Carrier rocket whole-process dynamics rapid simulation modeling method based on Modelica unified modeling language

Abstract

The invention provides a Modelica language-based rapid simulation method for whole-process dynamics of a carrier rocket, and aims to solve the problem of insufficient model reusability and adaptability in a traditional simulation method. By utilizing the object-oriented characteristic of the Modelica language, the method constructs a comprehensive dynamic model, and the model integrates a plurality of subsystems such as a propulsion system, a control system and an aerodynamic model. The method comprises the following specific steps: firstly, constructing an overall model of a carrier rocket, and defining time-varying parameters to accurately calculate key parameters such as propellant consumption, rotational inertia and centroid position; secondly, the influence of different design parameters on the rocket performance is analyzed through simulation, so that potential problems are recognized and solved in advance; and moreover, the simulation result can be analyzed, so that the accuracy and reliability of the model are ensured. The innovation of the invention lies in the modular design and multi-field collaborative simulation capability, which not only significantly improves the reusability and adaptability of the model, but also effectively supports the design optimization, task planning and control strategy formulation of the rocket. The method not only reduces the research and development cost, but also accelerates the progress of the spaceflight technology, and has a wide application prospect.

Description

Technical Field

[0001] The 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 space transportation vehicle composed of multiple stages. Its purpose is 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 key links in aerospace engineering. Dynamics simulation can help engineers analyze the performance of rockets under different design parameters during the design phase, thereby optimizing the design and ensuring flight safety and reliability.

[0003] Although there are relatively complete mathematical models for rocket dynamics simulation, in practical applications, it is still a time-consuming and laborious task to establish simulation models for specific problems. There are significant differences between different models and different applications. Traditional software design methods are difficult to provide a high degree of abstraction, resulting in low model reusability and inability to flexibly adapt to new situations.

[0004] The full-process simulation process involves multiple disciplines such as structural dynamics, multi-body dynamics, attitude control, aerodynamics, etc., which are related to the launch vehicle layout and shape selection, parameter design, etc., and requires the combination of multiple disciplines to conduct multi-physics field coupling simulation. It should be understood that the Modelica language is a language specially designed for unified modeling and simulation in multiple fields, and has unique advantages in multi-physics field coupling simulation. Because the Modelica language has the characteristics of object-oriented, equation-based, model reusable and hierarchical structure, it can effectively realize the full-process system modeling of the launch vehicle. Therefore, there is an urgent need for a solution that can build a full-process simulation model of the launch vehicle based on the Modelica unified modeling language. Summary of the invention

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

[0006] The technical solution to achieve the purpose of the present invention is: a method for rapid simulation of the whole 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, and the dynamic changes of the mass, center of mass and moment of inertia of the first and second stages of the launch vehicle with time; then, for the satellite and fairing with time-invariant parameters, respectively establish the parametric rigid body models of the satellite and fairing;

[0008] Step 2: According to the selected rocket engine model and the corresponding guidance, navigation and control system GNC, the first and second stage engine models of the launch vehicle, 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 the gravity model of the optional gravity field type. The gravity field type can be divided into no 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: The first and second stage parameterized rigid body models of the launch vehicle with time-varying parameters, the satellite rigid body model, and the fairing rigid body model are connected through connectors in the Modelica standard library. Based on the connectors, the motion parameters such as displacement, force, and torque are constrained to construct a constraint device model. This constraint device model can be reused in 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 method of building blocks and connecting lines to close the input and output parameters of each module and quickly build a full-process dynamics simulation model of the launch vehicle.

[0011] Furthermore, the parameterized rigid body models of the first and second sub-stages of the launch vehicle with time-varying parameters described in step one include parameterized design parameters such as 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, parameterized rigid body models of the satellite and fairing are established 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 launch vehicle are the current altitude atmospheric pressure, engine swing angle and 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 first and second stage attitude angles, the output parameter is the first and second stage engine swing angles of the launch vehicle, 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 restraint device of the launch vehicle are called by 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 first and second stage rigid body position vectors of the launch vehicle, and the output parameters are the current altitude atmospheric pressure, atmospheric density degree and atmospheric temperature, and the design parameters are zero sea level atmospheric pressure and density; 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 the rocket includes the reverse thrust on the first stage rigid body and the forward thrust on the second stage rigid body when the first and second stages are separated, the separation force applied to the fairing when the fairing is separated, and the separation force applied to the satellite when the satellite and the rocket are separated.

[0014] Furthermore, the constraint device model described in step 3 is composed of a constraint device, a signal generator, and a signal generator. The two ends of the constraint device are the action points 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

[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, speed and attitude of the first and second sub-stages, satellites and fairing rigid bodies, and enables 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 by a connector, wherein the connector includes flow variables and potential variables, and the sum of the flow variables is zero and 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 in step 4 includes at least the following sub-models:

[0022] Parameterized model of the rigid body structure with time-varying parameters of the first and second stages, parameterized sub-model of the rigid body structure of the satellite fairing, 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 acting at each separation stage.

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

[0024] 1) Based on the Modelica language, the model library of the developed launch vehicle model can have a clear model hierarchy, be reusable and scalable, and greatly improve the modeling efficiency;

[0025] 2) It can perform rapid numerical simulation of the whole 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 invention discloses a flow chart of a launch vehicle full-process dynamics simulation modeling method based on the Modelica unified modeling language.

[0027] Figure 2 The invention discloses 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 restraint device model based on 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 model 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 launch vehicle during the flight of the launch vehicle.

[0031] Figure 6 It 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 of the present invention; 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 It is a gravity field model based on Modelica language and a component parameter interface of the model with selectable gravity field types; 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 It is a large launch vehicle simulation result diagram drawn by the launch vehicle full process dynamics fast simulation model based on 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 change curve of the moment of inertia 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 invention discloses a method for rapid simulation of full-process dynamics of a launch vehicle based on Modelica language, comprising the following steps: step 1, rigid body modeling of a satellite, a fairing, a first and a second stage of a launch vehicle, a variable mass center, and a moment of inertia; step 2, modeling of functional units such as a first and a second stage engine, a control system, an aerodynamic module, and a reverse thrust rocket; step 3, constructing a constraint device model between the first and second stages, the satellite, and the fairing rigid bodies; step 4, based on the above rigid body model, the functional unit and the constraint device model, a full-process dynamics simulation model of the launch vehicle is rapidly constructed by building blocks and connecting lines;

[0035] Through dynamic simulation, the performance of the rocket under different design parameters is analyzed, and the simulation results are analyzed theoretically to verify the accuracy of the model, ensuring that it can accurately predict the rocket's flight trajectory, speed and attitude changes, increase the flexibility and scalability of the model, shorten the launch vehicle's R&D cycle and reduce costs. Specifically, the parameters of the full-process dynamic simulation model are configured, and key data such as the rocket's position, speed, acceleration, attitude, and engine working status during the simulation process are recorded, and the processed data is visualized, such as drawing the rocket's flight trajectory diagram, speed curve, acceleration curve, attitude angle curve, etc., so as to intuitively analyze the rocket's flight performance.

[0036] In the present invention, the functional units include: engine thrust acting on the launch vehicle, swing engine control system model for changing thrust vector, aerodynamic model, and atmospheric parameter model. The full-process simulation model of the launch vehicle includes at least the following sub-models: time-varying parameter first and second sub-stage rigid body structure parameterized sub-models, satellite rigid body model, fairing rigid body model, timing model, launch vehicle first and second stage engine model, swing engine control system model, atmospheric parameter model, aerodynamic model, and reverse thrust rocket model.

[0037] Specifically, the time-varying parameter parameterized sub-model of the rigid body structure of the first and second 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 stages; the timing model includes the start and shutdown moments of the first and second stage engines, the moment 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 second consumption, jet speed, jet pressure, nozzle area and other parameters; 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, and atmospheric density; 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 stages, satellites, and fairing rigid bodies. It is used to constrain the relative position, velocity, and attitude of the first and second stages, satellites, and fairing rigid bodies, and enables 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 according to the physical principles, force balance equations and torque balance equations of the full-process dynamics simulation model; a Modelica model is used to develop codes and define interfaces based on a standard library, wherein the interface is used to transmit force, torque, height and attitude angle, etc.; the parameterized sub-models are connected by connectors, and the connectors 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 stages, satellites, 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 respectively 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 solution and advantages of the present application more clearly understood, the present application is further described in detail below in conjunction with the accompanying drawings and embodiments. It should be understood that the specific embodiments described herein are only used to explain the present application and are not used to limit the present 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: Rigid body modeling 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 response to the time-varying mass problem, the following uses the first-stage flight segment 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, The 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 mass of the first and second level arrow bodies 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 mass center and the theoretical cusp

[0058]

[0059] In the formula is the distance from the center of mass of the first-stage propellant to 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 tip. is the distance between the mass center of the secondary propellant and the theoretical cusp at the initial moment, D F D 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), s is the distance between the satellite 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 own long axis coincides with O1X1), R O1 , R O2 I′ is the radius of the first and second stage propellant tanks, which is 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 rotational inertia of the satellite relative to its own body coordinate X-axis (assumed to coincide with the long axis of the rocket body).

[0067] The moment of inertia of the primary 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 arrow 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 It 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 is as follows: The moment of inertia of the launch vehicle system around the O1Y1 axis can be expressed as

[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 rotational inertia of the remaining fuel of the first stage and the fuel of the second stage about the O1Z1 axis, I Fz ,I Sz They are the rotational inertia of the two fairings and the satellite about the O1Z1 axis respectively. The derivation process will not be elaborated any more.

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

[0089]

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

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

[0092] During the static ground test of the engine, 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] In the formula, p e is the average static pressure of the gas on the nozzle section, p is the static pressure of the atmosphere 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 dynamic thrust or dynamic thrust. Note that the gas velocity u e point to The thrust is in the opposite direction of

[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. According to the characteristic that the static atmospheric pressure p decreases continuously with the increase of the altitude H, it is obvious that the engine thrust increases with the increase of the flight altitude H, that is, the engine thrust P is a function of the rocket's flight altitude H. The law of thrust changing 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 energy consumption and thrust are not constant. Figure 6 (b) shows a schematic diagram of the thrust change of a liquid engine, where the AB segment is the startup segment, the BC segment is the stable working segment, and the CD segment is the shutdown segment. The thrust and power consumption per second in the AB segment increase rapidly. Due to the unstable engine output, 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 power consumption per second in the take-off segment are approximately constant. The power 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 segments: 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 of the swing engine control system model:

[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 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 called control force and control torque respectively. 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 mechanisms, and one of the forms 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 midline 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 The engine can swing inside the cone, and the positive and negative swing angles δ are defined as follows: 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, otherwise it is negative.

[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 angle δ at the same time, then the thrust line of each engine and its installation axis must also form a positive angle δ. i (i=I, Ⅱ, Ⅲ, Ⅳ) is decomposed into the component P′ along the direction of 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" in the direction perpendicular to the installation axis i This can cause the rocket to produce a negative rolling motion around its longitudinal axis; conversely, if the engine swings at a negative δ angle at the same time, the rocket will inevitably produce a positive rolling 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 inevitably 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 are the following relationships:

[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. 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 direction 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 effective engine thrust; and the resultant thrust component along the O1Y1 and O1Z1 axis directions 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 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 I, II, and IV engines after swinging 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 there are

[0122]

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

[0124]

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

[0126]

[0127] The sum of the projections of the engine thrust 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 normal control force and 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 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, and the constraint equation is described as

[0139]

[0140] Find its first and second order derivatives with respect to time

[0141]

[0142] Thus, the constraint device model based on Modelica unified modeling language is obtained, such as Figure 4 shown.

[0143] Step 4: Based on the above rigid body model, functional unit and constraint device model, a rapid simulation model of the launch vehicle 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 parameters of the full-process dynamics simulation model, record key data such as the rocket's position, speed, acceleration, attitude, engine working status, etc. 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) give the displacement curves of the center of mass of the first and second stages of the launch vehicle in the direction of the rocket axis and the change curves of the 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 the 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 affected by a smaller positive thrust, and the acceleration decreases. The remaining diagrams show the thrust curves of the launch vehicle before and after separation, the changes in the aerodynamic force during the flight, and the attitude angle and the initial flight program angle. These results are consistent with the theoretical calculations, indicating the correctness of the full-process dynamics rapid simulation model of the launch vehicle established by the present invention.

[0148] The technical features of the above embodiments may 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-mentioned embodiments only express several implementation methods of the present application, and the descriptions thereof are relatively specific and detailed, but they cannot be understood as limiting the scope of the invention patent. It should be pointed out that, for a person of ordinary skill in the art, several variations and improvements can be made without departing from the concept of the present application, and these all belong to the protection scope of the present application. Therefore, the protection scope of the patent of the present application shall be subject to the attached claims.

[0150] Some of the techniques not described in the present invention belong to the known techniques in the art.

Claims

1. A fast simulation method for the whole process dynamics of a launch vehicle based on Modelica language, characterized in that: 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, and the dynamic changes of the mass, center of mass and moment of inertia of the first and second stages of the launch vehicle with time; then, for the satellite and fairing with time-invariant parameters, respectively establish the parametric rigid body models of the satellite and fairing; Step 2: According to the selected rocket engine model and the corresponding guidance, navigation and control system GNC, the first and second stage engine models of the launch vehicle, 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 the gravity model of the optional gravity field type. The gravity field type can be divided into no 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: The first and second stage parameterized rigid body models of the launch vehicle with time-varying parameters, the satellite rigid body model, and the fairing rigid body model are connected through connectors in the Modelica standard library. Based on the connectors, the motion parameters such as displacement, force, and torque are constrained to construct a constraint device model. This constraint device model can be reused in 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 method of building blocks and connecting lines to close the input and output parameters of each module and quickly build a full-process dynamics simulation model of the launch vehicle.

2. 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 first and second sub-stage parameterized rigid body models 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 with time-invariant center of mass, time-invariant mass, time-invariant moment of inertia, length, and diameter as design parameters are parameterized, and the output parameters are speed, displacement, and acceleration.

3. The method for full-process dynamics simulation 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 launch vehicle are the current altitude atmospheric pressure, engine swing angle and 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 first and second stage attitude angles, the output parameter is the first and second stage engine swing angles of the launch vehicle, 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 restraint device of the launch vehicle are called by 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 first and second stage rigid body position vectors of the launch vehicle, and the output parameters are the current altitude atmospheric pressure, atmospheric density degree and atmospheric temperature, and the design parameters are zero sea level atmospheric pressure and density; 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 the rocket includes the reverse thrust on the first stage rigid body and the forward thrust on the second stage rigid body when the first and second stages are separated, the separation force applied to the fairing when the fairing is separated, and the separation force applied to the satellite when the satellite and the rocket are separated.

4. The method for full-process dynamics simulation of a launch vehicle based on Modelica language according to claim 1, characterized in that: Step 3 The constraint device model 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 entity 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, speed and attitude of the first and second sub-stages, satellites and fairing rigid bodies, and enables 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 full-process dynamics simulation 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: Parameterized model of the rigid body structure with time-varying parameters of the first and second stages, parameterized sub-model of the rigid body structure of the satellite fairing, 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 acting at each separation stage.

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