 axis is

[0073] I Ry2 =I′ Ry2 +M R2 (D System -D R2 ) 2 (10)

[0074] Where I′ Ry2 It is the moment of inertia of the secondary rocket body relative to the Y axis of its own coordinate system (parallel to the O1Y1 axis).

[0075] The moment of inertia of the secondary fuel relative to the O1Y1 axis is

[0076]

[0077] The moment of inertia of the two fairings relative to the O1Y1 axis is

[0078] I Fy =2(I′ Fy +M F (D System -D F ) 2 ) (12)

[0079] Where I′ Fy is the moment of inertia of a single fairing relative to the Y-axis of its own body coordinate system (parallel to the O1Y1 axis).

[0080] The satellite's moment of inertia relative to the O1Y1 axis is

[0081] I Sy =I S ' y +M S (D System -D S ) 2 (13)

[0082] Where I′ Sy is the moment of inertia of the satellite relative to the Y axis of its own body coordinate system (parallel to the O1Y1 axis).

[0083] The combined formula ~ formula can be expressed as the moment of inertia of the launch vehicle system around the O1Y1 axis:

[0084] I Systemy =I Ry1 +I Oy1 +I Ry2 +I Oy2 +I Fy +I Sy (14)

[0085] Similarly, the moment of inertia of the launch vehicle system around the O1Z1 axis can be expressed as

[0086] I Systemz =I Rz1 +I Oz1 +I Rz2 +I Oz2 +I Fz +I Sz (15)

[0087] Where I Rz1 , I Rz2 are the moments of inertia of the first and second stage rocket bodies relative to the O1Z1 axis of the body coordinate system, I Oz1 , I Oz2are the moment of inertia of the remaining fuel in the first stage and the fuel in the second stage about the O1Z1 axis, I Fz , I Sz are the moments of inertia of the two fairings and the satellite about the O1Z1 axis respectively, and the derivation process will not be repeated here.

[0088] The time-varying inertia matrix of the launch vehicle system can be expressed as

[0089]

[0090] Step 2: Model the launch vehicle's first and second stage engines, swing engine control system and other functional units. Specifically:

[0091] Step 2.1: Modeling the first and second stage engines of the launch vehicle:

[0092] During the static test of the engine on the ground, in addition to the existence and mutual cancellation of gravity and the reaction force of the test bench, there is only axial force. It should be noted that this axial force is not a simple relative force. It also includes the axial force generated by the static pressure of the atmosphere on the surface of the rocket body and the static pressure of the gas on the engine nozzle section. These two parts of static pressure are called static thrust and are recorded as

[0093]

[0094] Where, p e is the average static pressure of the gas on the nozzle section, p is the atmospheric static pressure at the height of the test bench, S e is the nozzle cross-sectional area. is the unit vector in the direction of the rocket's longitudinal axis.

[0095] The total thrust P of the rocking engine is defined as the relative force and static thrust P st The sum of

[0096]

[0097] Corresponding to the static thrust, relative force Also called kinetic thrust or kinetic thrust. Note that the gas velocity u e point to In the opposite direction, the thrust is

[0098]

[0099] From the above formula, we can see that the engine thrust is not only related to the propellant consumption per second Average exhaust gas velocity u at the nozzle cross section e and its gas static pressure p eIt is related to the external atmospheric pressure P. Since the static atmospheric pressure P decreases continuously with increasing altitude H, it is obvious that the engine thrust increases with increasing flight altitude H. In other words, the engine thrust P is a function of the rocket's flight altitude H. The law of thrust variation with altitude is called the engine thrust-altitude characteristic.

[0100] During the ground test, since the atmospheric pressure p is equal to the ground standard atmospheric pressure p0 and its value is the largest, the engine thrust value is the smallest, that is,

[0101]

[0102] Under vacuum conditions, since the atmospheric pressure is zero, the engine thrust is maximum, that is,

[0103]

[0104] The above thrust calculation is often used in the preliminary ballistic design stage. The actual engine's fuel consumption and thrust are not constant. Figure 6 (b) shows a schematic diagram of the thrust change of the liquid engine, where the AB section is the startup section, the BC section is the stable working section, and the CD section is the shutdown section. The thrust and consumption per second in the AB section increase rapidly. Due to the unstable output of the engine, the rocket does not take off at the moment of ignition (fixed to the ground by a mechanical device). After the thrust stabilizes (enters point A), the thrust and consumption per second in the take-off section are approximately constant. The consumption per second and thrust drop rapidly to zero. The reason for the existence of this stage is that the shutdown process cannot be completed instantaneously. Based on this, the engine thrust model established in this study consists of three sections: Figure 6 (a): It includes three stages: startup stage (increasing from zero parabola to the size of the formula), stabilization stage (given by the formula), and shutdown stage (decreasing from the parabola of the formula to zero).

[0105] Step 2.2: Modeling the swing engine control system:

[0106] The reason why a rocket can fly and accurately make its payload hit the target or enter orbit is entirely the result of the action of engine thrust, control force and control torque. During the flight, the rocket is affected by the earth's gravity, aerodynamic force and engine thrust. Since the line of action of the earth's gravity passes through the center of mass of the rocket, and its size cannot be changed at will, it is impossible to generate a control torque on the center of mass of the rocket. Obviously, the force and torque to control the flight of the rocket can only be generated by changing the direction of aerodynamic force or engine thrust. The force and torque that control the flight of the rocket are respectively called control force and control torque. For launch vehicles, the mechanisms currently used to generate control force and control torque are in the following forms: air rudder, gas rudder, swing engine, swing nozzle and secondary injection and corresponding servo mechanism, and one of them can be used alone according to requirements, or two different forms can be used at the same time. The present invention adopts the swing engine control system modeling ( Figure 7 (a)), here we only discuss the control principle of the “×”-shaped layout swing engine control form.

[0107] The "X"-shaped engine group is composed of four swing engines connected in parallel. The swing engines are equivalent to being installed on the bottom edges of the four sides of a regular tetrahedron. The installation axis is equivalent to the center line of the bottom edge of the cone side, and forms an installation angle μ with the longitudinal axis of the rocket body, such as Figure 7 (b) The engine can swing inside the cone, and the positive and negative swing angles δ are defined as follows: when looking forward from the tail of the rocket, when the engine nozzle swings clockwise along the circumference of the rocket body, δ is defined as positive, and vice versa.

[0108] Obviously, when the engine is not swinging, the thrust line direction of each engine is along the direction of its own installation axis, and the direction of its combined thrust must be along the rocket body axis O1X1. This force can only make the rocket's center of mass move but cannot produce rotation around the center of mass. To make the rocket rotate, the engine must be swung.

[0109] If the four engines swing at a positive δ angle at the same time, then the thrust line of each engine and its installation axis must also form a positive δ angle. i (i=I, II, III, IV) is decomposed into the component P′ along its installation axis i and the component P″ perpendicular to the installation axis i , whose component P′ i The resultant quantity can make the center of mass of the rocket move, and the component P" perpendicular to the installation axis i This can cause the rocket to produce a negative roll motion around its longitudinal axis; conversely, if the engine swings a negative δ angle at the same time, the rocket will inevitably produce a positive roll motion around its longitudinal axis.

[0110] Similarly, when the engines forming the "X" shape swing at the same time according to a certain rule, the rocket will also produce corresponding pitch and yaw motions at the same time. δ ψ , δ γ and the engine swing angle δ I , δ II , δ III , δ IV There is the following relationship between them:

[0111]

[0112] Since the rolling motion of the rocket is very small when it is flying under the control system, the equivalent swing angle δ γ is also very small. When we approximately assume that δ γ When , the above formula can be simplified to

[0113]

[0114] The engine thrust after swinging is projected onto the axes of the rocket body coordinate system. Obviously, the resultant thrust component along the O1X1 axis can only push the rocket's center of mass to move, but cannot make it rotate around the center of mass, so it is called the engine effective thrust; and the resultant thrust components along the O1Y1 and O1Z1 axes can change the direction of the rocket's velocity and control the rocket's pitch and yaw motion around the center of mass, so they are called normal control force and lateral control force respectively.

[0115] Now let’s take the third engine as an example to derive the effective thrust and control force. III When the thrust P III Component P' on its installation axis III and the component P′ perpendicular to the installation axis III Can be expressed as

[0116] P′ III =P III cosδ III ,P″ III =P III sinδ III (twenty four)

[0117] Therefore, P′ III and P′ III The components on each axis of the rocket body coordinate system are

[0118]

[0119] In the same way, the thrust of the oscillating engines I, II, and IV can also be projected onto the axes of the rocket body coordinate system. In this way, the thrust components projected onto the O1X1 axis are superimposed to obtain

[0120] P x1 =(P I cosδ I +P II cosδ II +P III cosδ III +P IV cosδ IV ) (26)

[0121] Assume that the four engines are of the same model, so

[0122]

[0123] Where P is the sum of the thrust of the four engines.

[0124]

[0125] Therefore, the effective thrust of the engine is

[0126]

[0127] The sum of the engine thrust projections on the O1Y1 and O1Z1 axes is

[0128]

[0129] In the formula is the rocking engine control force gradient, P y1 、P z1 They are the normal control force and the lateral control force respectively.

[0130] In summary, the thrust of the swing engine can be expressed in the body coordinate system as

[0131]

[0132] like δ ψ are all small quantities, the above formula can be simplified to

[0133]

[0134] Then the thrust of the swing engine can be expressed in the launch coordinate system as

[0135]

[0136] In the formula is the transformation matrix from the rocket body coordinate system to the launch coordinate system, which is given by the coordinate transformation model library.

[0137] Step 3: Construct the constraint device model between rigid bodies, specifically:

[0138] Before separation, there is no relative motion or rotation between the first and second sub-stages. The constraint equation is described as

[0139]

[0140] Calculate the first and second order derivatives of it with respect to time

[0141]

[0142] This results in a constraint device model based on the Modelica unified modeling language, such as Figure 4 shown.

[0143] Step 4: Based on the rigid body model, functional unit and constraint device model, a rapid simulation model of the launch vehicle's full-process dynamics is quickly built based on the Modelica unified modeling language by building blocks and connecting lines. Figure 2 shown.

[0144] Step 5. Configure the parameters of the full-process dynamic simulation model, record key data such as the rocket's position, speed, acceleration, attitude, and engine working status during the simulation process, and visualize the processed data, such as drawing the rocket's flight trajectory, speed curve, acceleration curve, attitude angle curve, etc., so as to intuitively analyze the rocket's flight performance.

[0145] Example

[0146] In order to verify the effectiveness of the method of the present invention, the following simulation is performed and the results are analyzed.

[0147] Figure 8 .(a)(b)(c)(d) show the offset curves of the center of mass of the first and second stages of the carrier rocket in the direction of the rocket axis and the curves of the change in moment of inertia. It can be seen from the two figures that before the separation of the first and second stages, the center of mass position and moment of inertia of the first stage changed due to the consumption of the first stage propellant. After separation, the second stage propellant is consumed, and the center of mass position and moment of inertia of the second stage change. Since the second stage loses the effect of the after-effect thrust, it is only subjected to a smaller forward thrust, and the acceleration decreases. The remaining figures show the thrust curves of the carrier rocket before and after separation, the changes in the aerodynamic force during flight, and the attitude angle and the initial flight program angle. These results are consistent with theoretical calculations, indicating the correctness of the carrier rocket full-process dynamics rapid simulation model established by the present invention.

[0148] The technical features of the above embodiments can be combined arbitrarily. To make the description concise, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0149] The above-described embodiments merely represent several implementation methods of the present application. While the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the present invention. It should be noted that a person skilled in the art could make various modifications and improvements without departing from the spirit of the present application, all of which fall within the scope of protection of the present application. Therefore, the scope of protection of the present patent application shall be determined by the appended claims.

[0150] Some parts not described in the present invention belong to the known technologies in the art.

Claims

1. A fast simulation method for the whole process dynamics of a launch vehicle based on Modelica language, characterized by: The following steps are involved: Step 1: First, construct the parametric rigid body models of the first and second stages of the launch vehicle with time-varying parameters. The mass, center of mass, and moment of inertia of the first and second stages of the launch vehicle are dynamically changed over time. Then, for the satellite and fairing with time-invariant parameters, respectively, the parametric rigid body models of the satellite and fairing are established. Step 2: Based on the selected rocket engine model and the corresponding guidance, navigation and control system GNC, the functional units of the launch vehicle's first and second stage engine models, the swing engine control system model, the geophysical model, the timing model, the atmospheric parameter model, the launch point information model and the aerodynamic model are established through the object-oriented modeling method. The two functional units of the geophysical model and the launch point information model are used to configure a gravity model with optional gravity field types. The gravity field types can be divided into zero gravity field, uniform gravity field and point gravity field. The point gravity field type is the earth's gravity field considering the J2 item. Finally, the force model acting during the first and second stage separation, fairing separation and satellite-rocket separation is established. Step 3: Connect the first and second stage parameterized rigid body models of the launch vehicle's time-varying parameters, the satellite rigid body model, and the fairing rigid body model through connectors in the Modelica standard library. Based on the connectors, constrain the motion parameters such as displacement, force, and torque to construct a constraint device model. This constraint device model can be reused during the first and second stage separation, fairing separation, and satellite-rocket separation after calling different separation signals. Step 4. Based on the above rigid body model, functional unit and constraint device model, use the building block and connection method to close the input and output parameters of each module and quickly build a full-process dynamic simulation model of the launch vehicle.

2. The method for rapid simulation of the entire launch vehicle dynamics process based on the Modelica language according to claim 1, characterized in that: The parametric rigid body models of the first and second sub-stages of the launch vehicle with time-varying parameters described in step 1 include the parameterized design parameters of time-varying center of mass, time-varying mass, time-varying moment of inertia, length, and diameter, and the output parameters are speed, position, acceleration, and attitude angle. Subsequently, the satellite and fairing rigid body models are parameterized with the time-invariant center of mass, time-invariant mass, time-invariant moment of inertia, length, and diameter as design parameters, and the output parameters are speed, displacement, and acceleration.

3. The method for rapid simulation of the whole process dynamics of a launch vehicle based on Modelica language according to claim 1, characterized in that: The functional units in step 2 include: a launch vehicle first and second stage engine model, a swing engine control system model, an aerodynamic model, an atmospheric parameter model, a geophysical model, a launch point information model, and a timing model; The functional unit is constructed by object-oriented modeling method; the input parameters of the first and second stage engine model of the carrier rocket are the atmospheric pressure at the current altitude, the engine swing angle and the rocket attitude angle, the output parameter is the thrust vector, and the design parameters are the engine second consumption, jet speed, and nozzle area; the input parameters of the swing engine control system model are the attitude angles of the first and second stages, and the output parameters are the swing angles of the first and second stage engines of the carrier rocket, and the design parameters are the static amplification coefficients of the pitch, yaw, and roll channels; the timing model includes the design parameters of the first and second stage engine power on and off time, the engine full start time, the first and second stage separation, the fairing separation, and the satellite-rocket separation time. The first and second stage engines and the constraint device of the carrier rocket are called through the Constant function to complete the control of the flight progress by the timing model; the input parameters of the atmospheric parameter model are the rigid body position vectors of the first and second stage of the carrier rocket, and the output parameters are the atmospheric pressure at the current altitude, the atmospheric density, and the atmospheric pressure at the current altitude. degree and atmospheric temperature, and the design parameters are the atmospheric pressure and density at zero sea level; the input parameters of the aerodynamic model are the atmospheric density at the current altitude, the velocity vector of the first stage, the direction vector of the rocket axis, and the attitude angle of the first stage; the output parameters are the aerodynamic vector in the rocket body coordinate system, and the design parameters are the aerodynamic reference length and area of ​​the rocket; in the gravity model configured by the two functional units of the geophysical model and the launch point information model, the design parameters in the geophysical model and the launch point information model are used as Parameter parameters to participate in the particle gravity field calculation, so that the gravity can be automatically calculated by dragging any rigid body model into the full-process simulation model containing the gravity model; the force model acting during the separation of the first and second stages, the separation of the fairing, and the separation of the satellite and rocket includes the reverse thrust on the first stage rigid body and the forward thrust on the second stage rigid body during the separation of the first and second stages, the separation force applied to the fairing during separation, and the separation force applied to the satellite during separation of the satellite and rocket.

4. The method for rapid simulation of the entire launch vehicle dynamics process based on the Modelica language according to claim 1, characterized in that: The constraint device model in step 3 consists of a constraint device, a signal generator, and a signal source. The two ends of the constraint device are the points of action of the physical frame. Before separation, there is no relative motion or rotation between the first and second sub-stages, the satellite, and the fairing. The constraint equation is described as Where r A 、r B is the position vector of the entity frame interface at both ends of the rigid body to be connected, θ A ,θ B For the attitude angles at both ends of the rigid body to be connected, calculate the first and second order derivatives of the position vector and attitude angle with respect to time; This is used to constrain the relative position, velocity and attitude of the first and second sub-stages, satellites and fairing rigid bodies, and to enable the constraint device to release the constraint when receiving the signal from the timing model.

5. The method for rapid simulation of the whole process dynamics of a launch vehicle based on Modelica language according to claim 1, characterized in that: The rigid body models described in step 3 are connected by connectors, wherein the connectors include flow variables and potential variables, and the sum of the flow variables is zero while the potential variables are equal; A mathematical model is established based on the physical principles, force balance equations, and torque balance equations of the full-process dynamics rapid simulation model of the launch vehicle; a Modelica model is adopted, and code development and interface definition are performed based on a standard library; wherein the interface is used to transmit force, torque, and attitude angle.

6. The method for rapid simulation of the whole process dynamics of a launch vehicle based on Modelica language according to claim 1, characterized in that: The full-process dynamics simulation model of the launch vehicle described in step 4 includes at least the following sub-models: Parametric model of the time-varying parameter rigid body structure of the first and second sub-stages, parametric sub-model of the satellite fairing rigid body structure, timing model, constraint device model, first and second stage engine model of the launch vehicle, swing engine control system model, atmospheric parameter model, aerodynamic model, and force model of each separation stage.

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