Ballistic design complex process high abstraction and complex model layered decoupling method

Abstract

The invention relates to a trajectory design complex process high abstraction and complex model layering decoupling method, and belongs to the field of aircraft orbit design. Setting a minimum integral unit, and selecting a core modeling object of the carrier rocket according to the minimum integral unit; establishing a thrust equation of the carrier rocket engine; calculating a small deviation thrust in the engine thrust equation; establishing a propellant flow deviation equation according to the small deviation thrust; engine control force Fpk and control torque Mck are calculated; synthesizing equivalent three-channel aerodynamic force and aerodynamic moment of the rocket system; determining a pneumatic modeling tool; generating a required aerodynamic model; carrying out abstraction modeling on a trajectory section and a flight time sequence; carrying out complex program angle abstraction modeling on the multi-section ballistic trajectory; according to the method, the complicated process is highly abstracted, and the complicated technology is refined and decomposed, so that the trajectory design and optimization method which is high in universality and wide in application range is formed.

Description

Technical Field

[0001] The invention belongs to the field of aircraft trajectory design, and relates to a method for highly abstracting a complex process of trajectory design and hierarchical decoupling of a complex model. Background Art

[0002] The design of an aircraft begins with the trajectory and ends with the trajectory. The trajectory design plays a very important role in the overall design of the aircraft. The trajectory design is highly professional and technically difficult. The trajectory design methods of different types of aircraft are quite different, but the design steps are basically composed of complex processes such as simulation basic data configuration, model building, simulation drive configuration, and trajectory design method selection. Therefore, it is very difficult to highly abstract the complex process and refine and decompose the complex technology to form a trajectory design and optimization module with strong versatility and wide adaptability.

[0003] The essence of missile trajectory design is based on flight mechanics and control theory, combined with numerical analysis and optimization methods, to form a model decomposition of strong nonlinear, multi-constrained optimization problem. Modeling abstraction is the core difficulty of general software. Taking the launch vehicle as an example, the aircraft is a complex variable mass flexible body, and its motion characteristics include: multiple forces and torques act simultaneously, including thrust, gravity, aerodynamic force, control force, inertia force and more than ten types of forces and torques; multiple motions exist simultaneously, including center of mass motion, center of mass motion, elastic vibration, shaking, and even propellant flow; elastic deformation is complex, control object is nonlinear, parameters and interference change randomly, multiple coordinate systems exist and are converted to each other, and in some cases, decoupling simplification and conversion are required. For missile trajectory design software, it is necessary to perform functional decomposition of various requirements, form a general model and process abstraction, so that the complex missile trajectory design problem can be generalized through industrial software. Summary of the invention

[0004] The technical problem solved by the present invention is: to overcome the shortcomings of the prior art and propose a method for highly abstracting the complex process of ballistic design and hierarchical decoupling of complex models, so as to realize highly abstracting the complex process and decomposing the complex technology, thereby forming a ballistic design and optimization method with strong versatility and wide adaptability.

[0005] The solution of the present invention is:

[0006] A method for highly abstracting complex processes in trajectory design and decoupling complex models in layers, including:

[0007] Set the minimum integral unit and select the core modeling object of the launch vehicle according to the minimum integral unit;

[0008] Establish the thrust equation for the launch vehicle engine;

[0009] Calculate the small deviation thrust in the engine thrust equation; establish the propellant flow deviation equation based on the small deviation thrust;

[0010] Calculate the engine control force F pk and control torque M ck ;

[0011] Equivalent three-channel aerodynamic force and aerodynamic moment of synthetic arrow system; determine aerodynamic modeling tools; generate required aerodynamic models;

[0012] Abstract modeling of ballistic segments and flight timing;

[0013] Perform complex procedural corner abstract modeling of multi-segment trajectories.

[0014] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model in layers, the minimum integration unit contains basic position, velocity, and posture information; any other transformation of the design model and the addition of the design process are connected to the minimum integration unit; and the flexible interpreter and script are also connected to the minimum integration unit.

[0015] In the above-mentioned method for highly abstracting a complex process of trajectory design and decoupling a complex model in layers, the minimum integral unit satisfies the following equation:

[0016]

[0017]

[0018] Where x is the state vector;

[0019] t is the integration time;

[0020] The superscript · indicates the derivative of a variable with respect to time t;

[0021] is the derivative of position r with respect to time;

[0022] v is the velocity vector, a is the acceleration vector;

[0023] m is the mass, dt is the derivative with respect to time;

[0024] dm is the derivative of mass with respect to time;

[0025] represents the derivative of a variable with respect to time t;

[0026] x, y, and z are three-dimensional positions respectively;

[0027] V is the speed;

[0028] W is the apparent velocity;

[0029] g is the gravitational acceleration;

[0030] The subscripts x, y, and z are the three directions respectively;

[0031] The subscript a indicates component form.

[0032] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model layer by layer, the method for selecting the core modeling objects of the launch vehicle is:

[0033] The method of abstract ballistic segment mounting structure level is adopted; flight timing, mass, aerodynamics, and structure are all abstracted as configuration objects; the mission design takes the ballistic segment as the main line, and various ballistic configuration objects are mounted.

[0034] In the above-mentioned highly abstracted complex process of trajectory design and hierarchical decoupling method of complex models, the thrust equation of the engine rated working section is:

[0035]

[0036] Where, P0 is the rated ground thrust of a single engine;

[0037] S a is the engine nozzle area;

[0038] ΔP i is the small deviation thrust, which is calculated by the engine small deviation equation;

[0039] p0 is the standard ground atmospheric pressure, p is the local atmospheric pressure;

[0040] In the startup and shutdown working stages, the source variable uses the measured value provided by the engine unit, so the startup and shutdown stages ΔP i is zero, the following thrust equation is used to calculate the thrust

[0041]

[0042] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model layer by layer, the small deviation thrust equation is:

[0043] ΔP i =(c 11 ΔT Oz +c 12 ΔT fz +c 13 ΔP Oz +c 14 ΔP fz +c 15 ×(n x1 -1))

[0044] The propellant flow deviation equation is:

[0045]

[0046] Where Δ is the increment, which is the deviation from its standard value during flight;

[0047] ΔP i It is a small deviation thrust;

[0048] is the propellant flow rate per second;

[0049] T is the propellant temperature;

[0050] P is the engine pump inlet pressure;

[0051] n x1 is the overload coefficient along the longitudinal axis of the rocket body;

[0052] Divide by n x1 In addition, the subscripts of each symbol mean:

[0053] The subscripts Z, 1, and 2 denote booster, core first stage, and second stage, respectively;

[0054] The subscripts O and f represent oxidant and combustion agent, respectively;

[0055] c11 to c35 are the coefficients of the small deviation equation and are constants.

[0056] In the above-mentioned highly abstract and complex model hierarchical decoupling method of a complex trajectory design process, the thrust of the engine along the thrust line is assumed to be P; following the design practice, y is defined in the swing coordinate system. s The positive swing angle is δ1, z s The positive swing angle is δ2; the subscript s represents the swing coordinate system component; δ1 is the swing angle around -z s Axis swing; δ2 is around y s Axis swing; projecting the engine thrust onto the engine swing coordinate system requires rotating around z s Axis rotation δ1, around y s The axis rotates -δ2, and the force generated is:

[0057]

[0058] From the engine swing coordinate system to the arrow system, there are three rotations, namely, from the engine swing coordinate system to the engine installation coordinate system; then the installation angle and quadrant angle are rotated to the arrow body coordinate system in turn. The conversion matrix is:

[0059]

[0060] In the formula, δ f Swing the azimuth angle for the engine;

[0061] δA Mounting angles for the engine;

[0062] δ q Mounting angles for the engine;

[0063] The distance from the control force point of the kth engine to the rocket axis is R k ; The distance from the center of mass to the theoretical tip is X z , the distance from the engine to the theoretical tip of the rocket is

[0064] Then the position of the engine control force action point in the rocket body coordinate system is for:

[0065]

[0066] The projection of the kth engine force on the arrow system is:

[0067]

[0068] In the formula, is the transformation matrix from the swing coordinate system to the rocket body coordinate system;

[0069] F is the thrust vector;

[0070] The kth engine force acts on the arrow system with a torque M ck for:

[0071]

[0072] In the formula, is the force arm vector;

[0073] but:

[0074]

[0075] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model in layers, the equivalent three-channel and multi-segment combination is adopted to synthesize the equivalent three-channel aerodynamic force and aerodynamic moment of the arrow system; a tree data structure is used to encapsulate the aerodynamic data as an aerodynamic modeling tool; each channel corresponds to a set of default aerodynamic coefficient name identifiers and corresponding operation rules; the aerodynamic data file defined by the corresponding identifier is directly read and the required aerodynamic model is generated.

[0076] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model in layers, the method of abstracting the trajectory segment and flight sequence modeling is as follows:

[0077] The simulation-driven timing reconstruction module is adapted to the launch vehicle trajectory design;

[0078] Define flight sequence discrete events, allowing custom actions at the time of event occurrence; calculate complex trajectories consisting of several flight segments, and flexibly drive the trajectory in segments by specifying time or conditions;

[0079] The physical flight level is divided based on the characteristic values ​​of various flight parameters; the end of the flight segment is defined according to the characteristic values ​​of time and speed flight parameters.

[0080] In the above-mentioned method of highly abstracting the complex process of trajectory design and decoupling the complex model in layers, the complex program angle abstraction model of the multi-segment trajectory realizes a small-angle model and a full-angle model; realizes independent configuration of 3 channels; realizes arbitrary combination of 6 angle variables of pitch angle / angle of attack, yaw angle / sideslip angle, roll angle / roll angle; supports the splicing of any multiple-segment flight program curves in a single flight segment; and each curve supports 10 different curve forms.

[0081] The beneficial effects of the present invention compared with the prior art are:

[0082] (1) The present invention takes into account the mission requirements of various trajectory types, such as one-time orbital entry, glide segment, suborbital transition orbital entry, pure thrust vector control, main and auxiliary aircraft coordination, air rudder review control, atmospheric gliding trajectory, warhead reentry, return trajectory, etc., and highly abstracts the above trajectory design process to form a unified trajectory design model;

[0083] (2) The present invention considers the research on the integrated overall architecture modeling of the trajectory design process, the high abstraction of the trajectory design process, and the decoupling and decomposition technology of the complex process in a near-real application environment, and constructs the overall architecture model of the initiation, initialization, process monitoring, driving parametric design, calling simulation and saving of the entire design process. It is tested and analyzed in a laboratory environment to meet the requirements of trajectory design optimization and simulation with strong versatility and wide adaptability;

[0084] (3) The present invention establishes a core object architecture system that complements the missile / rocket trajectory segment and the structural level, realizes a single data source for rocket flight variation parameters and rocket characteristic parameters, and adapts to the modeling needs of different types of systems or systems of the same type with different functional requirements. At the same time, the ballistic segment is used as the main line to mount various ballistic configuration objects, further realizing flexible design of the ballistic segment. This technology has been verified by flight tests when the launch vehicle is performing a model flight mission;

[0085] (4) The present invention constructs a universal and standardized ballistic component model construction method under the model real model library and domestic platform environment, making full use of the componentization, modularization and hierarchical characteristics of the configuration components in the abstract ballistic design process, realizing the interface standard specification and rapid switching between multiple models, and unifying the model library and the underlying model encapsulation interface calling form of various aircraft ballistic design specifications. The compatibility test is completed with the basic component library for the final ballistic design of different platform models, and it has been formally applied in the development process of the model product, meeting the corresponding overall design requirements.

[0086] (5) The present invention takes into account the complex flight profile trajectory design function of maneuverable aircraft. By performing multi-segment program angle design, different combinations of the above configurations can be used within a certain section of the trajectory according to needs, which greatly broadens the support for ballistic flight types and can meet the current complex ballistic flight profile design needs of various types of aircraft in the atmosphere in my country.

[0087] (6) The universal modeling of complex timing of the present invention is particularly suitable for large-scale bundled liquid launch vehicles such as CZ-5 and CZ-7, where several engines work simultaneously in the same physical flight stage, but the physical shutdown timing of each engine is different. The above-mentioned flight segment modeling method can be flexibly designed to fully meet the ballistic design requirements of such complex flight timing. BRIEF DESCRIPTION OF THE DRAWINGS

[0088] Figure 1 The present invention provides a highly abstracted flow chart of the complex trajectory design process and a hierarchical decoupling flow chart of the complex model. DETAILED DESCRIPTION

[0089] The present invention will be further described below in conjunction with the embodiments.

[0090] The present invention provides a highly abstract and complex model layered decoupling technology for complex processes in trajectory design. It adopts the abstract trajectory segment mounting structure level method to abstract the flight sequence, mass, aerodynamics, and structure into configuration objects. The mission design is based on the trajectory segment, and each trajectory configuration object is mounted to ensure that the data source is unique, so as to realize the flexible design of the trajectory segment. The core technologies include: the core modeling object selection method of the launch vehicle, the engine thrust parameterized modeling, the general high-dimensional complex aerodynamic parameterized modeling, the flight sequence abstract modeling, and the trajectory design process abstraction method.

[0091] Highly abstract the complex process of trajectory design and hierarchical decoupling methods of complex models, such as Figure 1 As shown, the specific steps include:

[0092] Set the minimum integral unit and select the core modeling object of the launch vehicle based on the minimum integral unit.

[0093] The minimum integration unit contains basic position, speed, and posture information; any other transformations of the design model and additions to the design process are connected to the minimum integration unit; flexible interpreters and scripts are also connected to the minimum integration unit.

[0094] The minimum integration unit satisfies the following equation:

[0095]

[0096]

[0097] Where x is the state vector;

[0098] t is the integration time;

[0099] The superscript · indicates the derivative of a variable with respect to time t;

[0100] is the derivative of position r with respect to time;

[0101] v is the velocity vector, a is the acceleration vector;

[0102] m is the mass, dt is the derivative with respect to time;

[0103] dm is the derivative of mass with respect to time;

[0104] represents the derivative of a variable with respect to time t;

[0105] x, y, and z are three-dimensional positions respectively;

[0106] V is the speed;

[0107] W is the apparent velocity;

[0108] g is the gravitational acceleration;

[0109] The subscripts x, y, and z are the three directions respectively;

[0110] The subscript a indicates component form.

[0111] The method for selecting the core modeling objects of the launch vehicle is:

[0112] The method of abstract ballistic segment mounting structure level is adopted; flight timing, mass, aerodynamics, and structure are all abstracted as configuration objects; the mission design takes the ballistic segment as the main line, and various ballistic configuration objects are mounted.

[0113] Establish the thrust equation for the launch vehicle engine.

[0114] The thrust equation of the engine in rated working section is:

[0115]

[0116] Where, P0 is the rated ground thrust of a single engine;

[0117] S a is the engine nozzle area;

[0118] ΔP i is the small deviation thrust, which is calculated by the engine small deviation equation;

[0119] p0 is the standard ground atmospheric pressure, p is the local atmospheric pressure;

[0120] In the startup and shutdown working stages, the source variable uses the measured value provided by the engine unit, so the startup and shutdown stages ΔP i is zero, the following thrust equation is used to calculate the thrust

[0121]

[0122] Calculate the small deviation thrust in the engine thrust equation; establish the propellant flow deviation equation based on the small deviation thrust.

[0123] The small deviation thrust equation is:

[0124] ΔP i =(c 11 ΔT Oz +c 12 ΔT fz +c 13 ΔP Oz +c 14 ΔP fz +c 15 ×(n x1 -1))

[0125] The propellant flow deviation equation is:

[0126]

[0127] Where Δ is the increment, which is the deviation from its standard value during flight;

[0128] ΔP i It is a small deviation thrust;

[0129] is the propellant flow rate per second;

[0130] T is the propellant temperature;

[0131] P is the engine pump inlet pressure;

[0132] n x1 is the overload coefficient along the longitudinal axis of the rocket body;

[0133] Divide by n x1In addition, the subscripts of each symbol mean:

[0134] The subscripts Z, 1, and 2 denote booster, core first stage, and second stage, respectively;

[0135] The subscripts O and f represent oxidant and combustion agent, respectively;

[0136] c11 to c35 are the coefficients of the small deviation equation and are constants.

[0137] Calculate the engine control force F pk and control torque M ck .

[0138] Assume that the thrust of the engine along the thrust line is P; follow the design practice and define y in the swing coordinate system s The positive swing angle is δ1, z s The positive swing angle is δ2; the subscript s represents the swing coordinate system component; δ1 is the swing angle around -z s Axis swing; δ2 is around y s Axis swing; projecting the engine thrust onto the engine swing coordinate system requires rotating around z s Axis rotation δ1, around y s The axis rotates -δ2, and the force generated is:

[0139]

[0140] From the engine swing coordinate system to the arrow system, there are three rotations, namely, from the engine swing coordinate system to the engine installation coordinate system; then the installation angle and quadrant angle are rotated to the arrow body coordinate system in turn. The conversion matrix is:

[0141]

[0142] In the formula, δ f Swing the azimuth angle for the engine;

[0143] δ A Mounting angles for the engine;

[0144] δ q Mounting angles for the engine;

[0145] The distance from the control force point of the kth engine to the rocket axis is R k ; The distance from the center of mass to the theoretical tip is X z , the distance from the engine to the theoretical tip of the rocket is

[0146] Then the position of the engine control force action point in the rocket body coordinate system is for:

[0147]

[0148] The projection of the kth engine force on the arrow system is:

[0149]

[0150] In the formula, is the transformation matrix from the swing coordinate system to the rocket body coordinate system;

[0151] F is the thrust vector;

[0152] The kth engine force acts on the arrow system with a torque M ck for:

[0153]

[0154] In the formula, is the force arm vector;

[0155] but:

[0156]

[0157] Synthesize the equivalent three-channel aerodynamic force and aerodynamic moment of the arrow system; determine the aerodynamic modeling tool; generate the required aerodynamic model. The specific method is:

[0158] The equivalent three-channel aerodynamic force and aerodynamic moment of the arrow system are synthesized by combining the equivalent three-channel and multiple sections. The aerodynamic data is encapsulated in a tree data structure as an aerodynamic modeling tool. Each channel corresponds to a set of default aerodynamic coefficient name identifiers and corresponding operation rules. The aerodynamic data file defined by the corresponding identifier is directly read and the required aerodynamic model is generated.

[0159] Abstract modeling of trajectory segments and flight timing.

[0160] The method for abstract modeling of trajectory segments and flight timing is:

[0161] The simulation-driven timing reconstruction module is adapted to the launch vehicle trajectory design;

[0162] Define flight sequence discrete events, allowing custom actions at the time of event occurrence; calculate complex trajectories consisting of several flight segments, and flexibly drive the trajectory in segments by specifying time or conditions;

[0163] The physical flight level is divided based on the characteristic values ​​of various flight parameters; the end of the flight segment is defined according to the characteristic values ​​of time and speed flight parameters.

[0164] Perform complex procedural corner abstract modeling of multi-segment trajectories.

[0165] The complex program angle abstraction model of multi-segment trajectory realizes small-angle model and full-angle model; realizes independent configuration of 3 channels; realizes arbitrary combination of 6 angle variables of pitch angle / attack angle, yaw angle / sideslip angle, roll angle / roll angle; supports any splicing of multiple flight program curves in a single flight segment; each curve supports 10 different curve forms.

[0166] The more detailed implementation steps are:

[0167] Step 1: Method for selecting core modeling objects of launch vehicles

[0168] Aiming at the problem of selecting core modeling objects, a method for selecting core modeling objects for launch vehicles is proposed. This method is used to realize universal dynamic modeling of launch vehicles, realize universal description of flight timing, structure, mass, aerodynamics, etc. in any multi-segment and multi-module, and realize universal and flexible modeling configuration through unique design source.

[0169] The minimum integral unit is defined as containing basic position, velocity, and posture information. Any other transformation of the design model and addition of the design process are connected to the minimum integral unit, including flexible interpreters and scripts that are also accessed outside the unit.

[0170]

[0171] The above is the expression of the minimum integration unit, where x is the state vector, t is the integration time, and the superscript · represents the derivative of a variable with respect to time t. is the time derivative of position r, v is the velocity vector, a is the acceleration vector, m is the mass, dt is the time derivative, and dm is the time derivative of mass. It represents the derivative of a variable with respect to time t. x, y, z are the three-direction positions, V is the velocity, W is the apparent velocity, g is the gravitational acceleration, the subscripts x, y, z are the three directions, and the subscript a represents the component form.

[0172] Based on the minimum integral unit, the next step involves the core object selection method and modeling issues. Therefore, a core object selection method for general modeling is proposed. The flight sequence, mass, aerodynamics, and structure are abstracted as configuration objects by using the abstract ballistic segment mounting structure level. The mission design is based on the ballistic segment and various ballistic configuration objects are mounted to achieve the unity of the two, which ensures the uniqueness of the data source and realizes the flexible design of the ballistic segment.

[0173] Step 2: Engine thrust parameter modeling

[0174] Liquid rocket engines and solid rocket engines ultimately act in the dynamic equation in the form of force, so it is necessary to parametrically model the engine thrust. The process of parametric modeling is mainly to abstract the calculation process of the control force and control torque generated by a single engine into a general mathematical model, and based on the abstracted mathematical model. It includes the following aspects:

[0175] 1. Engine thrust calculation

[0176] The following thrust equation is used in each engine rated operating section (after the start-up process and before shutdown):

[0177]

[0178] In the above formula: P0 is the rated ground thrust of a single engine; S a is the engine nozzle area; ΔP i is the small deviation thrust, calculated by the engine small deviation equation; p0 is the standard ground atmospheric pressure, and p is the local atmospheric pressure.

[0179] In the startup and shutdown working stages, the source variable uses the measured value provided by the engine unit, so the startup and shutdown stages ΔP i is zero, the following thrust equation is used to calculate the thrust.

[0180]

[0181] 2. Engine small deviation equation

[0182] The basic form of the engine small deviation equation is to obtain the changes in engine thrust and flow rate according to the changes in propellant temperature and engine pump inlet pressure during flight. The impact of thrust changes with atmospheric pressure during flight needs to be considered additionally.

[0183] Taking a certain liquid oxygen-kerosene equation as an example, its small deviation equation form is:

[0184] ΔP i =(c 11 ΔT Oz +c 12 ΔT fz +c 13 ΔP Oz +c 14 ΔP fz +c 15 ×(n x1 -1))

[0185] Similar to thrust, the propellant flow deviation equation is:

[0186]

[0187] Where:

[0188] The meaning of symbols in the small deviation equations in this section is:

[0189] Δ: represents increment, the deviation from its standard value during flight;

[0190] Pi: represents engine thrust;

[0191] Propellant flow rate per second, (kg / s);

[0192] T: propellant temperature, (K);

[0193] P: engine pump inlet pressure, (MPa);

[0194] n x1 : Overload coefficient along the longitudinal axis of the rocket body, dimensionless.

[0195] Divide by n x1 In addition, the subscripts of each symbol mean:

[0196] Z, 1, 2: subscripts represent booster, core first stage, and second stage respectively;

[0197] O, f: subscripts represent oxidant and combustion agent respectively;

[0198] c11 to c35 are the coefficients of the small deviation equation and are constants.

[0199] 3. Control force and torque calculation

[0200] Assume that the thrust of the engine along the thrust line is P, and follow the design convention and define it in the swing coordinate system (the subscript s represents the swing coordinate system component, the same below) y s The positive swing angle is δ1 (around -z s Swing), z s The positive swing angle is δ2 (around y s To project the engine thrust onto the engine swing coordinate system, we need to rotate the z axis. s Axis rotation δ1, around y s The axis rotates -δ2. The generated force is:

[0201]

[0202] From the engine swing coordinate system to the arrow system, there are three rotations, namely, from the engine swing coordinate system to the engine installation coordinate system; then the installation angle and quadrant angle are rotated to the arrow body coordinate system in turn. The conversion matrix is:

[0203]

[0204] The engine swing azimuth is defined as δ f, the engine mounting angle is defined as δ A , the engine installation angle is defined as. Let the distance from the control force point of the Kth engine (subscript k) to the rocket axis be R k The distance from the center of mass to the theoretical tip is X z , the distance from the engine to the theoretical tip of the rocket is Then the position of the engine control force action point in the rocket body coordinate system is:

[0205] The projection of the force of the Kth engine on the arrow system is

[0206]

[0207] in, is the transformation matrix from the rocking coordinate system to the rocket body coordinate system. F represents the thrust vector.

[0208] The torque of the Kth engine acting on the arrow system is Mck. The expression is as follows:

[0209]

[0210] The force arm vector is defined as The expression is as follows:

[0211]

[0212] roll out

[0213]

[0214] Step 3: General high-dimensional complex aerodynamic parameter modeling

[0215] In view of the difficulty of general high-dimensional complex aerodynamic parameter modeling, 1) the types are complex. As a multi-stage aircraft, the aerodynamic shape of each stage of different models changes during the flight. The aerodynamic data obtained through wind blowing tests and numerical calculations are in various formats, including speed system, missile system, and combination; 2) the aerodynamic combination form is flexible (there are various forms such as head rudder, tail rudder, grid rudder, wing, etc., both local aerodynamic and overall aerodynamic) 3) the aerodynamic data type is complex (in the scheme development stage, aerodynamic data is limited, and it is necessary to support aerodynamic data expansion to complete the rapid evaluation of the impact of aerodynamic selection on rocket dynamics and the overall scheme; in the sample stage, the wind blowing data has a huge amount of data, high-dimensional aerodynamics with hundreds of thousands or millions of rows, which requires efficient analysis and calculation functions) Since the aerodynamic parameters are flexible and diverse, returning to the essence, only aerodynamic forces and torques are required in dynamic calculations.

[0216] Therefore, the combination of equivalent three-channel and multiple sections is adopted to support the combination of local aerodynamics of different sections and multiple aerodynamic types under one section. The equivalent three-channel aerodynamic force and aerodynamic moment of the arrow system are finally synthesized and connected to the dynamics integral in the form of mounting. At the same time, it supports centralized aerodynamics and distributed aerodynamics, high-dimensional aerodynamic balancing, aerodynamic switching of different coordinates; and the composite control function of aerodynamic force and control force. This satisfies the aerodynamic calculation of various complex-shaped aircraft.

[0217] In order to process complex aerodynamic data, the aerodynamic modeling tool intends to use a tree data structure to encapsulate aerodynamic data. The key nodes included in the hierarchy include:

[0218] Segment refers to different aerodynamic models caused by the change of aircraft shape (or other reasons, such as different aerodynamic models at different Mach numbers) during the flight process. At any time, only one segment model is activated. The aerodynamic force and torque information finally output by the aerodynamic model are provided by the activated segment.

[0219] Channel, the aerodynamic model of a section is a linear combination of one or more channels; each channel is a complete set of default aerodynamic coefficient variables + aerodynamic variables that are not automatically identified.

[0220] Aerodynamic coefficient, an aerodynamic coefficient entered separately under the channel;

[0221] The aerodynamic coefficient table processes aerodynamic data encapsulated in a table format, which is a combination of aerodynamic coefficients of multiple identical independent variables.

[0222] Each channel has a set of default aerodynamic coefficient name identifiers and corresponding calculation rules. The aerodynamic data files defined by the corresponding identifiers can be directly read and generated to generate the required aerodynamic model.

[0223] There are many states of aerodynamic coefficients in the aerodynamic model, such as velocity system or projectile system, synthetic angle of attack or non-synthetic angle of attack, derivative type or non-derivative type, including pressure center or excluding pressure center. The aerodynamic module needs to be compatible with the above various forms of aerodynamic data. After sorting and summarizing, the calculation method of aerodynamic force and aerodynamic moment is as follows.

[0224] The calculation method of aerodynamic force under the missile system and velocity system is:

[0225]

[0226] Where, the superscript ' represents aerodynamic force or aerodynamic moment, F is force, M is moment. Q is dynamic pressure, S ref For reference areas, angles and aerodynamic coefficients, see the table below for definitions.

[0227] Table 1 Angle definition

[0228] Serial number Aerodynamic coefficient meaning illustrate 1. α Angle of attack The unit is degree 2. β Side slip angle The unit is degree 3. Alf Total angle of attack The unit is degree 4. γ Azimuth The unit is degree 5. <![CDATA[δ x ]]> Roll rudder angle The unit is degree 6. <![CDATA[δ y ]]> Rudder deflection angle The unit is degree 7. <![CDATA[δ z ]]> Pitch rudder angle The unit is degree

[0229] The aerodynamic coefficient variable names, meanings and unit descriptions are shown in the table below.

[0230] Table 2 Aerodynamic coefficients

[0231]

[0232]

[0233] The calculation of aerodynamic torque varies according to the given aerodynamic coefficient, and can be divided into two main cases:

[0234] (1) The aerodynamic torque can be calculated based on the aerodynamic coefficient, that is,

[0235]

[0236] Among them, L ref is the characteristic length.

[0237] If you choose to input the center of mass information, you can get the vector of the center of mass in the initial parameter coordinate system, set it as Rg, and the components are Rgx, Rgy and Rgz. Then the final aerodynamic torque on the center of mass is:

[0238]

[0239] If no centroid information is found, then

[0240] M b =M b '

[0241] (2) The pressure center information is given in the aerodynamic coefficient, which is set as Rp. Its components in the initial parameter coordinate system are Rpx, Rpy and Rpz, then Rpx = XcpLref, Rpy = YcpLref, and Rpz = ZcpLref. At this time, the aerodynamic moment of the center of mass is calculated as:

[0242]

[0243] If no centroid information is found, then

[0244]

[0245] Step 4: Abstract modeling of trajectory and flight sequence

[0246] Research on process abstract design technology based on core object modeling selection, including: flight variation parameter objects based on ballistic segments, and rocket full feature parameter objects based on structural levels. Through the mutual complementation of core objects, it can adapt to the modeling needs of different types of systems or the same type but with different functional requirements.

[0247] The simulation drive is highly refined and unified, and a universal and fast timing reconstruction module is designed to adapt to the trajectory design of launch vehicles, etc., which greatly improves the efficiency of trajectory design, simplifies the time of model and data organization, and ensures the design quality.

[0248] Flexible definition of flight sequence discrete events allows custom actions at the time of event occurrence; complex trajectories consisting of several flight segments can be calculated, and the trajectory can be flexibly driven by segment simulation by specifying time or conditions.

[0249] In order to meet the design requirements of different types of launch vehicles, the "ballistic shutdown timing drive technology" research was carried out, and the concept of flight segment modeling was proposed. This flight segment is different from the concept of physical flight stage (rocket stage one, rocket stage two, etc.), but is based on the characteristic values ​​of various flight parameters to divide the physical flight stage. For example, during the flight of the first stage of the rocket, the end of the flight segment ("shutdown") can be defined according to the characteristic values ​​of flight parameters such as time and speed. In this way, a physical flight stage can be divided into any number of logical flight segments according to needs, greatly improving the ability to design the trajectory of different flight profiles of launch vehicles.

[0250] Step 5: Trajectory Design Process and Program Corner Design Abstraction Method

[0251] In order to cover the different ballistic design requirements of various types of existing maneuverable vehicles in the atmosphere, we carried out targeted research on the abstraction of the program angle calculation model, and finally formed an abstract modeling of multi-stage complex program angle design. This modeling method includes 2 types of angle models, 3 channels, 6 angle variables, and 10 curve types, that is, it supports 2 angle models of small amount and full amount, supports 3 channels independent configuration and any combination of 6 angle variables of pitch angle / angle of attack, yaw angle / sideslip angle, roll angle / roll angle, and supports any multiple flight program curve splicing in a single flight segment, and each curve can support 10 different curve forms.

[0252] Table 3 Supported flight program curve forms

[0253]

[0254] This achieves a highly abstract process for complex trajectory design and a hierarchical decoupling technology for complex models.

[0255] The present invention takes into account various trajectory types such as one-time orbital insertion, orbital insertion with gliding segment, suborbital transition, pure thrust vector control, main and auxiliary aircraft coordination, air rudder review control, gliding trajectory within the atmosphere, warhead re-entry, return trajectory, etc., and highly abstracts the above trajectory design process to form a unified trajectory design model.

[0256] Considering the application environment that is close to the real world, the overall architecture model of the integrated trajectory design process is modeled, the trajectory design process is highly abstracted, and the decoupling and decomposition technology of the complex process is studied. The overall architecture model of the initiation, initialization, process monitoring, driving parametric design, calling simulation and saving of the entire design process is constructed. The test and analysis are carried out in the laboratory environment to meet the needs of trajectory design optimization and simulation with strong versatility and wide adaptability.

[0257] By establishing a core object architecture system that complements the missile / rocket trajectory segment and the structural level, the data source of the rocket flight change parameters and rocket characteristic parameters is unique, which meets the needs of system modeling of different types of systems or the same type with different functional requirements. At the same time, the ballistic segment is used as the main line to mount various ballistic configuration objects to further realize the flexible design of the ballistic segment. This technology has been verified in flight tests when the launch vehicle is performing a model flight mission.

[0258] In the real model library and domestic platform environment, a generalized and standardized method for constructing ballistic component models is constructed, which fully utilizes the componentization, modularization, and hierarchical characteristics of configuring components in the abstract ballistic design process, realizes interface standard specifications and rapid switching between multiple models, and unifies the model library and underlying model encapsulation interface call form of various aircraft ballistic design specifications. The compatibility test is completed with the basic component library for the final ballistic design of different platform models, and it has been formally applied in the development process of the model product, meeting the corresponding overall design requirements.

[0259] The function of complex flight profile trajectory design of maneuverable aircraft is taken into consideration. Through multi-segment program angle design, different combinations of the above configurations can be adopted in a certain section of the trajectory according to needs, which greatly broadens the support for ballistic flight types and can meet the current needs of complex ballistic flight profile design of various types of aircraft in the atmosphere in my country.

[0260] The universal modeling of complex timing of the present invention is particularly useful for large-scale bundled liquid launch vehicles such as CZ-5 and CZ-7, where several engines work simultaneously in the same physical flight stage, but the physical shutdown timing of each engine is different. The above-mentioned flight segment modeling method can be flexibly designed to fully meet the ballistic design requirements of such complex flight timing.

[0261] Although the present invention has been disclosed as above in the form of a preferred embodiment, it is not intended to limit the present invention. Any person skilled in the art may make possible changes and modifications to the technical solution of the present invention by using the methods and technical contents disclosed above without departing from the spirit and scope of the present invention. Therefore, any simple modifications, equivalent changes and modifications made to the above embodiments based on the technical essence of the present invention without departing from the content of the technical solution of the present invention shall fall within the protection scope of the technical solution of the present invention.

Claims

1. A method for highly abstracting complex processes in trajectory design and decoupling complex models in layers, characterized by: include: Set the minimum integral unit and select the core modeling object of the launch vehicle according to the minimum integral unit; Establish the thrust equation for launch vehicle engines; Calculate the small deviation thrust in the engine thrust equation; establish the propellant flow deviation equation based on the small deviation thrust; Calculate the engine control force F pk and control torque M ck ; Equivalent three-channel aerodynamic force and aerodynamic moment of synthetic arrow system; determine aerodynamic modeling tools; generate required aerodynamic models; Abstract modeling of trajectory segments and flight timing; Perform complex procedural corner abstract modeling of multi-segment trajectories.

2. According to claim 1, a method for highly abstracting complex processes in trajectory design and decoupling complex models by layers, characterized in that: The minimum integration unit contains basic position, speed and attitude information; Any other transformation of the design model and addition of the design process are connected to the minimum integration unit; flexible interpreters and scripts are also connected to the minimum integration unit.

3. A method for highly abstracting complex processes and hierarchical decoupling of complex models in trajectory design according to claim 2, characterized in that: The minimum integration unit satisfies the following equation: Where x is the state vector; t is the integration time; The superscript · indicates the derivative of a variable with respect to time t; is the derivative of position r with respect to time; v is the velocity vector, a is the acceleration vector; m is the mass, dt is the derivative with respect to time; dm is the derivative of mass with respect to time; represents the derivative of a variable with respect to time t; x, y, and z are three-dimensional positions respectively; V is the speed; W is the apparent velocity; g is the gravitational acceleration; The subscripts x, y, and z are the three directions respectively; The subscript a indicates component form.

4. A method for highly abstracting complex processes and hierarchical decoupling of complex models in trajectory design according to claim 3, characterized in that: The method for selecting the core modeling objects of the launch vehicle is: The method of abstract ballistic segment mounting structure level is adopted; flight timing, mass, aerodynamics, and structure are all abstracted as configuration objects; the mission design takes the ballistic segment as the main line, and various ballistic configuration objects are mounted.

5. The method for highly abstracting complex processes and hierarchical decoupling complex models of trajectory design according to claim 3, characterized in that: The thrust equation of the engine in rated working section is: Where, P0 is the rated ground thrust of a single engine; S a is the engine nozzle area; ΔP i is the small deviation thrust, which is calculated by the engine small deviation equation; p0 is the standard ground atmospheric pressure, p is the local atmospheric pressure; In the startup and shutdown working stages, the source variable uses the measured value provided by the engine unit, so the startup and shutdown stages ΔP i is zero, the following thrust equation is used to calculate the thrust 6. A method for highly abstracting complex processes and hierarchical decoupling of complex models in trajectory design according to claim 5, characterized in that: The small deviation thrust equation is: ΔP i =(c 11 ΔT Oz +c 12 ΔT fz +c 13 ΔP Oz +c 14 ΔP fz +c 15 ×(n x1 -1)) The propellant flow deviation equation is: Where Δ is the increment, which is the deviation from its standard value during flight; ΔP i It is a small deviation thrust; is the propellant flow rate per second; T is the propellant temperature; P is the engine pump inlet pressure; n x1 is the overload coefficient along the longitudinal axis of the rocket body; Divide by n x1 In addition, the subscripts of each symbol mean: The subscripts Z, 1, and 2 denote booster, core first stage, and second stage, respectively; The subscripts O and f represent oxidant and combustion agent, respectively; c11 to c35 are the coefficients of the small deviation equation and are constants.

7. A method for highly abstracting complex processes and decoupling complex models of trajectory design according to claim 6, characterized in that: The thrust of the engine along the thrust line is P; following the design convention, it is defined in the swing coordinate system y s The positive swing angle is δ1, z s The positive swing angle is δ2; the subscript s represents the swing coordinate system component; δ1 is the swing angle around -z s Axis swing; δ2 is around y s Axis swing; projecting the engine thrust onto the engine swing coordinate system requires rotating around z s Axis rotation δ1, around y s The axis rotates -δ2, and the force generated is: From the engine swing coordinate system to the arrow system, there are three rotations, namely, from the engine swing coordinate system to the engine installation coordinate system; then the installation angle and quadrant angle are rotated to the arrow body coordinate system in turn. The conversion matrix is: In the formula, δ f Swing the azimuth angle for the engine; δ A Mounting angles for the engine; δ q Mounting angles for the engine; The distance from the control force point of the kth engine to the rocket axis is R k ; The distance from the center of mass to the theoretical tip is X z , the distance from the engine to the theoretical tip of the rocket is Then the position of the engine control force action point in the rocket body coordinate system is for: The projection of the kth engine force on the arrow system is: In the formula, is the transformation matrix from the swing coordinate system to the rocket body coordinate system; F is the thrust vector; The kth engine force acts on the arrow system with a torque M ck for: In the formula, is the force arm vector; but:

8. A method for highly abstracting complex processes and hierarchical decoupling of complex models in trajectory design according to claim 7, characterized in that: The equivalent three-channel aerodynamic force and aerodynamic moment of the arrow system are synthesized by combining the equivalent three-channel and multiple sections. The aerodynamic data is encapsulated in a tree data structure as an aerodynamic modeling tool. Each channel corresponds to a set of default aerodynamic coefficient name identifiers and corresponding operation rules. The aerodynamic data file defined by the corresponding identifier is directly read and the required aerodynamic model is generated.

9. The method for highly abstracting complex processes and decoupling complex models of trajectory design according to claim 7, characterized in that: The method for abstract modeling of trajectory segments and flight timing is: The simulation-driven timing reconstruction module is adapted to the launch vehicle trajectory design; Define flight sequence discrete events, allowing custom actions when events occur; Calculate complex trajectories consisting of several flight segments, and flexibly drive the trajectory in segments by specifying time or conditions; The physical flight level is divided based on the characteristic values ​​of various flight parameters; the end of the flight segment is defined according to the characteristic values ​​of time and speed flight parameters.

10. A method for highly abstracting complex processes and hierarchical decoupling of complex models in trajectory design according to claim 9, characterized in that: The complex program angle abstraction model of multi-segment trajectory realizes small-angle model and full-angle model; realizes independent configuration of 3 channels; realizes arbitrary combination of 6 angle variables of pitch angle / attack angle, yaw angle / sideslip angle, roll angle / roll angle; supports any splicing of multiple flight program curves in a single flight segment; each curve supports 10 different curve forms.

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